M11 Aerodynamcis,Structures and Instruments 2 Of2

CONTENTS

Page

Types of Generators 1 Permanent Magnet Machines 1 Rotating Armature Type 1 Rotating Field Type 4 Connection of Phases 7 Power in 3 Phase Systems 9

Frequency Wild Systems 10 Brushless Generators 13 Volt-a-geegulation . 16 .-.

'.\ i

16 -'\

\ r i- 7 @or Sensing ~ e g u l k t ~ \, >, , -, , '.> .,

1 '/r'ransistorised Voltage B&iylatoS ' 17 m i I , I i ., , \ , i I , ! I

19 ', '

Cons tv t Frequency Systemq ; ; : : i

I j

20 i j CSDU in a on-~arallegeh Systek I i i

I Indications ! : 1' ,/ 2.3 I , / L-

./ 8 . / /'

on-~balleled Systems i I., i -1' 26

, , ---1 \

! --// /:

1 I Fault Protection \,,, , ',.~,, '. 29 i 33

I : i M ~ ~ U ~ ; F C R Tripping ! ! ! i ! . I , ElectriQ Load Control: Unit , ,

, , 36 j 1 r :

i / 36 i ' ! ~ o d i ~ h e d d i n ~ : 1 , , , I ,

f' g-j~' , -,' .------I I ; 1~

/' L 1 !- _ ,-' 39 7-' L -

-

Flight Deck Indication 41 Manual Paralleling 43 Automatic Paralleling 44

Paralleled Systems 45 Load Sharing 49

Power in ac Circuits 49 Power Factor 51 Load Sensing 53 Reactive Load Sharing 54 Real Load Sharing 55 Fault Protection 56

Variable Speed Constant Frequency Generator 60 Operation 62 BIT 63

AC GENERATION

In ac generators, the rotating part of the generator is called the ROTOR and the stationary part is called the STATOR. There are three basic types:

* Permanent Magnet Generator * Rotating Armature Generator * Rotating Field Generator

Permanent Magnet Type

The rotor of the machine is a permanent magnet and as the magnet is rotated its magnetic field cuts the stationary output windings producing an alternating voltage output. This type of generator (or variations of it) is used as part of the most brushless ac generators (see later text in this book).

\

I i I i I

rotates acoutput -- -' from stationary

I windings ,

I i

Fig. 1 SIMPLE PERMANENT MAGNET ac GENERATOR

Rotating Armature Type

This type of ac machine is similar in construction to a dc generator in that the rotor rotates in a fixed field with the emf picked off via slip-rings. The rotor windings are laid in slots along the rotor periphery, the armature being laminated to reduce eddy current losses. The stator carries the dc excitation windings wound on the pole pieces to create alternate north and south poles around the stator. Figure 2 shows a single phase 2 pole machine with the output as shown in Figure 3 .

FED FROM dc SUPPLY

SINGLE PHASE ARMATURE WINDING

BRUSHES

-

EMF laoO 2100 240' 270' 3o0° 330P 360' EMF

0' 30' 60' 90' 120' !SOo 180

Fig. 3 GRAPH OF INDUCED EMF - SINGLE PHASE ac (The double arrows indicate the single-phase windings position)

One cycle of voltage is induced when the conductor moves through 360" past one pair of poles. If there are two pairs of poles then two cycles of ac will be produced.

The number of cycles of induced voltage of an actual generator will correspond to the number of pairs of poles in the generator and is called the frequency (0.

- 2 - rnoodull l A-740

f = Np 60 Hertz

where N = speed in rpm of the generator at which the generator must be driven in order to generate the required frequency

p = number of pairs of poles.

QUESTION: To provide an output of 400Hz what speed must a two pole machine be driven at?

ANSWER: Transposing the formula

I I ( 1 I I , ---I

' / /

An ac Cenerator, in which the I who_lc - ofthe \ output consisds o f a-singie windidg with the outer ends aonnected\to,a pair of slip-rikgs,- is-tegmed a

I I 'single phase geperator'. If thgre were tdo yindings at differknt angles congebted to $Gp rings then thik would bye two outputs &d would be kno- as a Ywb phase generator'. 1 I

1 I

- - I /

i J L-1 : i -.

FIELD COILS

CWINDMG AT 0°1

SLIP RINGS

PHASE C (WINLIING AT 240"

Fig. 4 THREE PHASE TWO POLE ac GENERATOR

- 3 -

rnoodull l A-741

Figure 4 shows a three phase system, in which the coils are at 120" to each other and a 3 phase output is generated. In other words it is really 3 generators in one with 3 separate outputs each one 120" out-of-phase with the next.

,' Fig. 5 CURVE YF-INDUCED EMF - 3 PHASE - 1 1

1 I I ' 1 I I I I

Three1 phase supplies are used extensively on aircraft - ak it is on most national grid systems. This type of generator, however, is not used,as a main generating source on its own as it\has the following disadvantages.

1 r --' ' 1 I '

I (a) As all the power is taken from the rotor, the efgective insulation and v'eritilation causes problems.

- - -

(b) All the (heavy) output is taken via slip-rings - and brushes.

(c) Centrifugal forces are considerable on the rotor windings.

Rotating Field Type

In this type of generator the dc field rotates and its field cuts the stationery (output) windings on the stator. The output windings consist of a number of coils connected in series and inserted in slots in the laminated stator to give a single phase output. The field windings are supplied with dc via two slip rings and brushes.

The principle of a two-pole single phase ac generator is shown in figure 6.

SINGLE PHASE

Fig. 6 WIRING SCHEMATIC - TWO POLE SINGLE-PHASE ac GENERATOR

STATOR WINDINGS

I

ac OUTPUT

DRIVE SHAFT

ROLLERM RACE

dc INPUT .-

I i', , FIELD WINDINGS ROTOR I

1 , I I 1 , I

/ Fig. 7 SEC'~'I,ON O$;R~TATING FIELD( I SINGLE~P~ASE ~ C ~ E N E R A T O R ;' -;

_I - I

The general arrangement of a single phase rotating field ac generator is shown in figure 7. Note the drive rotating the field windings, the power of which controlled by the dc input via the brushes. The ac output taken d&ctly from the stator windings. /\J-'

Fig. 8 ROTATING FIELD TWO PHASE ac GENERATOR

'. '. '. '. '.

time

Fig. 9 GRAPH OF TWO PHASE OUTPUT

If another set of single phase windings at 90" to the first set is added, then a two phase output is produced being 90" out-of-phase with the first (figures 8 and 9).

If afurther set of two coils is added and each coil in the system is spaced a t 60" to each other then we have a three phase system. Each pair of coils is-spaced at 120" to one angther so there are 3 phases where the 3 outputs are 120" out of phase YfigUre\lO). I

I I

Fig. 10 TWO-POLE THREE PHASE ac GENERATOR

The advantages of the rotating field generator over the rotating armature type are:

(a) Only two slip-rings and brushes are used taking less current, ie field winding current only.

(b) Less problems with centrifugal effects on rotor windings. (c) The output is taken from the stator, where ventilation and

insulation of windings is less of a problem.

Connection of Phases

Each phase of a three phase generator can be brought out to separate terminals and used to supply separate loads independently, which would require a total of six leads. However, a considerable saving in cable (and weight) and other advantages can be obtained by connecting together a lead from one end of each of the three phase-windings as shown in figure 11.

This shows that the three windings are connected to one point and a lead is taken from that point. This configuration is called STAR CONNECTION and the point where they meet is called the star point and the lead taken from the star point is called the neutral lead.

l line

I / ; , I

! 1 ! I ~: ! I ' i i .' j 1 i _ - - l " ' , Fig. 11 STAR CONNI$CTION GENERAT&

/ L - ~ - _ - ' ,, , -1 L l.?-l-: -- // i... .--I

Figure 11 shows that the line current = the phase current.

The phase voltage (Vphase) on an aircraft generator would be 1 15V and the line voltage (Vs,,) which is the sum of the two phase voltages across that line, ie two 115V phases at 120' phase angle, is 200V and mathematically is the same as multiplying the phase voltage by 43.

The big advantage of the star connection is that with the neutral line two voltages are available - 200 and 115. Aircraft ac generators are generally connected in star.

Another connection of the three coils would be to connect them as shown in figure 12, known as DELTA CONNECTION. In this case the three windings are connected in series to form a closed mesh, with the three output lines at the junctions.

l line L

l phase

l phase Vphase V line

l phase l llne -

B

l line L - t 1

Fig. 12 DELTA CONNECTION LAYOUT

As can be seen from figure 12

- L>

! 't f

In this connection the line carrent is composed of two components and mathematically it can be shown that:

I

IL = 4 3 1 ~ - / 1 1 1

, /

i 1 . , The delta connection does n f t have a neutral and cannot provide two outputs and must be connectdd to a balanced load, but does give a higher current outgut than a star connected system. ' I

I /

, I

NOTE: The reason why these interconnections can be used is that in a three phase system the instantaneous sum of theemfs or currents in a balanced three phase system is zero. Look at figure 13 and if, on the line indicated, you add the voltages together the sum is zero. The same applies to the current waveforms.

sum of instantaneous voltaees VI + VI + V3 = 0 sum of instantaneous currents 11 + I + 11 = 0

Fig. 13 THREE-PHASE WAVEFORMS

Power in Three Phase

The power in a single phase system is:

True Power = V x I x cos 0 Watts

True power in a balanced three-phase star or delta system must be three times that in a single phase system, so:

True Power = 3 Vph Iph cos 0 Watts

In a star connected system IL = Iph

So the formula can be written:

True Power = 3VphI~cos0Watts

and as VL = d3vph - --

Vph = & 63 - .

i - - \\

1- - - \

\ \ 1 I \

Then True Power = 3 x ~ I I C O S 0 watts I 1 63 I I I

I I

I ' I ' i / I

1 1 I

1 = d 3 ~ ~ 1 ~ c o s 0 w a t t s I

I

I - - , I

I 1 -7 I \, '\ I -- --/

/

1

1 \ i ' 1 For a kelta cqnnected sys te4 VL = Vph I

I ' I '

/ 1 'Then dhe fornula can be wri t tehL- :, - -

1 --- - - - - - --- / ' --2

True Power = ~ V L I ~ ~ C O S 0 watts

and as IL = d 3 1 ~

True Power = ~ V L I~cos0watts

So for star or delta connected systems there are two formulas for power.

1. True Power = 3Vph Iph cos 0 watts

2. True Power = 43 VL IL cos 0 watts

FREQUENCY WILD SYSTEMS

The first ac generators used on aircraft were rotating field genera-tors in what was called a 'frequency wild' system (the generators giving out a frequency depending on their rpm, which depended on engine rpm). The ac generator gave an output of 208V and was connected to a main control unit, which converted this ac into dc, so the aircraft busbar was 28V dc.

These generators were preferred to dc systems because they had a better power to weight ratio and were much less affected by poor brush performance a t high altitudes. The output voltage of the generator was controlled by controlling the field strength of the rotating field ac generator by a signal from the voltage regulator in the control unit which kept the voltage constant irrespective of load or speed variations.

At this time electrical heater mats for de-icing had created the need for large quantities of power. A s these elements are resistive, the variation of frequency had no effect, so the frequency wild constant voltage supply was fed to resistive circuits such as windshield heating, heater mats on the

- airframe, engine de-icing heaters etc. ' 1 I

I I

Figure 14 shows the basic layout of a iectified ac (frequeAcy wild) system. Note the symbols used to dehote 3 phase (three short parallel lines).

1

TO RESISTIVE CIRCUITS eg ELECTRICAL HEATERS

AC GENERATOR (CONSTANT VOLTAGE FREQUENCY WILD) DIRECTLY CONNECTED TO f DC BUS

Fig. 14 SIMPLIFIED FREQUENCY WILD SYSTEM

ENGINE GEAR BOX

- 10 -

rnoodull I A-748

TRU TRU CONTACTOR

- 4 6 . 2

111 = * 111 "r"

OUTPUT FREQUENCY PROPORTIONAL TO GENERATOR rpm VOLTAGE

REGULATOR

TO DC SERVICES

d - DC TO ROTATING FIELD

' I \ - The TRU transforms the ac voltage down to 28V and then rectifies it to dc

104V OUTPUT

EXCITATION WINDING

Fig. 15 GENERATOR WIRING DIAGRAM

This generator is of the rotating field type. A s can be seen from the intEnd-circuit diagram it hzsmO stator windings, one1prodming 208V and'thg other 104V. The speed-r%Z~2 ,i$3,300rprn to 1 O:lOEpin giying a

range of 165Hz to s05Hz. ~he'\rotor winding id supplied with via two slip-qngs. 1 I I I

I i 1 ' ,I / I ! I m- ' ,i , ,__ --/

{REQUENCY WILD \ / TO 2 0 0 ~ AC SERVICES \

112V DC

TO MIC (28V)

BUS +VE

Fig. 16 PRINCIPAL COMPONENTS OF A ONE CHANNEL RECTIFIED ac SYSTEM

Figure 16 shows the principal components of one channel of the system. The reason for the two outputs from the generator (2081 104V) is that the main busbars on older aircraft were 1 12V and 28V dc.

The supply from the 208V winding is fed via the compounding TKU to the frequency-wild services. From terminal A1 of the Main Control Unit a supply is fed back to the rotor winding to maintain the voltage constant.

In this system stable ac is provided by invertors.

A number of aircraft have a small ac generator which only supplies the anti-icing and de-icing systems. One such generator is shown in figure 17.

OUTPUT PROTECTION TERldINAL BOX I7UNSPORHgR

ROTOR \ /

STATOR I BALL RACE

COOLING A I R Durn

Fig. 17 FREQUENCY WILD THREE-PHASE ROTATING FIELD GENERATOR

The generator has a n output of 15kVA* a t 208V; its speed range is 6,700 to 10,700 rpm giving an output frequency of 335 to 535Hz. It is a rotating field generator, with six poles, which you can see in the diagram, they are often called 'salient' poles. The dc is fed to the poles via slip-rings and brushes.

*kVA is a term used to denote the power output from a generator and is kilo Volts Amps. (From module 3 volts x amps = watts. Not strictly true for generator outputs but the two can be equated as a very simple rule).

BRUSHLESS GENERATORS

There are obvious problems associated with brushes rubbing on commutators, so a generator designed without these has considerable advantages in terms of wear, sparking etc.

There are many types of brushless ac generators in use, we shall deal with two variations on the theme.

- -

GENERATOR " " .... -

4- ROTOR

WN ROTATING

ROTATING FIELDr ...

TRANSFPRMERL

Fig. 18 WIRING OF BRUSHLESS GENERATOR WITH VOLTAGE REGULATOR

With reference to figure 18. There are two main parts in its construction:

(a) The Exciter, which is a Rotating Armature, Star-wound a c generator.

(b) The Main ac Generator, which is of Rotating Field, Star-wound construction.

Operation

(a) Permanent magnets are interspersed between the main poles of the exciter field on the stator. A s the exciter output windings are turned, the weak magnet field of the stationary magnets is cut by the exciter windings inducing an emf which is fed to the rectifiers.


(b) Silicon diodes located on the rotor form a 3 phase, full-wave rectifier bridge, converting the output into dc, which is smoothed by a capacitor and fed along the rotor to the rotating field poles of the main generator.

(c) The rotating field induces emf's into the main 3 phase stationary output windings which are star-connected (externally in this case) to give phase voltages of 1 15V rms and line-to-line voltages a t 200V rms .

(d) Apart from feeding out to the loads, the generator output is fed back, via a voltage regulator, a s dc onto its own exciter field windings. The exciter field windings take over from the permanent magnets in controlling the output voltage of the generator. (In this machine, the permanent magnets play no further part in the operation as their flux densities are low compared to that of the exciter field coils).

(e) Two field windings are provided for the exciter. One (Fl) is I- permanently in circuit. The-second (F2) is only brought-into operation in the event of a temperature rise causing the resistance of F1 to increase to the point where the voltage regulator is unable to sustain sufficient current through it to maintain the voltage. (This is known as ~ i e l d Critical Resistance). As thk temperature rises the thermistor rebistance decreases and allows current to flow

-

' through F2, so assisting F 1. I

1 1

(f) 'Also wound on the of the exciter is a third wihding. This is a stability feed-back winding. Under steady operating conditions, no emf is induced into this winding, but in the event of rapid changes in field current, emf's are induced and currents a re caused to flow in the stability windings in the voltage regulator in such a direction as to prevent over-reaction. Thus the system is stabilised.

(g) The current transformer boost circuit assists the voltage regulator during load changes and during periods of large unbalanced loads on the generator.

Also incorporated in each generator is a thermostatic switch which switches on an overheat warning light if the temperature limit is being exceeded.

The second type of brushless ac generator is shown in the figure 19.

There are three main parts in its construction:

(a) The Rotating Permanent Magnet Generator (PMG) - sometimes known as the Pilot Exciter.

- 14 - moodull l A-752

(b) The Main Exciter, which is a Rotating Armature, Star-wound ac generator.

(c) The Main AC Generator, which is of Rotating Field, Star- wound construction.

GENERATOR CONTROL UNIT

3 PHASE 1151MOV 4(#3 Hz ac OUTPUT TO 3 SEPARATE

ROTATING FIELD

EXCITER

Fig. 20 PART DISMANTLED BRUSHLESS GENER

When the generator drive shaft is rotated the permanent magnet rotates and its field cuts the three stationary star connected coils and induces an ac current and voltage into them. This is fed externally to the Voltage Regulator section of the Generator Control Unit (GCU - the %rains7 behind the control of the generator).

The regulator section controls and rectifies this output and feeds a dc current to the main stationary exciter winding. This is controlled to keep the generator voltage output constant irrespective of load. This dc field is cut by the rotating star connected exciter armature winding which induces an ac into it. This ac is fed via the six diodes on the rotating armature to give 3 phase full wave rectification and therefore dc to the main field coil (also being rotated by the engine). This rotating field cuts the star connected windings of the main generator to give 1 15/200V 3 phase 400Hz output (400Hz provided the generator is rotating at the required rpm) .

Part of the output is fed back to the voltage regulator in the GCU which controls the dc to the exciter field-and hence the generator output.

1 VOLTAGE REGULATION ,

Voltage regulation is accomplished by varying the field strength of the ac generator's exciter field in order to keep\the output voltage konsttan't under varying speed and load conditions. There are two main methods of carrying this out: 1 ~

1. The exciter field is fed with dc, which is varied in strength using an ~ r r o r Sensing Bridge:

- -- - -

2. The exciter field is fed with a stream of pulses, the amplitude of which remains constant whilst the width of the pulses is increased to increase the overall field current and decreased to decrease field current.

Error Sensing Bridge Method

Operation (single generator running). The output of the ac generator is fed via a full-wave rectifier bridge, complete with choke/ capacitor smoothing, to an Error Sensing Bridge. A trimmer is used to set the circuit so that, when the ac output is correct (200V)' 170Vdc is applied to the bridge. The two constant voltage tubes in the bridge each maintain a voltage of 85V across itself throughout it's working range. Under correct output conditions the points 'A' and 'B' are a t the same potential and no current flows in the control winding of the magnetic amplifier.

Fig.

' ', ,>

/ I Incrcaked line~uoltage (cause increased speed or reduction in load)

cau$es an increased dc voltage to be to the error sensing bridge. DUX' td-the-tion of the conskant_voltag~ lubes, point A now becomes

i- pdsitive-with-iespect to point-B and-euirent flows through-th2 control winding from A to B. This has the effect of reducing the current in the exciter field and so reducing the generator output voltage back to normal.

Decreased Line Voltage (caused by decreased speed or increase in load) causes a n action that is opposite to that of the previous paragraph, with point B becoming positive with respect to point A. This causes current to flow in the opposite direction through the control winding, resulting in a n increase in field current and an increase in generator output.

Transistorised Voltage Regulation

The output from the PMG (figure 22) is fed to the star connected primary of a transformer in the GCU. The star connected secondary of the transformer feeds a combined voltage regulator and Transformer Rectifier Unit (TRU) to ensure the voltage to the field circuit is a constant dc voltage.

G E N E R A T O R OUlPUT -

i I

Fig. 22 TRANSISTORISED VOLTAGE REGULATION I ' I

The circuit continues through a contact of the Generator, Control Relay (GCR), through to the main exciter field windings, back into the GCU to a transistor in the output stage of the voltage regulator to e a r t h . This is the field circuit - note the GCR cdntact, this, i s very important because under fault conditions the GCR is tfipped and the field circuit will be broken.

The generator output is fed via therectifiers to a sensor in - the voltage regulator. It is compared t o a reference value and the difference signal will signal the amplifier to switch the transistor ON and OFF, this ON/OFF pulse is varied according to whether voltage is required to be increased or decreased, ie effective current to the field is increased or decreased.

EFFECTIVE FIELD CURRENT

NORMAL GEN VOLTAGE LOW GEN VOLTAGE HIGH GEN VOLTAGE TIME (NORMAL LOADING) (HIGH LOAD) (LOW LOAD)

Fig. 23 PULSE WIDTH MODULATION

- 18-

rnoodull l A-756

If the generator loading was to increase, the terminal voltage of the generator would decrease (due to internal voltage drop), this would be sensed by the regulator which would signal the switching transistor to increase the width of the pulse, ie stay on for a longer period of time. The pulsing is fast so the field sees a n effective current, which will increase in this case to increase the output of the generator.

If a very heavy load was taken off the generator, the terminal voltage of the generator would tend to rise. The regulator this time would signal the switching transistor to switch on for a smaller period of time, lowering the effective current, which will decrease the current to the field, lowering the output of the generator.

The voltage amplitude of the pulses remains the same, it is just the width of the pulses that is varied, and hence the name given to this type of regulation system is PULSE WIDTH MODULATION.

r -~ CONST@%L!-FR-EQIJENCY SYSTEMS; '~ ', ,, ',

i 1 I--- - - \, 1- ----\ '. For &craft where the load i s only resiitiye (heaters - anti-@ing, :I >,,

windstreen heaters etc) then ifrLquencii wild systems can bd used. 1f \ circhids that are inductive or r4active +ejused then the f~!e uencypf 1

I I I supply must be constant becidbse-the/im'pedance (resistance) varies Fi th the fr&uency (module 3 ~lecbqea l -~u*~en ta l s ) . 1 - - - / j

For w s t aircrqt the generator is conne1ct$d to a unit callled a Constant spegd l~r ive W i t (CSDU), thik in turn i i cpnnected to theiemgine. The

/ CSDUis j*d f o ensure the dederator runs at a constant speed and hence ' / conktanT freqyency, irrespective<fthe,speed L__ - - -._- of the engine. The speed of

the generator is 12,000 rpm, but many aircraft have running at 6,000 rpm or 8,000 rpm.

(On some aircraft a generator is used called a Variable Speed Constant Frequency (VSCF) generator which runs at a speed related to engine rprn and a constant frequency output is obtained electronically - more of this later).

The CSDU and the generator can be separate units but in later aircraft they are one unit which is called an Integrated Drive Generator (IDG).

In general there are two types of ac power distribution systems:

a) Non paralleled b) Paralleled

In a non paralleled system each generator supplies its own bus and all the services attached thereto. So, in general, if there are four generators then there are four buses each supplying their own services. The buses are interconnected by relays so if one generator output drops off then another generator can be switched in to supply some power to that bus. Sometimes called a Non Load Sharing system.

With a paralleled or Load Sharing system all generators share the load to the busses. This means that each generator is taking exactly the same load as each of the others.

CSDU in a Non Paralleled System

In this CSDU we are going to look a t an aircraft fitted with a JT8D engine, which is a twin-spool axial flow turbo fan engine. At the front of the N2 compressor is a vertical shaft called the Tower Shaft, which is driven by the N 2 shaft. The tower shaft drives the accessory gearbox and all the mechanical accessories such as the oil pimp, fuel pump hydraulic pump and theJ CSDU.

\ I I

The CSDU is capable of adding or subtracting from the speed received from the engine gearbox (4,300 to 8,600 rpm) to maintak khe generator speed' a t 6,000 rprn and the frequency at 400 Hz with small allowable tolerances. /

1 -. \

1

' I -

Fig. 24 CSDU/GENERATOR - ENGINE LOCATION

.- I N E G A T I V E I

P O S I T I V E OIFFEUENTIAL SPEED 1 OIFFLRENTIAC SPEED I - - I

OUTPUT SPEED

DIFFEREHTIAL SPEED

IN PUT SPEED

Fig. 26 PRINCIPLE OF CSDU DIFFERENTIAL DRIVE

This will make the pump act like a conventional axial piston pump and suppl$ fluid to the motor, causing it to rotate - in the s-&e direction as the whole unit is rotating. I

This means the rprn of the dotor is now added to the rprp of the whole unit to bring its output speed to that of the constant rpm required.,This is

--

called Overdrive. --

,

I

When the engine is running faster than the Straight ~ r i & speed the control pisfon moves the wobbler plate in the other direction causing the output from the pump to be the reverse from the Overdrive condition.

This will cause the motor to rotate in the opposite direction from the hydraulic unit, deducting its rpm from the hydraulic unit's rprn and causing the output rpm to remain constant at the required speed. This condition is called Under-drive.

The drive input and the output to/from the hydraulic unit goes through an axial differential gear box which houses a set of cyclic summing gears that sums the output from the hydraulic unit to the output shaft to the generator.

I Fig. 27 CSDU

Fig. 28 CSDU & STANDBY POWER PANEL

Indications

Figure 28 shows the CSDU and standby power control panel.

At a pressure of 120 to 160psi, the electromagnetic pressure sensor in the CSDU will cause an amplifier to ground an AMBER low pressure warning light.

At a temperature of 157°C in the oil reservoir, a bi-metal switch will ground the AMBER high oil temperature light. Both amber lights would be accompanied by MASTER CAUTION and ELECT annunciator lights.

Two temperature bulbs measure the oil temperature change either side of the oil cooler. One bulb measures the input oil temperature to the CSDU and is read on the meter on the power panel. A switch on the panel alters the circuit to include the oil out temperature bulb so that the meter now reads the rise in oil temperature through the CSDU. 5 to 10°C rise is normal.

Figure 29 shows the circuitry involved with the CSDU indications.

Fig. 29 CSDU INDICATOR CIRCUIT

WORM ENGAGEMENT- THREAD

AND GOVERNOR TO CSDU DISCONNECT .h SWlTCH

RESET HANDLE

Fig. 30 CSDU DRlVE 8~ DISCONNECT

In case of a mechanical fault on the generator there is a disconnect on the drive between the engine and the generator and, in case of serious jamming, there is a wasted drive shaft that will shear.

The disconnect solenoid is a guarded and lock switch on the power panel. The normally closed contacts of the switch place a ground on the line to prevent the possibility of a voltage pick-up from inadvertently tripping the CSDU.

When the switch is activated one pole of the switch sends a signal to trip the GCR which will trip the generator off-line. The other pole energises the disconnect solenoid. This allows a spring-loaded pawl to move into contact with threads on the worm gear. The input shaft acts as a screw in a threaded hole and input rotation causes the input shaft to move away from the input splined shaft, separating the driving dogs on the two shafts. When the driving dogs have been separated, the input splined shaft, which is still being driven by the engine, spins freely in the transmission without causing any transmission rotation (figure 30).

~ & e t may only be following anl engine I shutdown - by the solenoid nos'e pin

snaps 'into position. I I

I ! 8 8

I 1

Fig. 3 1 IDG SCHEMATIC

- 25 - rnoodull lA-763

Figure 3 1 shows an IDG schematic of a more modern aircraft (Boeing).

The IDG supplies 115/200V ac 3 phase 400Hz and consists of a CSD and a brush-less generator in a common housing. The gearbox input speed is 5,800 to 9,975rpm and the output speed is 12,000rpm. So in this system the CSD only adds rpm to the generator.

The governor adjustment allows adjustment of the IDG output frequency, if the frequency is just outside the 400Hz i- 5Hz, a governor adjustment may be performed. One turn changes the frequency 3 to 3.5Hz, counter- clockwise to increase and clockwise to decrease.

Note the input speed sensor and the disconnect mechanism. IDG temperature sensor resistance's sense the IDG 'oil in' and 'oil out' and sends signals via the Generator Control Unit (GCU) to the EICAS display.

Figure 32 shows the unit complete.

Fig. 32 INTEGRATED DRIVE GENERATOR

NON PARALLELED SYSTEMS

Figure 33 shows a typical system. It consists of 2 main generators (IDG L and IDG R) with an APU driven auxiliary generator provides a back up generator in flight and a self sufficient power source for ground operation. An external power source can be connected to the ac tie bus through the external power receptacle and the external power contactor (EPC).

( D I G I T A L DATA ----- CONTROL SIGNALS - POUER FLOU

ELEC SYS PANEL

I I I I

BTB BTB I

I I I I I FWD GALLEY I -- 1 7 - - . L------L-----A------------------------a7---------------- - i -

-- -- \ -, r- Y I '\ '1,

I ' I

' I Fig. 33 N&q P A R ~ L ~ L E D SYSTEM 1 , I

I I

I I 1 / I I

1 ' /

If the Arcraft is on the g o u n b y i t h - t h $ ~ ~ ~ running then tke-~u.&ar~ power 5reaker (APB) will be closed with,^?^ open, the ~dne-tor Circuit B r e e e r s (GCB's) will be open dnd both Bus Tie B r e a k e r s 1 ( ~ ~ ~ ' s ) contacts will be c l o s e q ~ o the APU geverator cad fied both main ljusbars and othek rielevant - busbars. It is jmPortant/to1note __- 1 ' there is ad interlock

prevents any _ - two sources'of power beingrparalleled to one another. Usually the power coming onto the system h a ~ ~ o r i t y .

If the engines are started and both generator outputs are okay then before the GCB's are energised the APB must trip and both BTB's must trip, leaving each generator feeding its own busbar and relevant busbar, ie non-paralleled. If one generator should fail in flight the APU may be started and its generator output can be fed to the relevant busbar.

For example, assume IDG L fails, then its GCB will trip, its BTB will close and the APB will close and feed that busbar. If no APU generator was available then both BTB's would close after IDG L GCB opened so the left busbar could be fed from the right generator (the generator could not give out more than its rated maximum so demand would have to be reduced and some services curtailed).

- 27 -

moodull lA-765

The generator control units (GCU's) provide automatic control and protection function for each channel by monitoring the IDG output. The unit contains the voltage regulator and all the circuits for fault protection. The power required to operate internal GCU circuits and external GCB and BTB contactors is derived from the IDG PMG source with backup from the aircraft 28V dc.

: I : : . : ; ' '"I""" SUPPLYlfRV R E W T O R CURREUT

I l l I

C W L l U G A I R VENT HOLES

\CIRCUIT BREAKERS

PHG I , GEMERATOR

POR VOLTAGE EXTERHAL C T ' S b GEMERATOR CT b 6CU W S I T I O U C BUS T I E SYITCH h 1 GEM COHT SU DRIVE DISC SY F IRE SYITCU F I E L D SUITCH IOG O I L T E W I b G IPU CHARGE PRESS SU A W SHUTDOVW AIR I6UD 6CB AUX CONTACT BTB AUX CONTACT

HICRO PROCESSOR '4 I D 6 O I L TEMP

F IELD OFF LIGHT BTB ISOLATE LIGHT DRIVE L I6HT M B CLOSE L TRIP BTB CLOSE L TRIP

GCR I U X CONTACT 6CR CLOSEITRIP d

ISOLATIOHt SERIAL DATA BUS

STORAGE (UMI) DATA L I M K TO lFRMl BPCU

I - - 1 // -- Fig. 34 GCU SCHEMATIC

115V AC KHlEI AIRlGNb SYITCH U I D CDWTACT P O f l l I O N EXTERNAL PWER INTERLOCK CURRENT TRANSFORMER SENSING AUTOUND WYER TRANSFER

---+UTILITY BUS RELAY

111s Irslmxrlws FAULT ISOLATION BITE .

Fig. 35 BUS CONTROL UNIT

The Bus Power Control Unit (BPCU) contains all the circuitry necessary for external power monitoring and protection, load shedding* on the utility and galley buses, tie bus differential protection, and control of the external power contactor (EPC), ground handling relay and ground service relays.

* Load shedding reduces the demand when it is likely to be greater then the generatorls can supply and is automatic in operation. I t is carried out on non essential services such as galleys etc.

Each GCU and BPClJ has built-in test equipment (BITE) with self-check and fault diagnosis capability. The BITE display and operating controls are mounted on the 'BPCU.

The GCB7s, BTB7s and APB are identical circuit breakers, the main contacts allow electrical power source to the main load bus or ac tie bus. The circuit breakers are of the latched7 type, ie when the close coil is energised, the circuit breaker closes, the permanent magnet provides the closed contact holding force, ie latches the circuit breaker in the closed position-;1 When a trip signal-is-applied,,the internal spring-assist s the eleAromagnetic field of the coil in breaking the magnetic latch:yThe, two

\ zenek diodes across the coil are'used to sbppress arcing df the contacts 1 I when the breaker is opening or8closing.J There are two of thkm since1

currCn)t flows through the coil ih one di'rqction for trippink &d the I oppdsite direction for closing.I '-/

' 1 /' 1 '-, I ( _ , I I 1 r-\, 1, I /'

Fig. 36 TYPICAL CIRCUIT BREAKER

The utility bus relay (UBR) connects utility busloads to the main generator bus. The electrical load control units (ELCUYs) connect aircraft galley loads and electrically driven hydraulic pumps to the main generator buses.

moodutl w67

The APU generator provides 115/200V 3 phase 400Hz either in-flight or on the ground, controlled and monitored by its own GCU.

FAULT PROTECTION

The GCU monitors the main generating channel, in the event of system faults it trips the GCR, which then trips the GCB. Once tripped there must be a reset procedure, which is to switch off the generator control switch and then switch it on again.

Typical fault protection circuits are:

(i) OVERVOLTAGE (130V) after an inverse time delay, trips the GCR and the GCB trips.

(ii) UNDERVOLTAGE (100V) after a time delay trips GCR and GCB. To prevent the GCR tripping on run-down (non-fault conditions) this is inhibited by under-f~equency/under-speed protection-circuits.

-

I - - \

(iii) 'UNDER-FREQUENCY (365Hz) after an inverse tim6 delay trips GCR and GCB trips. Inhibited by under-speed protectidn kircuit on run down (non-fault conditions). Note. Some aircraft not using a speed sensor on the IDG, have 'a frequency sensing circuit in the ,GCU and this does not trip the GCR it trips the GCB direct.

,

(iv) OVER-FREQUENCY (430Hz) trips the GCR and GCB trips. Again on some aircraft not using a speed sensor on the IDG, using a frequency sensing in the GCU, it trips the GCB direct.

(v) OPEN PHASE - typicaly lowest phase 6 amps and next lowest phase 40 amps, after a time delay trips the GCR and GCB trips.

(vi) OVERCURRENT - If the current drawn from generator exceeds a set value, trips GCR and GCB trips.

(vii) SHORTED PMG - Any permanent magnet generator winding shorted, after a time delay trips GCR and GCB trips.

(viii) UNDER-SPEED - IDG input falls below a set value trips GCB direct.

(ix) SHORTED ROTATING DIODE - Any rotating diode on the generator shorted, after a time delay trips GCR and GCB trips.

(x) DIFFERENTIAL PROTECTION - line to line, line to line to line, and line to earth faults between generator and busbar are detected by this circuit, ie any feeder fault. Compares the current going to and leaving the generator.

There are three current transformers (CT's), one in each phase line, connected to the star point of the generator. These may be external to the generator, or integral to the generator. A further three current transformers, one in each phase line, are downstream of the generator busbar.

PHASE LEADS FROH

CONNECTOR

I I Fig. 37 -CURRENT \ TRANSFORMER- I -.

The 'current transformer outputs are fed to a differential Ptect ion circuit within, the GCU. 1 1 I , '

I I

, / 1 I _ / Q 1 i ~ i ~ u r e 38 shows the principle of o p e T ~ o n using one phaseLline: ~ , d c h phase line is identical and there are thr\&relays in the GCU-'

I !? 1 I I / 1 ,

i i 1 ' i / ; i /' /

I STAR POINT ; I j 1 icuRE6JT ;

8 1

/ : \-.-_---

T R A N S F ~ E R S - .. . . . .- . . . - -- - . .. ../'

- - +"L.-o: - - -GCR then GCB

.... EARTH FAULT....-" _---' ----. ---.-..--... _.--

Fig. 38 PRINCIPLE OF DIFFERENTIAL CURRENT PROTECTION

Operation

Under no fault conditions the current sensed by the load CT's and the star point CT's will be the same. Current flows through the loads and aircraft structure and through the star point CT's to the generator.

- 31 -

rnoodull l A-769

The CT outputs are equal and opposite. When there is an earth fault (as shown in the drawing, the star point CT's have the load current and the fault current, (which flows through the aircraft structure through the star point CT's to the generator). So the star point CT's sense load and fault currents, the load CT's sense only load current.

When the fault current is typically 20 - 40 Amps then the star point CT output will be higher than the load CT and sufficient current is fed to the relay to energise it and signal the GCR to trip and hence trip the GCB.

BRUSHLESS ac GENERATOR

Fig. 39 GCU/BRUSHLESS GENERATOR RELATIONSHIP

Figure 39 shows the relationship between the generator and the GCU. Note the inputs and outputs to the GCU to include load shedding, differential current protection, voltage control, bearing and diode condition monitoring and communication with other GCUs.

- 32 - moodull lA-770

Manual Tripping of the GCR

Typically there are three actions which will trip the GCR, trip the GCB and disconnect the generator from the busbar. They are:

(i) Switching the generator 'OFF'. (ii) Operating the CSDU disconnect switch. (iii) Fulling the fire handle.

Figure 40 shows the electrical power distribution system for a passenger carrying aircraft.

- - - -. . -.

CAPT FLT INSTR BUS HOT BATTERY BUS

Fig. 40 TYPICAL AIRCRAFT DISTRIBUTION SYSTEM

We have discussed the ac generation control and also dc generation from the TRU's. Emergency ac power can be supplied from the static invertor, the Hydraulic Motor Generator (HMG) and a Ram Air Turbine (RAT). These will be described later.

It is important to note the function of two of the busbars, the ground handling busbar and the Ground Service Busbar (GSB). The ground handling busbar is .not connected to the main ac buses, it is supplied from external power or the APU. Relays in the BPCU control the coil of the ground-handling relay. This busbar feeds circuits such as cargo and service lights, and cargo doors (figure 4 1).

rnoodult 123n

The GSB supplies power to in-flight loads and can provide power on the ground for aircraft servicing operations. The GSB is energised from either external power, APU generator or the right main bus. Control relays are provided in the BPCU to operate an external ground service select relay and an external ground service transfer relay. This busbar feeds main and APU battery chargers and interior lights. For full list see the table below which shows typical circuits fed from all the ac busbars.

TABLE OF AC BUS SERVICES

AC 1 15V AC LEFT MAIN BUS

HYD PUMPS OVERRIDE FUEL BOAST PUMP EICAS, LEFT EFlS GRD PROX WARN WEATHER RADAR L ILS & RA DFDR L FCC, TMC & SERVOS MCDP

115VACLEFTTRANSFERBUS CAPT'S PHASE PlTOT HEAD HEAT AC STANDBY BUS PWR SOURCE -

L ENGPROBE HEAT, L AOA PROBE HEAT L DME, ATC, IRS, FMCS CMPTR & CDU

28V AC TRANSFER BUS LEFT L & R FLAP POS IND RUDDER TRIM POS

11 5V AC LEFT UTILITY BUS I I CARGO FAS, HEATERS

- - PASS READING LIGHTS - LAV WATER HEATERS AUS HEATERS -,

11 5V A CENTRE BUS C IR'S, RA & FCC

28V LEFT AC BUS CONT POSIT SENSORS & IND CHART LIGHTS

1 1 5V AC STANDBY BUS L HF COMM, L RDMl L VOR MKR 8 L ADC R ADF, C ILS L CSEU - 1U2L, L PROBE HT IND WARN ELECTRONICS B STDBY INSTR LIGHTS

11 5V AC CAPT'S FLIGHT INSTR TRANS BUS L IAS MACH, L ADC AIR DATA INSTR CENTRE EFlS CAPT'S INSTR LIGHTS L EFIS, ALT, AD1 & VSI

11 5V AC GND HANDLING BUS CARGO, SERVICE LIGHTS CARGO DRIVE & DOORS WATER LlNE HEAT

BUSES 11 5V AC GND SERVICE BUS

MAIN BATT CHARGER APU BATT CHARGER INTERIOR LIGHTS SERVICE OUTLETS ANTI COLLISION LIGHTS WING ILLUM LIGHTS POSITION LIGHTS POT WATER COMPRESSOR FUEL BOOST PUMP HYD QTY EXHAUST FAN WATER LINE HEAT SIDE WALL LIGHTS -

28V AC GND SERVICE BUS 1 EQUIPMENT CTR LIGH JS INTERIOR LIGHTS, PASS SIGNS NLG SERVICE LIGHTS ,

115V AC RIGHT MAIN BUS HYD PUMPS OVERRIDE FUEL BOOST PUMP- EICAS, R ADC, AIR DATA INSTR R EFIS, RDMI & VSI R ADF, DME, VOR, ILS & IRS R RA, ATC & FMC, R HF COMM R FCC & SERVOS

115V AC RIGHT TRANSFER BUS R ENG PROB HEAT R AUX PlTOT HEAT, R ADF

28V AC TRANSFER BUS RIGHT EMER NIGHT LIGHTS & WORK LIGHTS LAV DOME OCCUPY LIGHTS

115V AC RIGHT UTILITY BUS CARGO, RECIRC FANS PASS READ LIGHTS LAV WATER HEATERS, AUS HEATERS TAPE REPRODUCER

28V AC RIGHT BUS DOORS CONTROL INDR HYD OIL PRESSURE MAP LIGHTS FLOOD LIGHTS

Figure 4 1 shows the full busbar layouts of the ac and dc generation systems for a large aircraft. Note that the left and right main buses, left and right hand transfer buses and the ground service bus also feed to 28V ac buses. These will be fed via auto transformers, dropping the 115V bus voltage to 28V. Note also that the 115V ac ground hardling bus feeds a TRU to convert the 1 15V ac to 28V dc for the dc ground handling bus. It is not important you know the circuit in detail but you should have a good working knowledge of the system as a whole.

Fig. 41 TYPICAL ELECTRICAL SUPPLY SYSTEM

LOAD

D

llPlCAL t L N I O I M l l C

PmVIIIWI FOR CWIECTIU6 t l l t M L cr

D :ow;;;,;;,;;tar*rL

Fig. 42 ELECTRICAL LOAD CONTROL UNIT

Electrical Load Control Unit (ELCU)(Figure 42)

As mentioned earlier this unit connects the aircraft galley load and electrically driven hydraulic pumps to the main generator. The galleys require a large amount of power and the ELCU not only contains a three phase main contactor but consists of integral current transformers and sensing circuits for over-current, phase unbalance current, differential current, anti-cycle and lockout protection.

AC Load Shedding

Electrical load management/load shedding is to ensure that electrical loading on the generators stays within limits during both normal and abnormal aircraft operating conditions. To achieve this load reduction is carried out automatically during high demand periods. This load reduction is achieved by selectively de-energising non-essential loads and buses, as required, during aircraft conditions where the system is overloaded or where there is a high probability that normal procedures such as prior to, and a t engine start, will cause a n overload.

1

1 I

L AC BUS R AC BUS 1 I I BTB BTB 1

L UTILITY BUS &I L AFT GALLEY - 48 23 I , ([ORVARD GALLEY - 1s)

R U T I L I T Y BUS

Fig. 43 LOAD SHEDDING - GENERAL

System load shedding is controlled primarily through the BPCU as shown in figure 43 which trips the UBR and ELCU's. The BPCU monitors overload information from all main power sources as well as all main power breaker positions. The BPCU contains the logic for load shedding in the event of a system overload or generator loss in flight. Figures 44 to 47 show the action that occurs with each of the conditions indicated. Each figure is self-explanatory.

A L L GALLEY ELCU'S A L L U B R ' S

AUTOMATICALLY W I T H RESTORATION OF ANY TUO POUER SOURCES

I L L U M I N A T E D DURING GENERATOR OUT C O N D I T I O N AND EXTINGFISHED, A F T E R RESET

I I

I I / (G j , , I 1 / 1- ' 90 KVA FOR 4.5 M I H l 1 1 ~ . 5 KVA FOR 4!0 SECONDS 1 i - - -- I

7 X F T U B R P L E F T GALLEY ELCU

RIGHT BUS U M K E OVERLOAD 90 KVA FOR 4.5 R I N 112.5 KVA FOR 4.0 SECONDS

I R E RIGHT UBR FUD GALLEY ELCU'S

MANUAL RESET CYCLE L E F T U T I L I T Y BUS SWITCH OFF-ON'

B T B (R AC BUS)

L U T I L I T Y )

ELECT CONT PANEL ''.\ ''\

; ' i \ '

pEq R U T I L I T Y )

FUD GALLEY )

b E l C A S ADVISORY MESSAGES L (A) U T I L BUS OFF

INF~IGHT GENEI~ATOR LOSS i Fig. 44 LOAD SHEDDING - 1 I

I '

- I ~ & x 9 L Ac Bus

b E I C A S ADVISORY L (R) U T I L BUS

MANUAL RESET CYCLE RIGHT U T I L I T Y BUS SUITCH OFF-ON

ELECT CONT PANEL

IF TWO GENERATORS ARE OPERATING IN FLIGHT OR AT LEAST ONE GENERATOR IS OPERATING ON THE GROUND

Fig. 45 LOAD SHEDDING - 2

MESSAGES OFF

I I ENGINE START PANEL

pW+-= L U T I L I T Y

m-) L AFT GALLEY

FUD GALLEY - T R I P S BOTH UBR'S AND GALLEY ELCU'S

RESET I S AUTOMATIC U I T H GCB CLOSURE

SEtOHD.ENGINESTA4T TRIPS UBR AND GALLEY ELCU FOR ENGINE BEING STARTED

RESET I S AUTOMATIC U I T H GCB "OFF" I L L U M I N A T E S CLOSURE DURING ENGINE START

-- -

ENGINE START - THE ELEcTRIcAL LOAD ON THE APu GENERATORIS REDUCED T o ENSURE A PROPER AIR SUPPLY TO THE MAIN ENGINES DURING STARTING

Fig. 46 :LOAD SHEDDING - 3 I

OYERLOAD 9 0 KVA FOR 4.5 N I N

112 KYA FOR 4 SEC

-tp --- TRIPS - UBR'S AND RELAY

OYERLOllD 9 0 KVA FOR 3 SEC

Aum TRIPS ALL GALLEY - GALLEY ELCU'S ELCU'S

-CYCLING EITHER BTB SYITCH OFF-ON FOR CONDITION ~ E I C A S ADVISORY MESSAGES

L ( R ) UTlL BUS OFF

ON THE GROUND WlTH EITHER EXTERNAL POWER OR THE APU GENERATOR SUPPLYING THE LOADS. TWO CONDITIONS ARE CONSIDERED - ONE WITHOUT

THE ELECTRIC HYDRAULIC PUMPS RUNNING AND THE OTHER WlTH THE ELECTRIC HYDRAULIC PUMPS RUNNING

Fig. 47 LOAD SHEDDING - 4

Built In Test Equipment (BITE)

BITE is common on all types of aircraft and the following shows a n example of BITE on a large passenger carrying aircraft.

The GCU's and BPCU are individually responsible for isolating faults and storing the results in their Non Volatile Memory (NVM).

In this example there is an alphanumeric display located on the BPCU, which is a 24 character readout that displays messages describing what faults have occurred and which area of the system contains the problem. BITE defines which LRU has failed or if a failure has occurred in the wiring or sensors associated with the system.

The BITE tests can be performed with the aircraft completely powered or only with the battery switch 'ON'.

On the front of the BPCU are three switches, BIT (Built In Test), PERIODIC an-d-RESET. -.. I---~ L-

L ' I------., ' '.

md,ssage~,stqred ', in

, , , ,

The fault rhessage is triljs h s i a t d s .---/

is 1 1 H~splayed for ?K is displayed

I I

' i i r' -1 --

AUXILIARY POUER GENERATOR CONTROL

DATA LINK

LEFT GENERATOR CONTROL UNIT

RIGHT GENERATOR CONTROL UNIT

BUS POUER CONTROL UNIT

Fig. 48 AC GENERATION BITE - GENERAL ARRANGEMENT

Fault data from previous flights is retrieved by pushing the BIT switch during the 15 seconds the 'FOR PREVIOUS FLT PUSH NOW' message is displayed. Previous flight data is retrievable for up to six flights. For the messages that identify the GCU and BPCU as failed, a hexadecimal code number is displayed.

EXTERNAL PUR SYSTEH BUS WEI CCUTIOL UII

LEFT GEN POUER SYSTEM

OK

RIGHT GEN POUER SYSTEM

OK

APU GEN POUER SYSTEM

OK

LAST FLT 0 0 END OF DATA '

FOR PREVIOUS F L T PUSH NOW

NO SYSTEM PROBLEMS

--

EXTERNAL PUR SYSTEH

OK

LEFT GEN POUER SYSTEM

UNDER FREQ T R I P

I D G I P H G I U I R I N G

RIGHT GEW POWER SYSTEM

OK

APU GEN POWER SYSTEM

OK

LAST F L T 00 END OF DATA

FOR PREVIOUS F L T PUSH NOW

LEFT CHANNEL UNDERFREPUENCY T R I P

Fig. 49 TYPICAL BITE, DISPLAYS - BIT I

/

Pressing the periodic test is normally,performed at scheduled aircrkft -

maintenance checks. I -

1 I

I I I ! I

I ' m , I

Pushing and releasing the PEAODIC tdst switch starts the maintenance BITE test, which is a limited end-to-end test of the GCU and BPCU. The results of the test are stored in the NVM. When the test is complete and stored in the NVM, the contents of the NVM for that fligh-t are displayed.

( O I I EXTERNAL PUR S Y S T E V r771 L E F T GEN POUER SYSTEM

OK

R I G H T GEN POWER SYSTEM

GCU F A I L E D CODE 59

APU GEN POUER SYSTEM

OK

L A S T F L T 00 END OF DATA

, FOR PREVIOUS F L T PUSH NOW I I

P E R I O D I C TEST D I S P L A Y SEQUENCE

EXTERNAL PWR SYSTEM

L E F T GEN POUER SYSTEM

R I G H T GEN POUER SYSTEM

APU GEN POWER SYSTEM

RESET D I S P L A Y SEQUENCE

Fig. 5Q TYPICAL BITE DESPLAY - PERIODIC TEST

The NVM contents are the maintenance test results plus any faults detected for the last flight. For the messages that identify the GCU or BPCU has failed, a hexadecimal code number is displayed. The PERIODIC test switch can retrieve previous flight data in the same manner as the BIT switch.

The RESET switch clears the BPCU and GCU memories each time it is pushed. This action makes previous memory entries inaccessible. Each time the switch is pushed, the display will state that the BPCU, left, right and APU GCU memories were cleared when the power system name appears.

Flight Deck Indications

Fig. 51 EICAS DISPLAY - ELEC SYSTEM

The following is a description of the indications that are available on the flight deck of a modern aircraft - in this case based on a Boeing aircraft. Older aircraft will have electro-mechanical gauges giving indications of frequency, current and voltpge of-each generator. There-will also be gauges

\

foFbhefindication or TRU o ~ t ~ u t ; t e m ~ e % ~ t u r e etc. ' 1 - \

I I ' - I 1

I I

Generator load is displayed on the EICAS ELEC/HYD maintenance page for scale values between 0/0.5 and 1.5 (1 .OO = 90KVA). Also displayed on this page is the external power load. For each main power source the voltage and frequency is displayed. The single phase output of the static invertor is also displayed.

I I I

1 I

\

I 1 : 1 I !

/ , ' - 1

Before we look at a system, the requirements for paralleling two ac generators must be considered.

Before two ac generators can be connected together onto a common bus the following conditions must apply:

I I

f 1 , I#*? L C n lllt PYR

LMO - 0.78 0.85, 0.00 0.00 AE-v 1 115 120",l25 0 0 FREE 1 401 402 398 ' 0 0 be-r 10 67 48 0 PC-v 28 28 27 ' 28 IDC OUT 101 103 lD6

I 10 12 I

i -- L ,/c / I

nlvu PRESS LOU MOM' HIKH H1D ~ ~ _ - 0 9 C - 1 < 0 0 0.99 HID PRESS 3238 3210 2140 HID IV~P 50 47 115

k J


i I

, L--

1 i-- 1 /

I I I I

i , -

-- -

a) They must have the same output voltage.

b) They must be operating at the same frequency (speed).

c) They must be in-phase with each other.

d) The 'Phase Rotation' of multi-phase machines must be the same.

Figure 52 shows two single-phase ac generators about to be linked together by the closing of the contactor. (The same principles will apply to multi-phase machines). If any of a, b or c, above, are not being complied with, the result will be a voltage appearing across the contactor contacts. In each case (or in any combination of cases) the voltage will appear as an ac voltage at a frequency that is known as the %eat' frequency.

LAMP

GENERATOR 2 -

I

1

-

' I

Fig. 52 PARALLELING GENERATORS - SIMPLIFAED CIRCUIT

This beat frequency will increase or decrease depending on how far 'out' the generators are with each other. If a lamp is connected across each set of contacts, as shown, then the lamps will go on and off a t a rate determined by the beat frequency. As the generators drift into phase with each other, for instance, the lamps will dim and then glow brightly as the generators drift apart again. The correct time to close the contactor is when the lamps are out. This method of telling when conditions are right for paralleling is known as the Lamps Dark Method. There is also a Lamps Bright Method (cross-connection of the lamps) but it is the Lamps Dark Method which is the basis of all aircraft paralleling systems whether manual, semi-automatic or automatic.

Any attempt to parallel ac generators, without meeting the conditions stated above, results in a large circulating current between the generators as the contactor closes. This circulating current will pull the two rotors 'into line' and paralleling will occur, but there is a very real possibility of loss of power and damage being done to generators and drives.

This system has a built-in safeguard which prevents paralleling if the frequencies of the two generators are far apart. If that is the case, the lamps will be going on and off at such a rate that it will not be possible to close the contactor during a dark period. (In fact, the lamp will probably be on all the time as the time-off period is so short that the lamp has no time to cool and loose is luminescence).

The same thing applies if the generators are at the same speed but are a long way out of phase with each other. It is only when they are very close together that the lamp will be going on and off at a rate that is slow enough for manual paralleling to be put into effect.

The operator must be wary if the lamp is continuously O N or continuously OFF. In such a case, it is only necessary to switch O N (or OFF) a heavy load on either of the generators. This will be sufficient to alter the CSDU momentarily and cause the lamp(s) to start flashing on and off.

Anotfier-w~y is to adjust the Engineer's-Frequency Control if provided, but \ this s i l l &bsequently have to-Ble ~ tu rne ,d to its original sgtting:- , ,

i I I i

i I 1 I I I I

I I

Method of ~ a n u a l l ~ l ~ h a l l e l i $ ~ ,3 Phase ~enera tbrs I

I 1 L --, 1 1 1 ,I i i L- -_-.' I With rkfqrence to figure 53, asskmiiiiith&yo 2 generator is connectdd to

the syn&ronising Bus (Synch BUS). No b gtnerator is connected to it's own load busibut is i,@lated from thk Synch bdsljars by it's trip$e$l Bus Tie

I ~ r e a k 4 r (BTB). / ; / ; ,I i 1 : I ! 1 1 ,I , / /I 8 \ ..' , , ; I - ~

7- /' -.,I '--,

LSYNGHBUS

Fig. 53 MANUAL PARALLELING 3 PHASE GENERATORS

rnoodull l A-781

When No 1 is selected for paralleling, the two lamps will each receive supplies from identical phases (A and C) of the two generators. At the correct moment (lamps dark) the BTB is closed and the two generators are paralleled. It is subsequently possible to select and parallel Nos 3 and 4 generators a s required.

A Method of Automatic Paralleling of Generators

In this system the engineer is relieved of the task of closing the GCB when the conditions are right. He still decides whether or not generators are to be paralleled but the actual operation is automatic. This device will close the GCB when:

(a) The Frequency Difference (Beat Frequency) is less than 3-5 Hz.

(b) The phase voltage difference is less than 10V. 1 - - - -

(c) The 'out of phase'angle-diffednce is less thak90". - 8 I

With reference to figure 54, assume that No 2 generator is cofinected to the Synch Bus and the No 1 generator has been selected for paralleling with it. Transformer TI is connected between the C lines of the two generators and

I / therefore receiving the modulated~aveform, or beat frequency, that exists between them. ,

I

, I I

TO CLOSE NO1 GCB -

4-0 ; 0- nav DC

C PHASE OF NO I BUSBAR

C PHASE OF NO 2 BUSBAR

Fig. 54 AUTOMATIC PARALLELING

The output of the transformer is half-wave rectified by diode D 1 and applied to C1 and R1. C1 will charge to the modulated wave and, as the output dies away, will discharge via R1.

The RC time constant of this circuit allows C1 to discharge completely if the beat frequency is less than 4 Hz, but not if it is above that.

At beat frequencies above 4 Hz there is sufficient baselemitter voltage to cause Q1 to conduct and its collector is at almost 'earth' potential. The Zener diode 21 is not 'broken down' and transistor Q2 is not conducting.

If the beat frequency drops to 4 Hz or less, the capacitor C1 will discharge completely and so reduce the voltage across R 1 and the Base/Emitter of Q 1. Q1 will now cease to conduct and it's collector voltage will rise.

This will break down 21 and switch on Q2. The operation of relay RL1 will now close and the GCB and the two generators will be paralleled.

PARALLELED SYSTEM

Figure 55 shows an IDG used in a paralleled system.

The IDG has a variable speed input from the engine gearbox and through a

1'

RIG^ 6 l E F T H?OMUOCp UNnS (VARIABLE DISPLACEMEW PUMPS 6 FIXED DISPLACEMENT MOTORS

V A R W L E SPEED DRNE FROM ENOINE

INWCATOR

Fig. 55 IDG IN A PARALLELED SYSTEM

The governor receives a n oil pressure supply from the charge pump and senses output speed of the planetary differential. The flyweight arms are made of alnico, below these arms is a coil which has an electrical supply from a load controller. This supply influences the position of the governor for Real Load sharing.

FLYWEIGHTS

PRESSURE TO ONE SIDE OF THE CONTROL \ ELECTRO MAGNETIC PISTON

CONTROL

ADJUSTER

URE TO T H t OTHER SIDE OF THE I ] ' TO LOAD

CONTROL PISTON CONTTRC )SUPPLY PRESSURE

I I I

Fig. 56 G O V E ~ O R WITH MAGNETIC TRIM I

I

I

The nominal charge pressure is 250psi as regulated by the charge relief valve. If the oil pressure falls below 140psi, the charge pressure switch closes the DRIVE light in the IDG disconnect switch to code on.

I I

Fig. 57 PARALLELED SYSTEM FOR A LARGE AIRCRAFT

The disconnect system is as previously described, and resetting is only allowed on the ground with the engine shut down.

Figure 57 shows the paralleled electrical power system of a four-engined aircraft.

Note the Ground Handling bus (GH), the Ground Service bus (GS) feeding similar services as described in the non-paralleled system. Each main ac bus feeds the dc bus via TRUs. The GS bus feeds the Battery Chargers (B/C) for main (MN) and APU batteries. Standby ac power is available through a static invertor. The system shown shows two APU's and two external power sources, many aircraft now only have one APU fitted.

Figure 58 shows the ac generation layout.

-- A - - . . ---\

I 1 ' I

6WCH BUB

- -- TRANS

E SS ESS ESS ESS

6 + oq-1Ac :4p' sTBy * c

INVERTER

Fig. 58 AC SUPPLY SYSTEM LAYOUT

Assuming the APU 1 is running and the Auxiliary Power Breaker (APB) is made the Split System Breaker (SSB) will close and with the Bus Tie Breakers closed the APU can feed all the AC busses (AC 1, 2, 3 & 4). The Generator Circuit Breakers (GCB's) will be open; the APB cannot close unless all GCB's are open. If the second APU was fitted then the SSB will open to prevent paralleling of APU's.

A similar sequence would occur if ground power was being used instead of APU power, this time the external power contactor (XPC) would be closed and via the BTB's feed all ac busses. Once again the XPC will not close unless all GCB's are open. If the second external power is switched in the SSB will open.

If engine No 1 is started, assuming the voltage, frequency and phase rotation of its generators output is OK, then selecting closure of the GCB will trip the APB, when that has tripped the GCB will close, connecting No 1 generator to the synch bus and hence to all load busses. Limited loading usually applies.

Starting No 2 engine will allow its GCB to close to parallel the two generators providing the auto-parallel senses that conditions are correct. Sim.ilar action takes place when starting engines 3 and 4; there is an auto-parallel circuit either side of the SSB.

So when all engines are running all GCB's, BTB's and SSB are closed and all the generators are paralleled. -----

- -

Should a fault condition occur such as a &merator failure then its GCB will trip leaving the other three gederators to power the four loaid busses. Even if three generators failed one gen'erator can feed all four load busses but some load shddding will need to take place.

'

-- - I I

If a generator will not load shafe, the BTB *ill trip allowing the genkrator to power its own bus in isolation from the others. If the load b J s shorted then the GCB and BTB will trip isolAting power from the shorted area.

I

If there is a-short on the synch bus two BTB's and SSB win trip, removing power from the shorted area without loss of busses.

In this system then:

a) Attempting to switch any of the 8 power sources onto the bus system will not be effective unless its voltage, frequency and phase rotation is correct.

b) Attempting to parallel drferent types of power will result in the existing power tripping off before the selected power breaker closes.

The two main units in the system are once again the GCU which contains, voltage regulator (action similar to that previously described), reactive load division circuit, generator control relay (GCR), GCB Control, BTB Control, essential relay control, auto-parallel circuits for generators and fault protection circuits.

The other unit is the Bus Power Control Unit (BPCU) or Bus Control Unit (BCU) which contains XPC control, APB control, SSB control, ground handling and ground service relays, SSB Auto Parallel, external power fault protection and a galley power trip signal.

From the diagram note the essential ac bus which feeds radio and flight instrument systems is normally fed from No 4 generator, but a flight deck switch can select any of the other three generators.

Load Sharing

As the generators are paralleled then load sharing must take place. (It would be convenient at this point to do some revision (CAA JAR module 3) on ac power circuits).

POWER-IN AC CIRCUITS -. - \ - -- r -- --

I I ' -',

a) In a purely resistive dircuit, ALL of the current holes work and I POWER is produced. I I

i I

I I l

I i

1

b) 1 In a purely inductive circui< , ,+he current does do ;work m,d NO

\ I - /

I POWER is produced. 7- I I

\\ '\\ I ,' I I \ i I 1 ' il \

) 1 In a,hurely capacit rvq circuit, $hb current does Lo! work and NO 1 POWER is produce?. 1 I

J I '

I ' I

! I / , : lL.,,' i [ ,,/ )'

j L...../'

; 8,

A pr@tical circuit will cont~~esis tanc~, .~inductance and L - dapkqitance, .. and if we take the example of an ac generator supplying aircraft systems (mainly inductance and resistance) then the current will lag in the supply voltage.

Fig. 59 PHASOR DIAGRAM

The phasor diagram shows the current lagging the supply voltage by phase angle 0. From our previous theory, power is only produced in an ac circuit when current and voltage are inphase. So we need to split the current I into its two components as shown.

The component 'in phase' with the voltage also known as the ACTIVE or REAL component and the component at 90" to the voltage known as the QUADRATURE or REACTIVE component.

It is very important to realise that only one current (I) flows in the circuit and this is the current that is measured by an ammeter in the circuit.

The power in a purely resistive ac circuit is found by multiplying together the rms voltage and current. It follows then that in a resistive reactive circuit, power dissipated can be found by multiplying together the voltage and the component of current in phase with it.

IREAL - - Icos0 I-

- -

TRUE~OR REAL POWER = v x I cos ~ l \ w i t t s (w) or KILOW+TTS (kw) I '

This then gives u s the actual power being used by the system. I

The component of the current that-does no work in the system still_flows through the system cables andproduces power which as w'e p o w cancels over one cycle so no net power is produced. I I

PREACTIVE = V x component of current a t 90" - - -

- - V x IREACTIVE

As sin 0 - - IREACTIVE

IREACTIVE - - I sin 0

REACTIVE POWER = V x I x sin 0 VAR OR KVAR

The unit of reactive power is VOLT AMPS REACTIVE (VAR) .

If the supply voltage is multiplied by the current (I) this will give u s the APPARENT POWER being dissipated, we know that this is apparently available but because current and voltage are not in phase then that is not the true power available from the system.

APPARENT POWER = V x I VOLT AMPS (VA) OR KVA

Power Factor

As we have seen we can work out the apparent power of a system in KVA. What we need to know is how much of this available power is producing actual work done in a circuit, ie producing True Power. So the ratio of

TRUE POWER APPARENT POWER

Is called the power factor (pf).

Example

If 4OKVA generator produces a power output of 30kW then the power factor is:

pf = 30 = 0.75 , 40.- - - -- --- -- - - . 1- - \ - \

\ /- - 8 \ , - \ \

So in this case the factor of podei being qskd is 0.75 and tde gene;?tor is producing 0.75 of its output as True Power,,ie producing power in the system. So obviously the higher the power factor the better1 Aircraft ' generation systems are typically 0.75 A . 9 pf . A pf of 1 (unity3 would mean that all of the power produced ia Being-u$d as true power and-the circuit

I 1 I -- ,," must be purely resistive. I ' \, \\, I 1 I I I I I /

pf = TRUE POWER ( T P ~ I

I l

APPARENT 'POWER (AP) ' I

' I

, L -- - ' , i-

I ----'/

I ___I

As TP - - V x I C O S 0 and AP - - v x I then pf - - V x X C O S 0

vxx

pf - - COS 0

So another formula for power factor is that it equals the cosine of the phase angle.

If we look back at the triangle related to impedance

then the cosine 0 = R z

So another formula for power factor is

Summary

The true power is produced when current and voltage are in phase.

TP = V x I x C O S 0 k W RP = V x I x s i n 0 kVAR A P = V x I kVA

- - cos 0

ACTIVITY -

A 200k 40KVA ac generator has an output current of 100 amps a t a phase angle of 30" lagging. Find the:

(a) True power (b) Reactive power -

(c) Power factor I

I I I I

So when a load is switched onto an ac generator it consists-of two components REAL and REACTIVE, so to ensure that each generator shares these two components equally hoo load sharing circuits are required. The principle is described using figures 60 and 61, using two generators in parallel to make it easier to understand. A sensing current transformer is mounted on one phase line of the generator, this being the primary and the secondary being connected across a sensing resistor. Each current. transformer is connected together by a load sensing loop.

Consider the case of two generators sharing the total load equally, ie each carrying 60 amps. To understand the operation of the loop, it is necessary to 'stop' the action a t one moment in time. At this moment the output of each current transformer (CT) is trying to:

a) Push a current around the outside of the loop, through the secondary winding of the other CT.

b) Push a current through its own sensing resistor. c) Push a current through the other sensing resistor (it is in

parallel with it).

- 52 -

moodull l A-790

- i - - - -- -- , .

- - \ ,'- Fig. 60 LOAD-SENSING - 1 L \

I i ,

1 / '\ > 1 ' \

I 1

I I I

If the two generators are sharing the loaq edually assuming ie&h CT /output is 0.5X then the sensing loop c d ~ e n t w h i c h is always the abefage of CT secondab currents in this case is also 0.6~. A s both of the f2~-out@ut/s are ,' trying to drive currents througp their op\po;ite sensing resistofsas well as their o-, then CURRENT IN THE SENSING RdlSISTORS as they

I I 1 1 cancel each othdr out. I 1 1

I I ' i / 1 L/,'

' /

/ I !

-i \

-- -- --

4 25A

Fig. 61 LOAD SENSING - 2

If the load becomes unbalanced such that generator 1 takes 90 amps and generator 2 takes 30 amps then the sensing loop current remains the same

at: 0.25 + .75 = 0.5A. 2

But the current flow through the sensing resistors is no longer equal and opposite and it is the DIFFERENCE current between the CT and the sensing loop current which flows through the sensing resistors. Note they are in opposite directions, so one signal will signal the system to decrease its loading, ie generator 1 and the other signal will signal its system to increase its loading, until they again balance.

Figure 62 shows a REACTIVE LOAD SHARING circuit of four generators in parallel. The sensing resistor is in the reactive load division circuit., which is in the GCU.

c---- 5 . 2 5 5 . 2 5 -

I R E A C T I V E REACTIVE - R E A C T I V E , REACTIVE

LOAD 20 LOAD 70 LOAD 60 I LOAD 60

I /

A L L CONTACTOAS

CLOSED P O S I T I O N (AUX CONTACTS

u I s c u 1 GCU I I :::%ER GCU 3 GCU 4

Fig. 62 REACTIVE LOAD SHARING

The total reactive load on the aircraft is 2 10 amps and there is unequal load sharing. The average current in the sensing loop is

2+7+6+6 = 5.25 amps 4

The difference current between loop and CT output goes through the sensing resistor so in GCUl 5.25 - 2 = 3.25 amps flow. In GCU2 5.25 - 7 = 1.75 amps flow (in the reverse sense) and in GCU 3 and 4 6 - 5.25 = .75 amps flow (in the reverse sense).

So signals from the sensing resistor are fed to the voltage regulator which will increase the generator excitation to increase the reactive loading of IDG 1 and decrease the excitation to IDGs 2, 3 and 4 to decrease the reactive loading in proportion to the sensing currents.

So remember, REACTIVE LOAD SHARING CONTROL MODIFIES GENERATOR EXCITATION

I REAL REAL REAL LOAD 20 LOAD 60 L O A D 60 i

A similar action takes place in the Real Load sharing circuit (figure 63). The sensing resistor this time is in the LOAD CONTROLLER and its signal is fed to the magnetic trim coil on the speed governor on the CSDU. Movement of the governor piston will modify the hydraulic fluid fed to the drive to increase the drive torque of any generator carrying too little real load and decrease the drive torque of any generator carrying too much real load. So in the example the signal will be to increase the drive torque of IDG 1 and decrease the drive torques of IDG 2, 3 and 4.

So remember, REAL LOAD SHARING CONTROL MODIFIES GENERATOR DRIVE TORQUE

Note. In both systems the current transformers are shorted out when the relevant generator is not load sharing by contacts of the GCB or BTB. This ensures that the CT does not overheat with possible burnout.

- 55 -

rnoodull l A-793

FAULT PROTECTION

Most of the fault protection circuits are the same as those in a non- paralleled system. However there are two protection circuits that you will only find in a paralleled system, OVER EXCITATION and UNDER EXCITATION. If in a paralleled system one generator is taking more reactive load than the other generators, then it will have increased excitation with respect to the others. The reactive load sharing CCT should balance u p the loads, if it does not, then the fault could be in the reactive load sharing circuit, or the generators voltage regulator.

E X 1 IPU PWR 1 CEN1

AUX POWER PU CEN 2

m Pmll

CLOSE

-

I

TRIP I

O M 1 GEN 1 'OWER GEN3--- \ GEN4

I I

Fig. 64 TYPICAL CONTROL PANEL

The over excitation circuit will sense this when the excitation reaches a certain level and trip the BTB, if the fault was on the load sharing circuit then, the generator will feed its own busbar in isolation. If the fault was in the voltage regulator then as soon as the BTB trips that over excitation becomes an over voltage fault and the GCR is tripped and then the GCB is tripped to take the generator off line.

A similar action will take place with the under excitation of one generator to a set level, and again trip the BTB, if the fault is an under-voltage fault then the generator will be tripped off-line.

Figure 64 shows the control panel of a four-engined aircraft with a paralleled generation system.

The layout goes from engine (CSD) to GCB (load bus), BTB and synchronising busbar. Note the kW (Real Load) / kVAR (Reactive Load) meters, just to the right of the meters is a switch marked kVARS. The metersnormally read kW but rpressing-tkq switch and they-will-read, kVAR, henciS7pC &crew can monitor tealand-rehctive load sharifig. -1n-the'bottom

1 I \ ' right h , 4 d comer there is a frequency meter and ac voltrne??r; just below is a select panel which you can s e ~ e h relev$ntlsystem display. Fkom this panel GEN TEST can be selected and YTfIG voltagy will appear on the voltmeter.

/' ' 1 i-, I / < , 1' / On later &craft with CRT dispfays;the-electrical power pagk canbe,brought

\ \ up on the EICAS or ECAM systkni, which shpws a coloured ~siiEjplifkd layout 1 I of the system. \ 1 I

1 i 1 ' ' I

I ' I

I ' ' I

I 1 , 1 ' i ,/ I ' 1

Fig. 65 EICAS ELECTRICAL POWER SYNOPTIC PAGE

The electrical power maintenance page can also be brought up to show ac voltages and frequency, and also the loading on each generator indicated as a percentage of the maximum.

moodull lA-795

Fig. 66 ELECTRICAL POWER MAINTENANCE PAGE

--

- -

BITE circuitry is provided in the 1 3 G ~ o ~ t r o l Units (BCU7s) and GCU7s to identify an electrical power system failure. , I

1 ' An example (Boeing) of one test is using the Central Maintenance Computer (CMC) and the Control Display unit-FDU).

BCU 2

Fig. 67 BITE TESTING - BCUIGCU

Ground tests are selected from the CMC menu on the CDU. Selection of electrical system ground tests are made by pressing the line select key next to 24 ELECTRICAL.

If inhibited is displayed above the electrical power generating system EPGS (BCU 1 & 2), the ground test enable page will appear telling you the conditions that must be met before the test can be completed.

0 < E I C A S I I A I N T P A G E S

< G R O U N D T E S T S 0

G R O U N D T E S T S X X l X X

(24 E L E C T R I C A L

(26 F I R E P R O T E C l I O N

f i 3 - b B R O U N D T E S T S l l l X I

0 < E P G S CBCU 1 ) 0 a < E P G S ( B C U 2) I 3

Fig. 68 ELECT POWER GROUND TEST SELECTION 7 - 1 - - -

I 1, , \

7 - I- i- r- -, \, , - - I i i \

By pressing the line select key next to the, &stem to be testedithe test is activated which takes 10 seconqs Whenltl$e test is completed PASS or FAlL i appear's. If the system failed pqessing the line select key ne? to FAIL causes

I - - / / the grdund test messages to appear. The<following information Lappears on the pa&e,er I I N I--, \\, - --

, I

/' I I i

\ I 1 ~a i i ed unit and 1 cause I 1 I i I ' I i I c ~ c ' r h e s s a ~ e nurnbkr)and AT^ $ring diagram number , li - / ~ ~ u i ~ r n e ~ n b m b e r o, f the failed unit I

- - -- EIGAS-ale& message Flight Deck Effect

(-1 < t P S l IL ICU 1 ) -

Fig. 69 EPGS GROUND TEST

- 59 -

moodull l A-797

VARIABLE SPEED CONSTANT FREQUENCY (VSCF) GENERATOR

With this type of generator, the CSDU has been removed and as the name implies, the variable rpm engine input produces an initial 'frequency wild' supply within the generator which is converted electronically to a constant frequency output. The benefits include: weight saving; reduction in direct operating costs; less maintenance; improved reliability, and less stock required (less number of parts compared to the IDG system). The VSCF generator was designed as a one for one replacement for the IDG and does not require any changes to the aircraft wiring or plumbing.

C N L

COOUMG oRAln O I L &* 1WSTllVCTIMl \ SERVICE PORT A1 R PLUG PORT IDEM1 EXHAUST P U T E

Fig. 70 VSCF - GENERAL VIEW

The VSCF generator system produces constant frequency, 3 phase 1 15V ac power. The unit is installed on the front side of the engine accessory gearbox with the input flange mating with the gearbox mounting pad and is installed with a Quick-Attach-Detach (QAD) adapter kit. Typical mass is about 1401bs (63.5 kg). The unit is made up of eight Shop Replaceable Units (SRU's): speed increaser, generator, inverter, ac fllter, dc filter, CT/EMI (Current Transformer and Electromagnetic Interference) module, generator converter control unit (GCCU), and a heat exchanger.

The generator consists of an input "speed increaser" gearbox, a spindle, a stator assembly, a rotor assembly, and a pump for circulating the oil. The speed increaser provides a 2.96: 1 speed ratio between the engine gearbox input shaft and the high-speed generator rotor, which operates at speeds between 13,705 and 26,120rpm.

A short spindle between the generator-input spline and the speed increaser gearbox shaft has a shear section which provides protection to the accessory gearbox in the event of a generator mechanical failure.

The stator assembly consists of three armatures: the main ac stator, the exciter stator, and the permanent magnet generator (PMG) stator. The rotor assembly consists of a shaft, permanent magnet rotor, exciter armature, rotating rectifier, and main dc field. The oil pump is mounted within the generator frame and is driven by a gear on the generator shaft. The oil pump drive gear ratio keeps the maximum speed of the pump below 12,000 rpm. The pump contains rotor-type elements that draw oil out of the sump in the inverter power module.

There are 6 ac terrninals at the top of the generator. TI, T2, and T3 are the power terminals; T4, T5 and T6 are the neutral terminals.

T1 and T4 are for phase A and the leads are colour-coded red. --

\

\ -

-\

T2 &dl ~ 5 a r e for phase C and the-leads\& colour-codedi61uF-. ' , 1 ; \ ', 1 I \

T3 and ~ i 6 are for phase B and ihk leads b e colour-coded ykllow. I 1 1 I 1 I

i ,I 1 ! i The rnh connector is on the tdp ~fthg/~e/nerator and contains -A c i rck?~ for low oil ipdessure indication and bu7lt~iii~te>t\,(~1~) power, oil /temperathe r --- indica~oh, discoqnect, BIT serih bata poi$ dpd alternate DPCT.

!? i I / 1 i I I

There y-9 GCU (p6).

and I -,

F, for I L--- -

the holtage ! '

/

' I regulator legds from

3 PHASE OUTPUT 3 PHASE 4 WlRE

the

INPUT 3 WIRE 3 PHASE 115V

VARIABLE SPEED AC 4 WIRE mHZ

DRNE FROM 1370 TO 3 PHASE PWN 1 15V AC M) KVA POWER

GEARBOX 2545 HZ 270V DC WAVEFORM 400 HZ

Fig. 71 FUNCTIONAL SCHEMATIC

FIELD EXCITATON FROM VOLTAGE REGULATOR


GENERATOR OUTPLn SIGNAL + 8 , .

e POWER SUPPLY FROM PMG

CONTROLS. MONITORS 6 HAS BIT FUNCTION

VOLTAGE SIGNAL 6 INVERTER STATUS

OUTPLiT VSCF OUTPUT TRANSISTOR SIGNAL CONTROL

Operation is as follows (refer to figure 71):

1) The generator converts shaft energy a t variable speeds to three phase electrical power a t 11 5V. The output frequency of the generator depends on the engine speed; it varies from 1370 to 2545 Hz.

2) The dc filter rectifies the generator output to 270V dc. The dc filter uses large capacitors to remove the ripple from the dc link voltage.

3) The inverter uses six large transistors to convert the dc link voltage to a three phase, Pulse Width Modulated (PWM) waveform. Inside the inverter is the neutral forming transformer. This adds a neutral lead to the three-wire output of the transistors. The neutral forming transformer permits the transistors to equally share a load that is not balanced.

4) The ac filter uses capacitors to change the PWM waveform from the inverter to a three phase sinusoid a t 11 5V 400 Hz. -

-

-

I \ -

I

5) The Current ~ransformef. /~lectro Magnetic Interference (CT/EMI) filter monitors the output current of the VSCF. It also removes unwanted

I

signals from the VSCF output. ' --

i ..

6) The generator/converter cqntrol unit (GCCU) uses a rni'cro~'r~drqcessor to bperate. It gets power from the PMG in the generato{ housing. The GCCU has ,two important functions: I

I I I I 2 ,

i

a) The GCCU controls the generator and the inverter. - 1 ) The voltage regulatotin the GCCU sends a signal to the

field of the generator to control the generator output voltage.

2) The GCCU also controls the output transistors in the inverter. It makes them come on and go off a t the correct times to make the output voltage waveform.

b) The GCCU also monitors the operation of the VSCF. 1) The GCCU monitors the dc Link voltage from the dc filter.

It uses this signal to control the generator output voltage. 2) The GCCU monitors the generator output voltage. It uses

this signal as feedback for the voltage regulator. 3) The GCCU monitors the operation of the internal circuitry

of the VSCF. The GCCU looks for failures and stores failure data for use during BIT testing.

4) The GCCU monitors the output of the VSCF through the CT/EMI filter. This information helps the GCCU find failures that are external to the VSCF.

BIT

A BIT feature is included in the VSCF unit, which is designed for two levels of interrogation: flight-line maintenance and shop maintenance. Three possible reasons for performing a BIT check are on ac system malfunction with illumination of a fault light, low residual voltage (5- 10 volts), or a scheduled maintenance check. Flight-line BIT is done using a two-throw, centre off switch on the converter module with BIT fault data displayed by two red LED'S located adjacent to the switch. DC power for the BIT is through the "low oil pressure" light and requires the "LIGHTS" switch in the 'E3RT' position. One LED is labelled "VSCF Fault Detected" and illuminates with the switch in the 'INDICATE position if BIT has detected a fault within the unit (this information would have been stored by setting a latch relay).

This indicator identifies a failed unit and is considered a "go/no-go" indicator. The second LED is labelled "Aircraft Open Phase Detected" and illuminates, again with the switch in the ?NDICATE7 position, if an open phasF fCu1t was detected duringtheprevious flight cycle. iTh&INDICATE' positisn i2 also used to electric'.$lj, ESet\th'e unit if it had Qenrdisconnected via the disconnect switch. The opposite ;witch position is lpbklled 'LAMP TEST' L d when the switch is in this posiiiqn (normally done fust ) , the LED'S should illuminate. If they do not, a fail9 indicator is likely]

I I 1 --

I L -,/'

Fig. 72 VSCF BIT

- 63 - moodull lA-801

The second level of BIT is intended for the shop technician performed on a removed unit. This level uses information stored in non-volatile memory for the previous 20 flight cycles (defined by engine run cycle-PMG speed).

(

Relative system performance parameters and protection trips/faults are stored. The information is accessed via an RS232 serial data port (pins 17, 18, 19 on the main connector) by any "IBM compatible" personal computer. This level of BIT, used in conjunction with common shop test equipment, will identify a failed SRU or sub-SRU with a minimum accuracy of 95Oh1 without using a rotating drive-stand. A trouble-shooting program for the computer will be supplied to supplement the component maintenance manual for this test.

- 64 -

rnoodull l A-802

CONTENTS

Page

APU Generator 1 Hydraulic Motor Generator 3 Ram Air Turbines 4 Back-up Generators 7 Rotary Invertor 7

Voltage Control 9 Frequency Control 10 Torque Switches 13

y Normal Operation 13 Failure Operation r-- -- 14

Static ~ ~ v ~ r - k o r c 'h14 operation I I [ .\I f 4

Extern$/, Ground Power I I i 16 q C External Power I I 1b AC lExternal Power -- , ' f7

~ransfdrrhers -- -,:<19 ~ u h e n t Transformer I

1 1 '\ 12

Operation ,-\ I ' I

'12 0 ~ u t o Transformer 1 ,

1 221

~ect i f idrs / 1 ' I 29 -sin&1ea~hase Full Wave Rectifier-' ,2 5 BFidwRiZtifier - 26 Three Phase Half Wave Rectification 27 Three Phase Full Wave Rectification 27

Circuit Protection 29 Fuses 29 Circuit Breakers 32 Reverse Current Relays 34 Reverse Current Circuit Breakers 34

EMERGENCY POWER GENERATION

The emergency power generation for an aircraft with dc generators must be capable of, according to AWN 8 1, maintaining an adequate supply automatically to a suitable bank and pitch indicator for a minimum period of 30 minutes. This may be achieved by, depending on the size of the aircraft:

k Main batteries * Separate emergency batteries * Standby invertor fed from the batteries J; There may also be separate standby batteries for radio

The emergency power generation for an aircraft with ac generators could be:

a) Auxiliary Power Unit (APU) b) Hydraulic Motor Generator (HMG)

- c) Ram Air Turbinee(&%T) --

, - d) Static Invertors , \

-- r-- - I , , e) ~ack-up-~ener+ to r s , ,

I 1 ' 1 '

i I I

/ ,I ~ 1 i I APU ~ d n e r a t o r ----, , ' I I

1 ,

1 -- i / 1 r -I, \

The ougput of this generator is used to suqply the busbars d?ring servicing ' I and ma!inkenance of the aircraft bd the grohdd, so the aircraft play be

1 indeperidknt of grdund support equipment; vowever, it can pe used as an emergency sour6ef of power in flibht, in thd event of failure 04 the main engine gener,afors agyou have already-seen-from the AC Generation Book.

- - - - pH 1 i

The generator components are usually the same as the main engine IDG's and are interchangeable except the APU generator is contained in its own housing (often of cast magnesium alloy) and has a different input spline and mounting flange.

WASTED SHEAR SECTION

\ KEYHOLE LOCATING

OIL INLET PORT ELECTRICAL

CONNECTOR

OIL OUTLET PORT

TERMINAL BLOCK WITH 4 STUDS

LOCATING PINS (3)' \

MOUNTING FLANGE

Fig. 1 APU GENERATOR

The GCU as with the main generators provides automatic control and protection functions, it has BITE with a self-check and system fault diagnosis capability.

The APU being a small gas turbine engine of its own is electronically controlled via the electronic control unit (ECU) which supervises all operations of the APU. The APU speed, and therefore generator speed, is achieved by fuel control and a typical speed on a modern passenger carrying aircraft of the generator driven by the gearbox is 12,000rpm.

The APU can also provide pneumatic output (for cabin conditioning and pressurisation etc) and the electrical power and pneumatic output availability may depend on aircraft height. For example, on the Boeing 757 the APU is certified to start up to 30,00Oft, it is capable of supplying 1 15V ac 3 phase electrical power up to the service ceiling of the aircraft. Pneumatics are available up to an altitude of 17,500 feet. Each aircraft's specification for the APU will obviously vary. The APU is started by its own dedicated battery with its own battery charger. - - -

. / L. --- . I

a , D I G I T A L OATA

----- CONTROL SIGNALS

- POWER FLOU

ELECSYS PANEL

Fig. 2 AIRCRAFT ELECTRICAL SUPPLY - OVERVIEW

With reference to figure 2, assuming the left IDG failed, then the GCB would be tripped. If the APU was then started and selected 'ON' the APB and left BTB would close allowing the APU to feed the left busbar leaving the right IDG to feed the right busbar, its BTB being open (non-paralleled system).

HYDRAULIC MOTOR GENERATOR (HMG)

The HMG system provides a backup source after loss of all generated electrical power. A hydraulic motor supplied in this example by the left hydraulic system drives the generator. The generator has an output of 5KVA 30 120/208V ac a t 400 Hz and 50 ampere, 28V dc power which provides power to the standby system, captains flight instruments, selected navigation, communication, lighting and anti-icing loads when both main ac buses are un-powered in flight.

RETURN TO LEFT HYDRAULIC GENERATOR

HYDRAULIC RESERVOIR Q 0 0 Q MOTOR 4 W CONTROL GENERATOR UNIT

LEFT HYDRAULIC SYSTEH PRESSURE MOTOR GENERATOR

SHUTOFF VALVE

I - - -- -

1

1 ' I Fig. 3 HYDRAULIC MOTOR GENERATOR SYSTEM

1 --__ -- I I

I !

! '1 i I

The gederator portion of the HMG i s compgsed of a permanent magnet generator, excitek,generator and; main genfrdtor I with rotors iniunted on a common shaft,.' The permanent ,magnet.generator supplies 3' phase 800 Hz power-to the GCU for excitation control and protective funcfions! The main generator has two output windings as mentioned previously, one produces ac and the other is full wave rectified to provide dc.

Fig. 4 HMG SCHEMATIC

- 3 -

rnoodulll A 806

An example of HMG Power distribution is shown in figure 5. The ac bus transfer relays connect the left and right ac transfer buses and the captains flight instrument transfer bus to the HMG ac output. The two 1 1 5 / 2 8 V ac single phase auto transformers supply 28V ac loads from the left and right ac transfer buses. A dc contactor connects the hydraulic motor generator dc output to the hot battery bus.

BTB BTB _I C

1

INSTR BUS VOLTAGE

1

BUS RLY

I

- - " XI"" BUS RLI

28V AC BUS

HOTOR

STATIC K10564 HID

INVERTER GEN DC

K104 MAIN Kt09 STBY BAT. RLY - PUR RLY HAIH 7

BATTERY i &

I I I

I

Fig. 5 HMG POWER DISTRIBUTION

RAM AIR TURBINES (RAT)

Figure 6 shows a RAT from a Boeing 777-200. The unit is automatically deployed into the airstream when there is a loss of hydraulic power in the three hydraulic systems. I t can also be deployed from the flight deck if necessary. The ram air drives a turbine via the propellers, constant speed being maintained by governor control and pitch change mechanism controlling the pitch of the blades. The turbine drives a hydraulic pump, to provide an emergency source of hydraulic power for operation of flight controls, also drives a generator to provide an emergency source of dc power which is converted to dc by the TRU7s. The generator capacity is 7.5kVA.

As can be seen from figure 7 the ac output from the RAT is converted to dc and supplied to the captains and first officer's instrument buses. The captains flight instrument bus supplies power to the battery bus 2. The battery bus supplies power to the hot battery bus and the static invertor, which provides the single phase 11 5v 400 Hz for the ac standby bus.

RAT - GENERATOR --

1 1 \

1 I

I - - - - >

m: T H I S SHOWS THE CONDITION WHEN THE RAT GENERATOR SUPPLIES POWER TO THE STANDBY SYSTEM

Fig. 7 STANDBY POWER - GENERAL

Figure 8 shows the electrical system layout of an A320. In the case of simultaneous loss of two engine driven generators the RAT is automatically deployed and the emergency generator driven by a hydraulic circuit from the RAT supplies ac and dc busbars via a third TRU.

DC I IlPP1

T BUS 1 I702 PPI

7- -

I

Fig; 8 i- A320 SYSTEM , ., , \

Some earlier RATS had just purely mechanical governors as ?hewn in figure 9, their Air Driven Generator (ADG) providing a constant frequency output to

I - - the emergency ac buses. - , -

I

I I

Fig. 9 APU WIFUNG SCHEMATIC

BACK-UP GENERATORS

The following is a description of the Boeing 777-200 which is a good example of the sort of back-up generators found on aircraft.

It has two back-up frequency wild generators whose output is fed to a convertor to provide a constant frequency output of 115V 3 phase 400 Hz. The generators are driven through the accessory gearbox and are therefore operating as long as the engines are running.

Figure 10 shows the system. If required the left back-up generator feeds the left transfer bus and the right back-up generator feeds the right transfer bus. If both transfer buses require power the right back-up generator supplies power.

SYSTEMS A R I N C 629 BUSES I 1

t t

----- ANALOG CONTROL

Fig. 10 B777 BACK-UP SYSTEM

INVERTORS

ROTARY INVERTOR

In earlier aircraft, and still used in a number of smaller types of aircraft today, a rotary invertor is used to provide the ac power on an aircraft when the main generating system is dc. One early example of a rotary invertor is shown in figure 1 1.

SLIP RING

FIELD WINDING FAN

I /

I

COMMUTATOR BASE

ARMATURE SUPPORT

- _ Fig. 11 SINGLEARMATUW ROTARY INVERTOR- _

The one shown in figure 11 is &own as a sihgle armature m e and comprises of one adnature rotating in a c o h o n field system. The dc kection is a four- pole 2 d ~ c o r n ~ o u n d type; the st& wou_nd ac winding being lbcated in the

I slots of tlpe armature beneath the dc windings. I 1

Fig. 12 ROTARY INVERTOR WIRING DIAGRAM

- 8 -

1 CARBON PILE M I , I 0

I

moodull 1A-811

--

DC INPUT

/ I ,

CONTROL - -

COIL - * 3 PHASE AC OUTPUT

C 7 *

3 PHASE RECTIFIER '

The ac output is taken from three slip rings and is 3 phase 115V 400 Hz. The dc input being via the commutator a t the other end of the machine.

These machines are used on a fairly constant ac load, so control of the output is achieved by carbon pile control as shown in figure 12.

If the output should rise this will be sensed a t the output, converted to dc by the three phase rectifier and fed to the control coil of the carbon pile regulator, which will de-compress the carbon pile, input current will fall, motor speed will fall and output will return to normal.

The later rotary invertors were basically a dc motor driving an ac generator in one unit.

DRIVE SHAFT I

2 8 V DC 3PHASE -

SUPPLY 115V400 HZ OUTPUT

,

Fig. 13 INVERTOR - GENERAL LAYOUT' ---

The dc motors were generally of the compound type and the ac, generators of the star wound rotating armature type. Most invertors are self-cooled by means of fan assemblies fitted to the rotor shaft. In these invertors two separate control systems are required. , -

--

1) Control of output voltage would be achieved through controlling field strength of the generator.

2) Control of output frequency would be achieved through controlling field strength of the motor and hence motor speed.

Voltage Control

In this type of control the generator excitation field system has two windings: the main excitation field fed from a dc source and a control field fed from the magnetic amplifier. The current in the control field is passed in such a direction as to produce a mmf, which assists the main field.

- 9 -

rnoodull l A-312

1 15V 3 PHASE 400HZ OUTPUT

+ AMPLIFIER GENERATOR SECTION OF + DC INVERTER INPUT

Fig. 14 VOLTAGE CONTROL

The value of the control field c ~ r r e n t is determined by the amount of variation of ac outpiit voltage from the r'equFed-level?, Even at the desired-lev,el there is a standing current flow in the control field. I

1 I I ' I

An incre4se in voltage output 4 1 be sensed by the voltage error sensing circuit. The output of which will takethe,magnetic amplifier away from the saturation level. 1 ., , \ I - '

\ , 1 ipp ' The impebance of the magnetic vplifier will increase causing a reduction in magnetic lamplifikation output to the generator control field. The total field strength of the~generator excitation field , wil1,decrease reducing the generator output' voltage to the desired level. ' /

-- - .- -- -- I

Frequency Control

The motor of an invertor being essentially a shunt machine allows the use of a control field incorporated into the motor main field system in the same manner as for voltage.

The motor control field is so wound that the passage of current through it produces a magnetic affect to assist that of the main rotor field system. By varying the amount of current through the control field the motor speed and thus output frequency can be varied in relation to the variation from the desired level.

3 PHASE AC OUTPUT

MAGNETIC AMPLIFIER

Fig. 15 FREQUENCY CONTROL - 1

1 - FREQUENCY

ERROR

Any increase in frequency output will be detected by the frequency error sensing circuits. The output of which will increase the impedance of the magnetic amplifier causing a reduction in output current to the motor control. We need to look a t the operation of the frequency sensing contro

I

,.... . -..-....-....--....----.-.., SERIES I CONTROL1 ! RESONANT CIRCUIT -

i

j . j MAGNETIC

SENSING + ....................... .........--.................

I I j AMPLIFIER I

1

"""

RESONANT CIRCUIT

li I

4 1 I

I I

'ROL 2

..........

b

;-m i TO

MOTOR i FIELD !

11

.........,

PARALLEL

t i CON1

Fig. 16 FREQUENCY CONTROL - 2

- A

- 11 --

moodull l A-814

...........................-,.......--..-.....-.......---....

I - - I - -

..............

This circuit is fed from the ac output of the invertor. The tuned circuits are resonant at dissimilar frequencies above the nominal value, and the control of the magnetic amplifier depends upon the difference in current in the two control windings.

CURRENT

RESONANCE

100 2 h 300 400 5&0 600 700 800 - --

FREQUENCY (Hz) I

Fig. 17 GRAPH OF 'FREQUENCY AGAINST C U F N T I i I

I I , , I _ '\ I ---, '

\ ',\ ' - -- ' I

At the nominal frequency (400 Hz) the cu6ents in the control &indings are the same, but the? are wound sb that theit' magnetic effects cancel each other out. I , I '

1 - i-

* s

If the-frequency-rises above nimal- (motor speed increases)-more current will flow in the series resonant circuit than in the parallel resonant circuit. A larger current now flows in control winding 1 than in control winding 2 and the unbalance will increase the motor field current, decreasing its speed/ frequency.

If the frequency falls below normal, the parallel circuit increases current through control 2 and the series circuit reduces current through control 1. The magnetic unbalance is now reversed, and the motor field current is reduced, increasing motor speed and restoring frequency to normal.

Efficiency

Invertors have a relatively low efficiency, about 50 to 55% on full load. When lightly loaded their efficiency is even less, which presents a considerable drain on the main electrical system. Some of the systems would have phase B earthed, ie use the airframe as the phase B line.

Torque Switches

These are used as a means of detecting whether the nominal frequency and voltage of the aircraft instrument supply circuit is within its prescribed limits. Used to operate the changeover circuits, from NORMAL to EMERGENCY in the case of invertor circuits, or to operate indicator circuits.

CONTACT BREAKER ARM CONTROL SPRING

CONTACTS

RIGHT HAND CONTACTS

Fig. 18 TORQUE SWITCH

I ,

The switch consists of a three phase motor, coupled to a switching device, by means of a rack and pinion gearassembly, the motor stator being star wound. -

-

The front compartment of the switch unit houses the switch mechanics, which consists of a control spring connected by a gear wheel to the pinion on the rotor shaft, and a set of contact assemblies. The contacts, being adjustable, are set to operate a t a low operating pressure. The contact arm assembly is mounted on the geared shaft, which also carries the control spring. The front compartment of the switch unit has a Perspex (transparent plastic) cover to allow visual observation of the switch operation.

Normal Operation

When the correct voltage and frequency and the correct phase sequence is connected to the terminals the windings are energised and the rotor will turn. This will turn the pinion and the gear wheel until the torsional resistance of the control spring equals the torque exerted by the motor. In this condition the motor will stop. The contact breaker will have opened or closed the contacts at this point.

Failure Operation

Should the supply voltage and frequency decrease, the motor torque is reduced, and the shaft will move under the influence of the control spring to the de-energised position. The contacts will now operate.

STATIC INVERTOR

As the name implies the components are of the solid state type. Some aircraft may use them as a normal source of ac power, but on most modern passenger carrying aircraft they provide an emergency source of ac power in the event of main generation ac failure. The output is single phase 11 5V 400 Hz. There are many different types, but we will look at one example.

- - -- - -- - , - - '\ -

\ !r .--

28 V FILTER DC NETWORK I / \

CONSTANT i CURRENT GENERATOR

I I 400Hz SQUARE -- -

VARIABLE ' PULSE WIDTH

7 @ ODD HARMONIC 1

FILTER' 1 15V 40OHz I - SINGLEPHASE

VOLTAGE SENSOR

CURRENT C C u C n m

OUTPUT

Fig. 19 STATIC INVERTOR PRINCIPLE

Operation

DC is supplied to a filter network, a pulse shaper, a constant current generator, power driver stage and the output stage. After variations in the input have been filtered out, dc is supplied to the square-wave generator, which provides the first stage of the conversion from dc to a c and also establishes the correct frequency of 400 Hz.

- 1 4 -

moodull l A-81 7

This output is then supplied to a pulse shaper circuit, which controls the pulse width of the signal and changes its waveform before it is passed to the power driver stage. The purpose of the turn-on delay is to allow the voltage to stabilise before there is an output to the power driver stage.

The power driver supplies a pulse width modulated output to control the output stage, this signal having a square waveform. The power driver also shorts itself out each time the voltage falls to zero (during notch time).

The output stage also produces a square wave output but of variable pulse width. This output is finally fed to a filter circuit, which reduces the total odd harmonics to produce a sine wave output a t the correct voltage and frequency.

The voltage and current sensors produce a rectified ac feedback signal which controls the notch time of the pulse shaper output to maintain the invertor output within the required limits.

( The standby ac bus is usually the only ac bus which can be maintained in the event of failure of all ac power sources. As shown in the next diagram if ac power fails the invertor will power the ac standby bus. The dc input being powered from the hot battery bus (battery). The standby bus typically feeds engine instrumentation, standby ignition and a compass system.

- - . --

(-) FY, .C I

OR F S S OC FAILS

STANDBY POWER LIGHT (GREEN) \ illuminated when STBY AC BUS is

.

powered from the INVERTER. \ STANDBY MASTER SWITCH OFF - the STBY AC BUS Is dead. NORMAL - STBY AS is powered from ES AC providing

both ESS AC and ESS DC are available. - should either ESS AC or ESS DC fail STBY AC BUS is powered from standby inverter which is powered from dc.

MANUAL - STBY AC powered from inverter ON

Fig. 20 STANDBY POWER CONTROL

EXTERNAL/ GROUND POWER

Some aircraft use power from a ground power unit (GPU) for servicing and maintenance of aircraft systems. So the GPU output must be connected into the aircraft at a convenient point, usually located low down on the side of the fuselage. In the majority of large passenger carrying aircraft the aircraft is completely independent of GPUs as they use the APUs, however should there be an APU failure then connectors are available for dc and ac external power.

DC External Power

Figure 21 shows a dc external supply socket, ie the end of the lead from the GPU and the 3-pin plug, which is on the aircraft. The plug has two positive pins, one being shorter than the other and one long negative pin.

-- AIRCRAFT SKIN , -

r- -' - -\ I

-, '\ I

\

I I I I I

I I 1 ' SOCKET i

I -- ,

P -- ,

I 8 I

Fig. 21' DC GROUND POWER AI~CRAFT CONN&~FION ' ,

L - - , L-

-- - - A - 1 _

Figure 22 shows a simple dc power system. When the power selector switch is placed to 'EXT' the positive supply from the short pin will energise the external power relay, allowing the GPU supply to be fed to the main busbar.

DC BUS BAR

EXTERNAL ? - -

- : POWER ( 0-411 I SELECT

+ *- +

EXTERNAL POWER CONNECTOR

TO BATTERY SYSTEM

Fig. 22 SIMPLIFIED GROUND POWER CIRCUIT

- 1 6 -

rnoodull lA-819

Note that if the socket is withdrawn with the circuit still live' then the external power relay will de-energise before the main pins are disengaged from the socket, this ensures there is no arcing at the main pins on the socket.

AC External Power

The standard ac connector/receptacle pin arrangement is shown in figure 23.

It includes four ac pins, for phases A, B, C and N (neutral or ac ground) and two shorter pins interlock pins E and F. The relays connecting the ac external power to the bus system depend on the external interlock circuit joining the two short pins E and F, with F usually being the return line.

DC INTERLOCK RELAY SUPPLY PINS

PHASE PINS

GROUND PIN I I

- -

Fig. 23 ARRANGEMENT OF AC CONNECTOR PINS

SUPWRT ROD

IHDICATIOH LIGHTS

EXTERNlL WER 1

Fig. 24 EXTERNAL POWER PANEL - EXAMPLE

If the external power lead is inadvertently removed while ac power is still flowing, the E and F pin interlock circuit is broken before the longer ac pins separate from the lead. When the interlock circuit is broken, all relays holding external ac power on the electrical system will relax. Therefore there will be no current flowing, and therefore no arcing when the lead pulls away from the large pins.

Figure 24 shows an example of an ac power cable connected to the external power receptacle. Inside the panel are two indicating lights, a white ac CONNECTED light and a clear PWR NOT IN USE LIGHT.

Figure 25 shows the ac external power system of a passenger carrying aircraft.

EXT PYR

-- -

L - -

\

GROUND ,, / HANDLING c

RELAY I

1

LBTB a- I ,' -

GROUND POWER, T-R UNIT

/ ' I

-

Fig. 25 AC EXTERNAL POWER SYSTEM

The Bus Power Control Unit (BPCU) controls the closing and opening of the External Power Contactor (EPC) allowing external power to be connected to the ac tie bus. Connecting external power of acceptable quality illuminates the AC CONNECTED, NOT IN USE and AVAIL lights on the flight deck. Pressing the EXT PWR switch closes the EPC causing the EXT PWR ON light to illuminate and connects external power to both ac buses via closed bus tie breakers (BTB's) .

The 1 15V ac ground handling bus is automatically powered on the ground if external power is of acceptable quality. The BPCU controls a three-position ground-handling relay and automatically gives priority to external power of the APU generator which is operating.

Note the TRU converts the 11 5V ac to 28V dc for the dc ground handling bus. These busbars are only powered on the ground, feeding such services a s cargo/ service lights, cargo drive/ doors, fuelling valves and control.

The ground service busbar power is supplied to equipment which must be powered in both air and ground modes. In the air this bus is fed from the right ac bus. If this bus is not powered, the ground service switch allows the ground service bus to be powered from either external power or APU.

The BPCU controls, electrical system external power monitoring and protection and BITE. The unit monitors the quality of power applied to the aircraft, before the external power contactor (EPC) is allowed to close the BPCU checks for, overlunder voltage, underlover frequency and phase sequence.

Other protective functions while the external power is connected are: ( I ) tie bus differential protection, which isolated any short circuit fault on the

i external power feeder or tie bus, (2) overcurrent, (3) underlover voltage, (4) overlunder frequency, (5) open phase and (6) overload, in each case the EPC is tripped.

GRWWD HUIOLIffi R E U Y GROUlD IERVtCE SELECT R E U Y 6RCW0 S E l I C E TRANSFER R E U Y EXTEIWAL m E a , c o m r ~ c T a a FLIGHT c m p h n m r n r IIIDICLTIWS ELECTRICAL LOAD C M R O L M I I S U L L E V U)bb R E U Y S (OPTICAIL) U T I L I T Y BUS I E U Y

DATA W E S TOIF1101 S N ' S

Fig. 26 BUS POWER CONTROL UNIT

TRANSFORMERS

CURRENT TRANSFORMER

Within the electrical power system, there are a large number of current transformers used. First it would be useful to re-cap on the theory.

The Current Transformer is designed to enable circuit currents to be measured without breaking into the circuit, as is necessary with an ammeter or its shunt. The output of the current transformer may be applied directly to an instrument or be used in control circuits.

It works on the principle of mutual inductance but its construction and mode of operation are different to that of the power transformer. Current transformers use the load's supply cable as the primary winding. The diagram shows the principle.

SECONDARY COIL

- , - - ', - -,

I

I I I I

' i Fig. 27 CURRENT TRAN$ORMER , PRINCIPLE - - -. \

, I ' i t operation ! 1 I

I I / 1 ! I

Wheli the load passes througl! the suppiy cable, it creates la magnetic field along it=hole - length which-is const&$i building-up, collapsing, reversing, building-up etc. ItpiSth-i~fluX which induces emfsihto the coils of the secondary winding.

As the ring former and secondary coil only take up a very small length of the primary cable, it is obvious that, whatever happens a t the secondary, the effect on the primary will be virtually nil. The primary, which depends on the load, may therefore be regarded as a constant current/constant flux supply.

The voltage in the secondary winding causes a current to flow through its load and through the secondary winding. This produces a secondary flux which opposes the primary flux and so keeps the corepu to a very low level. This is a most important point to remember because, ifthe primary is operated with the secondary winding disconnected from its load there will be no secondary emf to oppose the prima ry emf:

This will result in a high core flux; increased eddy currents in the core; increased voltages in the individual secondary coils and overheating.

The result is that the current transformer will burn out. (Even if the mistake is realised and the system is switched off before it actually burns out, the core may be pre-magnetised, or biased, and cannot therefore be relied upon to be accurate).

If it is necessary to operate the primary when the secondary load is disconnected, short together the secondary terminals. This will cause a secondary current and f lu and so keep the core flux to a minimum.

If the current transformer is supplying a load such as an ammeter, then the actual connections may not matter and the ammeter will indicate whichever way it is connected-up. This is not true, however, when the current transformer i s feeding signals into control circuits, where it is essential to get the phasing right.

If the secondary connections are crossed, the output will be turned through 180" causing untold havoc in the control circuit. It is absolutely

\ essential to get the secondary_connections correct. In the same way, if a current transformer is being fitted over its primary cable,-it is absolutely essential to physically position'it the right way round.

SUMMARY 1

1. When fitting a current transformer, ensure it is fitted THERIGHT WAY ROUND.

2. When connecting the secondary to its load, CONNECT IT CORRECTLY.

3 . NEVER operate the primary circuit with the secondary open- circuited --------- IT WILL SHORT IT OUT

- -

4. NEVER operate a current transformer on anything other than its DESIGNED LOAD.

5. In some cases, the current transformer and its load are a matched pair. (They may even carry the same serial numbers). If one is changed, then the other must also be changed.

6. When they are used in control circuitry, remember that the secondary output is a supply source proportional to the primary current flow.

Figure 28 shows a current transformer as we have seen used in real and reactive load division circuits, sensing for open phase and differential current protection circuits.

SE LEADS FROM ERATOR TERMINAL

\- CONNECTOR

Fig. 28 CURRENT TRANSFORMER

The current transformer assembly shown comprises three toroidal current transformer sections in a single package. They are rated a t 250 amperes primary current and have a primary to secondary current ratio of 1000: 1. Each tryisformer consists of-1000 tuTns,of number 28-wire woiind,on a

I-- , toroidal'core and are capable of 9perating.over a frequency r ~ g e - f r o m 350 to 440 Hz. As indicated in thd theory the current transforher must dever be left open circuited so usually auxiliary contacts of the Bus Tie ~realkers (BTB's) or Generator Breakersl(GB's) short them out when tlde system is

L-. - - , not being used. / I

-- i -- ' \ \ 1 _ _ - ' \ ,

AUTO TRANSFORMER I

' I

I

These A-usdd extensively f q pbwer distribution in a modeA-aircraft, - but again a-re-cap on the theoryLwill help. ,

STEP-UP SECONDARY S,

LOAD

, STEP-DOWN

6 S2 B

Fig. 29 AUTO TRANSFORTVIER THEORY

LOAD

This is a special type of transformer that has only a SINGLE winding, which serves as both the primary and the secondary. It follows that a portion of the winding is common to both the input and to the output:. It may be used either as a step-up or as a step-down transformer.

If an ac supply is applied to the primary terminals, an alternating current will flow through those coils connected across P1 and P2. This will set u p an alternating flux, which will link with all of the turns on the former, inducing a voltage in each.

The output voltage is therefore that which appears in the coils across terminals S 1 and S2. Loading the secondary will have the same effect as described for the Power Transformer. If the current flow is considered for one particular half-cycle, it will be seen that the primary and secondary currents are opposing each other in the common portion of the winding.

The actual current flow in the common portion is therefore the difference between the two currents. This means that the cross-section area of the copper can be decreased in the common portion, bringing about a saving in weight.

Fig. 30 SYSTEM CIRCUIT

This saving is obviously most beneficial auto-transformers where the input voltage and the output voltage are very close together and the vast majority of the winding is common.

One disadvantage of this type of transformer, especially when used as a step-down is that, in the event of an open-circuit occurring in the common portion of the winding, the input voltage will be applied to the load.

In the power distribution system the auto-transformer is typically used to step the 115V 400 Hz down to 28V 400 Hz ac. Figure 30 shows the layout of a non-paralleled electrical system, note the 28V ac busbars all feed via auto-transformers from 1 15V busbars.

The 28V ac left and right buses feeding position sensors and indicators, hydraulic oil press indication, map, chart and flood lights.

The 28V ac transfer bus right feeding, emergency night-light and work lights.

ipp- -- - <- I \

-- - \ - i- - The 2 8 0 ac transfer bus left feeding flap ahd rudder trim position ' :

' I indication.

I I

I I

The 2 8 ~ ac ground service bus, feeding idterior lights, passebger signs and ,, I i--- / 1 service lights. ---_ .. , ,

I 1 '., \ I -

7 -

The iuto-transformer is also dsdd in individual circuits to s tep down from 1 1 5 ~ / to other values of ac, typically in extei-nal lighting. 1 i 1 I 1 I

, !,

' '-- I - I

Within an electrical power system, as we have seen, rectifiers are used, a re-cap of the theory would help.

Single Phase Half Wave Rectifier

With reference to figure 3 1, when terminal A is positive with respect to B the diode conducts, this causes a current to flow around the circuit and a voltage will be developed across RL. When the input polarity reverses terminal A will be negative with respect to B and the diode will switch off.

The voltage developed across RL is therefore half-sine-waves and is known as a half wave rectifier. The output being dc, albeit variable. The average value being half that of the supply, ie peak x 0.3 18, (assuming no losses). The output dc 'ripples' have a frequency equal to the input frequency of the ac supply, ie ripple frequency = supply frequency.

moodull l A-827

+ Peak voltage v

- 0.318 v

-

Fig. 31 HALF WAVE RECTIFICATION

Single Phase Full Wave Rectifier

As the name implies this uses both half cycles of the input wave form. Figure 32 shows diodes Dl and D2 used with a transformer, which is centre tapped a t C. The point C can be considered as neutral with

A

i terminals A and B swinging alternately positive and negative about it.

When A is positive to C, Diode Dl conducts with D2 switched off. On the other half cycle of input, B is positive to C and D2 conducts with Dl switched off. The output is therefore un-directional, with both diodes alternately conducting, giving a full wave output across RL. The average output voltage is 0.637 x peak (assuming no losses), ie average of the

- - - -

supply -

Peak voltage v

AC Supply -Average voltage - 0.637 v

Fig. 32 FULL WAVE RECTIFICATION

The output dc 'ripple' is therefore twice the input supply frequency. Having to use the double winding on the transformer makes this component more bulky in size and therefore more expensive.

A point to note about this circuit is that when Dl is conducting, the voltage across the load resistor RL is the peak voltage. With D2 cut off the voltage across C-B is in series with this voltage, so these two voltages combine to give a total of twice the peak voltage.

This will act as a reverse voltage across D2 SO the peak inverse voltage for the diodes must be twice the peak voltage on either half of the secondary of the transformer.

Bridge Rectifier

This is also a single phase full wave rectifier, and has advantages over the previous circuit in that the transformer does not need to produce twice the voltage required and the secondary is in use all the time. Unlike the previous circuit where only half the secondary winding was used a t any one time.

Figure 33 shows a bridge rectifier. Assume the top of the secondary winding of the transformer to be positive (positive half cycle), trace the current flow through the load using the arrows shown.

Fig. 33 BRIDGE RE~TIFIER -: FIRST HALF ~ F F L E !

I I

' I

, - / , /

Fig. 34 BRIDGE RECTIFIER - SECOND HALF CYCLE

On the next half cycle (figure 34) assume the bottom of the secondary is positive and trace the circuit through the load following the arrows. Note the direction of current through the load is the same during each half cycle, ie it is dc.

Note that in this circuit the two non-conducting diodes have twice the supply voltage across them, (load/supply voltage + supply voltage = twice supply voltage). However, this voltage is shared between the two non- conducting diodes in series, therefore the peak inverse voltage per diode is the supply voltage. As before the ripple frequency is twice the supply frequency.

Typically all four diodes are available in one package.

Three Phase Half Wave Rectification

In order to obtain three phase half wave rectification a diode must be inserted into each of the supply lines to the load and the return from the load to the supply MUST be to the star point of the three phase system.

Therefore this form of rectification can only be used where there is a star connection using a neutral line. Assume this star connection is the secondary of a three phase (DELTA-STAR) transformer as shown in figure 35.

Figure 36 shows the waveform of the three phase supply and the resultant supply voltage to the load.

Fig. 35 DELTA STAR TRANSFORMER -

o m I VOLTAGE

Fig. 36 WAVEFORMS - THREE PHASE RECTIFIER

Note that the ripple frequency of this rectifier output is three times the supply frequency; with three dc output voltage 'blips' for one sequence of the three phase ac supply.

Three Phase Full Wave Rectification

This form of connection does not require a neutral line, so can be used on either Star or Delta connected systems. Figure 37 shows the diode circuit diagram.

rnoodull l A-830

LINE

LINE

LINE

Fig. 37 FULL WAVE RECTIFIER CIRCUIT

The arrows show the time in the three phase cycle when phase A is maximum and passing peak current ti the load (say 10 amps). After passing through the load, the current splits into two, of five amps each to return to the B and C lines back to the supply.

- - , - - - - - - - - , - - , \ -

The output ripple frequency is six times the supply frequency. \. '

Fig. 38 THREE PHASE FULL WAVE WAVEFORM

The use of rectifiers in the systems we have covered have mainly been three phase rectifier systems, ie in the transformer rectifier unit (TRU) there are star/star and star/delta systems. Within a brushless generator a three phase full wave rectifier system, and in a voltage control system a single phase bridge rectifier system.

- 28 -

rnoodull lA-831

CIRCUIT PROTECTION

Fault conditions such as a short circuit or overload will cause excessive current to flow in that circuit which, if left unchecked, would produce sufficient heat to cause considerable damage to the cables. Damage would not necessarily be confined to a single circuit as the heat generated by the first failure could also burn the insulation of adjacent cables within the loom, causing failure of those circuits and creating a condition for a fire.

It is essential therefore that a means of protection is integrated into electrical circuits. Devices normally used for this purpose are fuses, current limiters, limiting resistors, circuit breakers and reverse current relays/circuit breakers.

The protective device shall be of a type and capacity such as to ensure that it will perform the duty for which it is installed, and should be capable of sensing small sustained overloads, ignore short duration surges, and clear the circuit as quickly as possible on short circuit.

Fuses

This consists of a low melting point fusible element enclosed in a ceramic casing (to protect the element 'and localise any arcing when the-fuse %lows'). The fusible element is made of tin, lead, alloys of ;tin and bismuth, silver and copper. The fuse is in series with the circuit and its primary purpose is to protect the cables of the circuit against the flow of short circuit and overload currents. Under those conditions the fusible ele d interrupts the circuit.

- -

Construction and current ratings are varied in order to provide a suitable selection of protection for specific electrical installations and individual circuit requirements. Such things as, maximum current, supply voltage, breaking capacity time/current characteristic and ambient temperature must be taken into account.

The current rating of a fuse is the current it will carry continuously without deteriorating.

The minimum fusing current is the current, which will cause the fuse to operate under given conditions in a given time, typically four hours.

Fusing factor - - rated minimum fusing current current rating

It follows the fusing factor must exceed one.

- 29 -

moodull I A-832

A fuse providing close excess current protection is one with a fusing factor of 1.5 or less.

A fuse has an inverse timelcurrent characteristic, ie the higher the overload current the faster it will blow.

An example of a light duty fuse and its holder is shown in the next diagram.

SECURlNO NUT FUSE HOLDER

CAP I /

N S E RATING / METAL END / Pa- -- "

FUSE I -. . . . - --. . \.

8 - -- - - \\ -\ r--., I 1

Fig. 39 FUSE & FUSE HOLDER I

1

I I I I I /'

~ e a G duty or High Rupturing CQsity (HRC) fuses are nqrr&ly,us/usdd for such !circuits as main electrical distribkti'on. They consist bf a c e r d i c tube which a.pumber of idehtlcal fusible', elements are cbnbected in parallel to the end contacts. The tube is filled with a pacqng medium of granular magnesite (magnesium oxide), kieselguhr or chalk (calcium carbonate) to damp d o h the explosive effect of the resulting arc whkn-rupturing takes place. --Thepaked tube is sealed at each end with metal end caps, formed into mounting lugs, and fire clay cement.

CERAMIC BODY INERT GRANULES

FUSIBLE INERT ELEMENT GRANULES

CERAMIC BODY \

Fig. 40 HRC FUSES

When a n excessive current is flowing each element will be heated close to its melting point until one element fails, and being connected in parallel, will transfer its share of the load to the remaining elements causing f-urther failures in quick succession. Note the mounting lugs, through which the fuse is connected into the circuit.

Current limiters are used mainly for the protection of heavy duty power distribution circuits and consist of a high melting point single strip of tinned copper shaped a t each end to form mounting lugs. The central portion of the strip is, in some cases, "waisted" to form the fusible area. It is enclosed in a ceramic housing which has a glass or mica window in the side to allow for visual inspection of the fuse element. The timelcurrent characteristic of the device will allow a considerable overload current in the circuit before rupturing occurs, and are therefore useful in circuits carrying occasional current surges, such as starter motor circuits.

GLASS

Fig. 41 CURRENT LIMITER -

The Air Navigation Order lays down that spare fuses, for all electrical circuits, which can be replaced in flight, must be carried. This is 10% of the number of each rating employed or three of each rating employed, whichever is the greater.

Limiting resistors provide another form of circuit protection for circuits which normally have high initial starting current, eg engine starter motor, invertor circuits and circuits containing high capacitive loads. In order, therefore, to keep the initial current surges within a reasonable limit the starting section of the appropriate circuit incorporates a limiting resistor which is shorted out when the current has fallen to a safe level.

Figure 42 shows a limiting resistor in a typical turbine engine starter circuit utilising a time switch for the control of the limiting resistor.

moodull lA-834

SHORTING LIMITING BATTERYRELAY STARTING RELAY RELAY RESISTOR

BREAKER

SWITCH '1 y L - - - - -

Fig. 42 STARTER MOTOR CIRCUIT --\

c- - - ( r - - - . '\

1 - - \ The limiting resistor is in series with th& starter motor, on istart, the back emf (bemr) of the motor is zerd so to prevent overloading 06 the motor

I windings, the initial current to the motor is decreased. , ---A , I

When the starter push switch !is operated, current will flod ffom-the busbar:to energise the main starting relay and close it contact~which in turn cduses theyme switch td oberate, apd supplies currdnt to the starter motor via the liniiting resistor. When the niotor has built hp to sufficient speed, the lime switch energisesl the shorting relay. The current now pass. zrectly to the motor,. the'llm?tjng resistor being shbrted -- A out. Cwrent to the motor is now limited by the bemf of the motor. When the engine reaches 'self-sustaining' speed the starter motor circuit is switched off.

Circuit Breakers

Disadvantages of the fuse are that it takes longer to replace than to reset a breaker and spare fuses have to be carried for in-flight replacement.

All the modem aircraft's circuits are protected by manually operated circuit breakers. The circuit breaker is really a combined fuse and switch unit. It has a mechanical trip device actuated by the heating of a bimetallic element through which the circuit current passes.

All manually operated circuit breakers used on aircraft must be of the "trip free" type, ie they cannot be maintained closed or held against a fault current when any part of a circuit is carrying overload current.

Figure 43 shows a typical manually operated circuit breaker. Manual operation is usually by means of a single push-pull button, pushing to close the contacts, pulling to open.

SHAKE PRYOF WASHER

Fig. 43 CIRCUIT BREAKER

t LOAD

I A SUPPLY

Fig. 44 CIRCUIT BREAKER OPERATION

I In figure 44(a) the normal circuit current passes through the main contacts and bi-metallic (thermal) element. Should a fault occur causing the current to exceed its normal value, the bi-metallic elements curvature will change (caused by the difference in the coefficients of expansion of the materials).

The latch mechanism is released as shown in figure 44(b), this then breaks the supply to the circuit (load) and the push/pull button on the front extends. When the bi-metallic element has cooled down, ie goes back to its normal shape, the button may be pushed in to re-engage the latch mechanism providing the fault which caused the tripping has been cleared.

The manufacturers of circuit breakers produce characteristic curves of' rated current against time at various ambient temperatures to allow you to work out tripping times and overload currents.

In three phase circuits, three pole circuit breakers are used, all three sets of contacts open and close together when operated manually and open an overload in any one of the three phases.

Reverse Current Relays

We have seen the use of the principle in dc generating systems; the differential cut-out trips the generator off line on a reverse current of 15-20 amps.

Reverse Current Circuit Breakers

A further category of circuit breaker designed for the control and protection of circuits in which the current flow is normally in one direction only, which trip automatically if the current in the controlled circuit should undergo a reversal of direction as shown in figure 45.-

J -- - / > -, , --

Circciit ,breakers of this specialiskd fo&;are normally incoiporated\,in certain :aircraft dc power supply systems, where their primary function is the isolation of any generator which, for, ariy reason, takes cQrrent from the main busbar for unduly pdotraste&p,eriods. 1 ,

\ - - -

This t i e of circuit breaker is not a f f e 2 e ~ by the moment&-rever%als of current in the generator output circuit, which takes place Fhen the generator, with falling output, cuts out under its normal automatic control system. Such a circuit breaker is normdlly closed and takesno active part in the general --- functioning of the power supply system, opening only if the normal generator controls failto cut out a negative output generator.

The circuit breaker embodies a spring loaded contact assembly, which is closed manually by a setting handle and is then held in this condition by a latch assembly. The controlled generator is then connected to the busbar through a series connection consisting of the main contacts and a single turn coil.

The main contacts open to disconnect the generator from the busbar when the latch assembly is released; this action is performed by the displacement of a trip lever, which in turn is operated either manually (by pressing the trip button), by remote control through the attracted armature electromagnet system, or automatically when a reverse current condition, of sufficient magnitude and duration, develops in the single turn coil. After tripping by any of these methods, the circuit breaker must be reset by operating the setting handle.

RUBBER SHROUDED MANUAL TRlP BUTTON SETTING LEVER

TERMINAL TO

TERMINAL TO GENERATOR

Fig. 45 REVERSE CURRENT CIRCUIT BREAKER

1 I BUS BAR

MANUAL TRlP PUSH BUlTON

Fig. 46 REVERSE CURRENT CIRCUIT BREAKER OPERATION

The single turn coil shown in figure 46 is located over a soft iron yoke provided with pole faces between which is pivoted a permanent magnet armature. With current passing through the single turn coil a magnetic field is established between the pole faces of the soft iron yoke. Polarity of this field (and hence the direction in which it tends to displace the pivoted permanent magnet armature) is dependent on the direction of current through the coil.

The polarity of the yoke field, with current passing through the single turn coil in the normal direction, is such that the armature takes up a position in which one of its arms is poised above an ext,ension of the trip lever.

While the other end lies immediately below, and just clear of, a setting lever. The same position is maintained by the armature when there is no current through the single turn coil.

With a reversal of current in the coil the polarity of the yoke field is also reversed, and a force tending to deflect the ends of the pivoted armature towards the trip lever extension and the setting lever is established.

The setting lever, which is contacted almost immediately by the armature, on the initial movement of the latter in the reverse current direction controls an escapement mechanism. This mechanism in turn controls the rate at which all further rotary movement of the armature (after the setting lever has been contacted) is made.

-- _ --

The-design is such that the rate of arrn$ure movement is-approxi$ately proport!ional to the force applikd to the setting lever by theipermanent m a 6 e t armature. I I

I

The force exerted on the settikg leverby the end of the amqature u,Ader reverse current conditions is determined\by the intensity of they6keifield and (hys varies with the magnitpde of revkrse current in thesing16' turn coil. I i \ I

i I Sustained relierse current, if slightly above the designed dinirnum value, results iKa smhl but constant -----a foiTeeb;ing applied to the-setting lever, and the escapement allows the armature to turn quite slowly until; after an appreciable delay, it displaces the trip lever. This unlocks the latching assembly and so permits the spring-loaded contact assembly to open.

Large reverse current, if sustained, will result in the circuit breaker tripping after a shorter delay, while an extremely severe reverse current will cause almost instantaneous tripping.

If the reverse current is not sustained and falls to less than the designed minimum trip value before the permanent magnet armature, restrained in its movement by the escapement, has been able to operate the trip lever, the armature is returned to its original position and no trip action takes place.

Remote tripping of the circuit breaker is effected by means of a tripping solenoid which, when energised, attracts a hinged soft iron armature to displace the trip lever and so release the mechanical latch assembly.

The supply to the tripping solenoid is controlled by a push-switch, arld the remote trip operating circuit is interrupted, when the circuit breaker is open, by the normally open contacts of the miniature switch shown in figure 46. The normally closed contacts of this switch, which open as the main contacts close, control a warning light or indicator.

The contact assembly also incorporates a set of auxiliary contacts which open and close with the main contacts; these auxiliary contacts complete the circuit to the field winding of the generator when the circuit brealter is closed. A mechanical indicator, visible through an inspection window in the breaker casing, provides a visual check on the breaker state, ie set or tripped. The manual trip push is formed at the end of a spring-loaded plunger which, when pressed, swings the permanent magnet armature on its pivot until it displaces the trip lever and trips the circuit breaker.

Circuit protection for dc generation systems, ac generation systems, battery systems and galley systems has been covered in the previous three

i books on Electrical Power. -

-

- 37 -

rnoodulll A-340

LBP Dec 03, Feb 04, April 04, April 07

Addendums module 11A book CABIN EQUIPMENT pending amendment action in response to student feedback after taking the CG examinations.

*** If there is partial failure of the passenger emergency lighting system the aircraft can be dispatched provided the passenger compliment is reduced to that number that can be carried in that part of the aircraft with a serviceable emergency lighting system. Check the MEL for the specific aircraft. The maximum number of emergency lights that can be out is 25% (www2. faa.gov/ certification/ aircraft).

*** A similar regulation applies to inoperable passenger exits. The passenger compliment is reduced and passengers are not seated near that exit. Again the MEL is consulted.

*** A hand held microphone is not allowed on public transport aircraft with a flight crew of more than one.

*** - - - -

Passenger-seats may face in any direction. -

*** Hand held fire extinguishers - numbers on aircraft. EASA25 states at least 1 on the flight deck and in the passenger (pax) compartment 1 f7 to 30 pax), 2 (31 to 60 pax) and 3 (61 or more p e ) minimum.

---

AN (Airworthiness Notice) 60 states that'in addition to extinguishers on the flight deck and in the galleys etc, the number related to numbers are (up to 60) 2, (61 to 200) 3, (201 to 300) 4, (301 to 400) 5, (401 to 500) 6, (501 to 600) 7, (60 1 or more) 8. A t least half but not less than 2 to be of the BCF type.

. . -

NOTE. AN60 now reassigned to cover design and installation of In-flight Entertainment Systems (IFE). IF YOU GET THIS Q IN THE CAA EXAM - MAKE A NOTE - TELL THE PERSON IN CHARGE AND WRITE A LETTER TO THE CAA. AN60 was withdrawn Nov 2003. IF YOU GET THIS Q IN THE CAA EXAM - MAKE A NOTE - TELL THE PERSON IN CHARGE AND WRITE A LETTER TO THE CAA.

*** Not more than 25% of floor path lighting may become inoperative in the event of the fuselage splitting in half in a crash landing. Was in AN56.

*****

NOTE: It is possible that some of the above statements may not be too meaningful when read out of context, so it is suggested that the appropriate book/subject be read first then the information above be checked against that topic.

CONTENTS

Page

Cabin safety equipment Life jackets First aid kits Smoke hoods Megaphones Public address system Hand held fire extinguishers Escape ropes Escape chuteslrafts Life raftlescape slide Life rafts Identification of emergency equipment Emergency lighting Other emergency equipment

Cabin general equipment Seats Flight deck seats cabin crew seats Passenger seats Berths Seat belts In Flight Entertainment

Aircraft equipment Cabin layout Cargo Air-stairs

Answers to self assessment questions

CABIN EMERGENCY EQUIPMENT

The cabin is the only area the passengers get to see. They tend not to bother if something technical goes wrong (though often they are not told), but if something fails in the cabin, like the In-Flight-Entertainment packs up, or the galley ovens fail then passengers get disgruntled very quickly.

Cabin appearance (decor) is also important to the passenger. If the cabin looks good it is assumed the rest of the aircraft has been well cared for - though that is not always true.

The design of the aircraft and its equipment (by regulation) has to take into account many emergency situations - such as:

-k Fire (and smoke) within the cabin (fire extinguishers - smoke hoods - firelsmoke detectors - fire proof materials for seating - floor path lighting etc) .

* - Emergency evacuation on-land-,(ropes/rope ladders - escape

- . -

slides). * Emergency evacuation on water (escape slides - life rafts - life jackets

- floatation cushions etc). * Survival after crash landing in desert or artic conditions. A Rapid deceleration1 flight- tlirough turbulent air (deat belts). * Cabin decompression (oxygen) - dealt with in the book entitled Cabin

-

Pressurisation. I * Medical emergencies (oxygen and first aid kits). * General emergencies (emergency lighting - crash axe - torches etc)

- .--

Note that all these have one thing in common - in a normal situation - - they are not used, but it is vital that they work when needed.

Regulatory Requirements

These change from time to time but you need to be aware of them and where they are stated. Currently the information regarding the carriage of equipment is laid down in the Air Navigation Order 2005 (ANO). This publication is dealt with in detail in module 10 Air Law. Briefly it states:

An aircraft shall not fly unless it is so equipped as to comply with the law of the country in which it was registered. The equipment carried must be clearly identified and be so positioned to be able to be used by the person for whom it is intended.

It lays down scales of equipment and specifies when the equipment is to be carried.

Emergency Equipment Required

The scale and quantity of equipment to be carried varies between the size and function of the aircraft and the terrain over which it is operating. Large commercial air transport aircraft have to carry the most extensive range. You need to bear in mind that not all aircraft need to have all this equipment fitted, so you should consult the A N 0 and Airworthiness Notices (ANs) and CS25 and 2 3 (downloadable free from www.easa.eu.int) for details.

In general the Certificate Specifications for aircraft (not including rotorcraft) are:

A CS-VLA Single engined piston (diesel or spark ignition) aircraft

with not more than 2 seats and not more than 750kg maximum certified take-off weight.

* CS23 Normal, utility, aerobatic and commuter aircraft. Maximum seating, excluding pilots, for normal, utility

-- and aerobatic aircraft is 9 and maximum weight is 5670kg. It states that commuter aircrafttare propeller

I driven <win engined aircraft with a m d m u m passenger seating of 19 and a weight limit of 8618kg. C323 includes specifications for piston engines, turbo jet and turbofan engines!

* CS25 Turbinei powered large aircraft. I r--- - I

In general CS25 states that the ;oliowing equipment is to be f;ltted to aircraft when flying for the purpose of public t!-ahsport:

I '

1 ;

1 Crew segts (flight crew anh cabin cr&w) must have multi-doint safety harnesses - unless allowed otherwise by the CAA.

2 Passenger seats (which may face in any direction) - lap straps (or lap straps with a shoulder strap or a full harness) on all seats, with a child restraint device for any child under the age of 2. A means of indicating to the passengers that belts are to be worn. Seats must be designed to take a 77ko person.

3 A life jacket for every person on board - when flying over water more than gliding distance from land.

4 Life rafts sufficient for all persons on board for aircraft capable of carrying 20 or more people - if flying over water more than a certain flying time from an airfield where a landing can be made.

5 Additional floatation equipment for 1 / 5th of the total occupancy - for manoeuvres over water.

6 A first aid kit - security sealed, to contain a minimum specified list of equipment sufficient in size for the number of passengers.

7 Public address system. 8 Smoke hoods for all the crew. 9 Portable battery powered megaphone.

10 Means of escape from the cabin (slides or ropes) - if the sill of any passenger door is more than 1.82m from the ground in any state of undercarriage collapse. Rope and attachment strength to be not less than 1779N.

11 Portable fire fighting equipment and gloves. 12 Torches. 13 Appropriate survival equipment if flying for long periods over uninhabited

tropical/polar regions. 14 Cabin escape routes to be illuminated and emergency exits signed. Supply

to be independent of normal electrical supply and, when armed, automatically activated if normal lights fail or aircraft is put through an impact of more that 2g.

15 Number, location and size of aircraft exit doors and escape hatches are specified. Must be openable from inside and outside the aircraft and clearly marked.

16 Any birth or litter must have a restraint system and withstand certain minimum inertia loads.

17 Baggage and cargo areas-must-be placarded to show maxlmum weight contents critical load distribution.

-

18 Maximum burn rates are specified for cabin materials. 19 If the commander cannot see all the passengers, a means of indicating to

the passengers that belts are to be warn (an illuminated sign). 20 All passengers must be able to escape within 90 seconds in an emergency

with aircraft on the ground. i # I

2 1 Minimum number of fire extinguishers available. I

Figure 1 shows, typically, the emergency equipment carried on an aircraft. Take few minutes to study the drawing %dnote the considerable . range --- of equipment - for this particular aircraft - with other aircraft having, possibly, different equipment. The important thing to remember is that all aircraft must meet the minimum emergency requirements as specified in the ANO, CS25 (large aircraft) CS23 (small aircraft) CS29 (large helicopters) and CS27 (small helicopters).

LIFE JACKETS

Called life preservers in CS25.

The A N 0 states that on public transport aircraft flying beyond gliding distance from land there must be one for each passenger and all the crew (demonstration jackets not counted). The jacket should be accessible from the normal seated position.

- 3 -

rnoodull l A-845

Fig. 1 CABIN EMERGENCY EQUIPMENT - EXAMPLE

Usually stored under the passenger seat, but sometimes stored in arm rests and within Velcro covered panels close to the seat. The location of the life: jacket must be clearly placarded. The jacket must have a whistle and a waterproof lamp. They are lifed from manufacture or overhaul (also the CO:! bottle), with a typical period being 6 years.

Equipment includes:

* A whistle (for children under three this need not be fitted). * A water activated cell that operates a light. * A COz bottle for inflation purposes - with a pull chord for

operation. x An inflation mouth piece. * Straps for tying around the body. * A spray hood (not always fitted).

All work carried out (by an approved organisation) on lifejackets must comply with the appropriate manufacturer's instructions (repairs, replacements, etc).

An inspection record with all the particulars of the lifejacket including details of mandatory modifications and inspections and the inspectors stamp and signature, should be kept at the maintenance base. The lifejacket should be serial numbered to identify it against it's record. - -

The jacket is made of bright coloured nylon material and buoyanyis obtained by inflating the jacket using a carbon dioxide (C02) cylinder attached to the jacket. The CO, is released by operating a mechanism a t the base of the cylinder.

- -

INFLATOR 1 v4q ' *V> - 04 ~5

PROTECTION { @<x COVEA .% 1-

"JERK TO'INFLATE" TAG

Fig. 2 TYPICAL LIFE JACKET

- 5 -

moodull l A-847

A standby mouth inflation valve is also provided should the C02 system be inoperative or the jacket requires topping-up. Deflation can be achieved by depressing the NRV in the mouth-piece.

To assist search and rescue the jacket is equipped with a light and water operated battery. Some jackets may also carry additional equipment, eg fluorescent sea marker dye; shark repellent; signalling devices; rain and sea-spray hood, etc.

Lifejackets are packed in nylon or plastic valises and instructions on their use must be displayed in the cabin and on the jacket.

Inspection

Inspections and tests are carried in accordance with the manufacturer's instructions and inspections may be ccuried out at 6 monthly, 12 monthly, or a t 18 monthly intervals. - - -

I -- r-

~nspec'tidns should be carried oyt f n cld& premises on smooth bork surfaces. The plastic material of the jacket and the webbing of the strapsshould not come into contact with oil, grease, or acid. I

I

For test purposes a supply of clean dry - low pressure air, and a water manometer I should be available. I

\

I

The lifejacket s h d l d be inspect& as follows:

- Check jacket against it's record card. -If-necessary, clean with a recommended solvent or warm water. The material should be inspected for slits, tears, holes, discolouration, adhesion of seams and general condition. Webbings and cordage's should be inspected for security of attachment, discoloration, cuts and tears. Metal and plastic components should be inspected for security of attachment, damage, and deterioration. Check any adhesive joints. The lamp should be checked for security of attachment, as should the wires and battery. If the battery is not water operated (rare these days) then it should be changed every 6 months. Check the water operated battery for moisture ingress, and a voltmeter test should read NIL volts. A n insulation test should give a minimum reading of 5 megohms. Check the C 0 2 cylinder for change by weight, and check also for signs of damage and corrosion. Check the mechanism for damage and corrosion. If the life of the cylinder (which maybe 10 years) has expired then it must be changed. Check that label on jacket or valise has recorded on it the date of next service.

rnoodull l A 3 4 8

Inflation Test

This may be done by pressurising the jacket (to a very low pressure) and checking the pressure drop over a period of time or immerse it in a clean tank of water and check for bubbles or use a n acid free soapy solution.

Repairs

Minor damage may be repaired if an approved repair scheme is available. The area of the leak may be marked using white chalk only - this to be removed after repair. Repeat inflation test.

FIRST AID KIT

A first aid kit must be carried and it needs to be security sealed. May contain specialist resuscitation equipment. The first aid kits can contained controlled drugs and care needs to be taken where they are located.

The kit must contain bandages and other equipment (as listed in the ANO), the quantity relating to the number of passenger seats in the aircraft.

SMOKE HOODS OR MASKS

Equipment for the protection of eyes, mouth and nose (CS25 and the ANO) for a minimum period of 15 minutes.

. .. .

-

The crew only (at present) are provided with smoke hoods. ~ h e s e , as with life jackets, need to be easily accessible from the normal crew station. The smoke hoods are contained in a box and are also lifed from manufacture or overhaul. The purpose of the smoke hood is to cover the eyes and nose and mouth of the crew to allow them to help passengers in an emergency.

Most crew emergency oxygen masks incorporate a face mask (smoke hood) to cover the nose and mouth.

MEGAPHONE

On aircraft where there are 19 to 99 passengers a portable battery powered megaphone is required to be carried. The scale increases with the size of aircraft, eg 100 to 199 - 2 megaphones, 200 plus - 3 megaphones. Figure 3 shows a typical megaphone.

rnoodull I A-849

The megaphonels need/s to be mounted securely in a clearly identified location. I t needs to be able to be removed easily by the crew. Periodic testing and renewal of the batteries is the only maintenance required.

- - -

, Fig. 3 PORTABLE MEGAPHONE I I

I

PUBLIC ADDRESS (PA) SYSTEM +ND INTERPHONE SYSTEI~ I 1 I ~

The PA system needs to be capdblk-of-br6adcasting to all areas bf the aircraft from locations such as the chief cabi$ officer's,ldcation (various g$ley positions) and the flight deck - this via the pildt's normal microphone (this will be part of his/her head fitted set as'a hand held Aicrophone is not allowed on transport aircraft with a flight crew of more than one).

I

The PA address must interrupt - all - video -- and audio channel[~ (within the aircraft zones selected) on the in-flight entertainment system and be capable of broadcasting to all speakers located around the aircraft including toilets and galleys and the speakers in the in-flight entertainment system (head-phones, armrest speakers etc). Aircraft speakers are often located in the roof space. The PA system must be serviceable for all flights with passengers.

Applies to aircraft with more than 19 passengers.

HAND HELD FIRE EXTINGUISHERS

Both water and BCF extinguishers will be found in the aircraft cabin. These are for crew use only. They will be clearly identified with water in a red container and BCF is usually coloured green. (Note. BCF is currently causing concern because its adverse affect on the ozone layer and current EU regulations for ground based extinguishers is to have them all coloured red with only a small square of the appropriate colour showing).

rnoodull l A-850

Operation of the bottle is required to be by a single handed operation. It is removed from the bracket, the safety catch is pushed up with the thumb, it is aimed at the fire and the trigger squeezed.

The bottle itself has a 'full indicator' disc, which is pushed out as the trigger is depressed, giving indication of a partially (or fully) discharged bottle.

The cabin crew will make an appropriate entry in the log to inform the engineers if the bottle has been operated. It must then be changed. In any emergency situation where bottles are likely to be used they should all be checked. Normally checked by weighing, but some bottles have a pressure gauge with green and red indicator segments. Some have a 'test soft spot' usually on the bottom - if this can be pressed in with the finger the bottle needs changing.

Fig. 4 TYPICAL HAND HELD FIRE EXTINGUISHER

With reference to figure 4. The transit pin should be fitted whenever the bottle is being moved for maintenance purposes. It is important that it is removed after fitment to the aircraft. When fitting, also check that the extinguisher fits into the bracket in such a way that the bracket trigger guard or lug fits so as to prevent inadvertent firing of the bottle whilst in the bracket.

- 9 -

rnoodull l A-851

Bottles should be weighed periodically as laid down in the approved maintenance schedule. It is also checked for damage, that the transit pin is removed and that the full disc or fired indicator is still in place. The life date should be checked.

Written on the side of the extinguisher will be the full weight a t manufacture /overhaul. Also stated will be the type of extinguishant, manufacturer, details of how to operate and life date.

The bracket will require inspection whenever the bottle is removed to ensure it is not damaged or corroded and that the quick release mechanism functions correctly. As with all emergency equipment fire extinguishers need to be clearly labelled and are located near to normal crew stations.

Next to the extinguisher you will often find a pair of fire gloves.

The numbers of hand held fire extinguishers required as stated in EASA CS25 are:

--

-

* At least one in the flight deck. , , I * At least one in each class A or B cargo compartment. Also class E if

accessible to the crey. k At least one in eacq gklley. I * '

[ I

A minimum number c!f-extinguishers in the pass'engg compartment must be of the BCF type (Hidon 121 1 or similar).

I I 1

Numbers required depends on bassenger n'umbers: 1 , Passengers - Extinguishers - ,

7 to 3*-- '1 31 to 6 0 2 61 to 200 3 201 to 300 4 301 to 400 5 401 to 500 6 501 to 600 7 601 to 700 8

For more information on hand fire extinguishers see the book in the -LBP series module 7 entitled Safety.

ESCAPEROPES

If the door sill of an aircraft is higher than 1.8m (6 ft) from the ground in any configuration of undercarriage collapse, an assisted means of escape has to be provided. For flight deck crew this normally consists of a retarded rope system, or rope ladder.

ESCAPE ROPES

ESCAPE ROPE -

Fig. 5 TYPICAL ESCAPE ROPE STOWAGE - -

The rope assembly is securely fzed to an anchor point normally-above the escape window. In an emergency the free end of the ~ope/ladder is removed from its stowage and deployed to hang out of the dv (direct vision) window, the crew climb out of the dv window and climb down the ladder or are lowered slowly by the rope retard mechanism in the rope deployment unit. -

-. .- -. - - --- -- ..

Must have a minimum strength of 1779N (4001bf/182kg).

Minimum number and location of emergency exits are specified for the flight deck to include, in some cases, exists above head height.

Some aircraft have inertia escape ropes which lower the person to the ground. The pilot holds onto the rope handle and starts to descend. A s the rope is deployed it turns paddles within a small chamber filled with hydraulic fluid. This has the effect of slowing the descent rate to an acceptable level.

Periodically these units must be removed and the ropes and inertia units checked for damage, security and operation. Weights are used to check the rope deployment rate).

For passenger emergency evacuation escape slides/chutes are used.

moodull l A-853

Fig. 6 SLIDE/RAFTS DEPLOYED AFTER DITCHING

Fig. 7 SLIDES DEPLOYED ON LAND

moodull 1 A-854

ESCAPE CHUTES/ SLIDES

For passengers the normal method of emergency escape is the self-inflating slide. Often called a slidelraft it inflates automatically when the doors are opened in the armed emergency mode. On many aircraft the slide doubles as a life raft in that once inflated on a ditched aircraft it will float on the water and can be detached from the aircraft and used as a raft.

GAS

TRIGGER MECHANISM -!&

TLIDE ASSY.

Fig. 8 SLIDE STOWAGE - 1

Normally - - -. - housed - within the decor panel of each door (though some are housed in a container next to escape h a t c h e s ) : - ~ i ~ r e s 6 and 7 shows the rafts deployed on water and land. Typical stowage examples are shown in figures 8 and 9. When the door is opened the slide is caused to operate (door operation is covered in the book entitled Structures 1 in this series).

The unit consists of a gas bottle with some aircraft having an air charged bottle at 3000psi (2 1MPa) and others having a nitrogen charged bottle a t about 3,500psi (24MPa). The bottles are fitted with pressure gauges that can be seen when the slide is packed.

On operation the compressed gas passes into the slide tubes via an air aspirator. This draws in extra air from the atmosphere (slide tubes have a capacity far too great to be inflated by any reasonably sized gas bottle alone). A pressure relief valve is fitted to prevent the slide over-pressurising and a fusible plug is fitted to the bottle to prevent the bottle bursting due to excessive heat.

Slide illumination is provided automatically when deployed and the slide itself is neatly and carefully folded into the packboard and laced up.

ESCAPE SLIDE

I

Fig. 9 SLIDE STOWAGE - 2 I

Operation I

On older aircraft a member of the ,crew would' lift the girt bar o u t of i ts stowage position on th,e door and lock it into the floor brackets (figure1 11) before the door is opened. When the door is opened (outwards) the slide pack would move with the door and the slide would be-pulled out of its pack and its falling weight would trigger the inflation bottle.

With later aircraft girt bar engagement is automatic when the slide arm lever is moved to the arm position after the door is shut and locked.

With the door armed (arm/dis-arm lever on the door moved to the armed position) and the door release lever moved to open, the door lock bolts/catches disengage and, with gas assisted doors, the door opening jack pushes the door open. The girt bar (girt bar and top end of slide) is attached to the floor and the slide itself i s attached to the door.

The action of the door opening pulls on the slide pack and releases the lacing/Velcro which allows the slide to fall from the packboard and hang outside from the aircraft floor. The action of the slide dropping out of the door pulls on the bottle actuation cable and operates the firing disc inside the bottle. The bottle discharges through the aspirator. This inflates the slide by sucking large quantities of air from the atmosphere into the side tubes.

HARNESS ATTACHMENT TO DOOR

LOWER SHELF HINGE

HARNESS

PRESSURE GAGE

LOWER SHELF

NO. 1,2, OR 4 PASSENGER DOOR (LINING NOT SHOWN) - --- -

PACKBOARD

ESCAPE SLIDE

Fig. 10 SLIDE STOWAGE - B757

moodull l A-857

- - - - -- ~-

,~-

LANYARD 1 I , \

, L--, , \ ,

I I , , . ,

1 1 I

Fig.\ 11 GIRT BAR 0PERA;ION - OLDER AI~cRAFT

I

I

Figure 12 shows the deployment olf-the-slide based on the Boeing 757. The slide is a two tube slide, for added rigidity longer slides are four tube slides - two on each side one above the other. Slides have emergency lights positioned along the top of each tube which come on when deployed.

blank

NO. 3 EMERGENC'I =IT DOOR

ESCAPE SLIDE

W O R CUTOUT

NO. 3 EMERGENCY

1. DOOR OPENS 2. ESCAPE SLIDE STARTS TO INFLATE

3. ESCAPE SLIDE FULLY INFLATED

Fig. 12 ESCAPE SLIDE DEPLOYMENT

Most aircraft have a system of levers and push/pull rods to engageldisengage the girt bar. The arm/disarrn lever on the door moved to arm will cause either the girt bar to 'extend' out either side of the door to engage with locking fittings on the aircraft floor at each end, or to move a locking latch from the door to the floor fittings.

R E L I E F

- -ASPIRATOR

I N F L A T E D S L I D E I R A F T CROSS SECTION ( S C H E M A T I C ) P R E S S U R E GAGE

S A F E T Y

I

, , 1 I I

I

I Fig. 13 ?NF&ION SYSTEM -

I I \ I I

I , I

, -

- ---- - -

CONDITION : Armed.

show disarmed

Fig. 14 GIRT BAR ENGAGEMENT MECHANISM

Figure 13 shows the general arrangement of the air aspirator inflation system. Take a moment to study it. Note how the inflation cylinder (bottle) is connected to the aspirators and note the details of the bottle itself.

Figure 14 shows a girt bar locking system using sliders that effectively extend the bar's length. Shown in the armed position with both sliders out and engaged with the floor mounted lock fittings. When the door opens (in the armed state) the operating mechanism moves with the door - complete with the slide pack, leaving behind the girt bar which has attached to it the girt (the top part of the slide).

The arrows indicate the movement during disarm when the sliders are pulled out of their floor mounted lock fittings.

Figure 15 shows how the door operating mechanism moves away from the girt bar when the girt bar is engaged and the door is opened.

, Most modern aircraft use slides as life rafts, with inflatable canopies provided in the emergency kit. If this is not the case separate life rafts are provided. These are normally located by the doors, often in the roof space immediately above each door.

PERATING MECHANISM

FLOOR MOUNTED LOCK FITTING ith slider attached to floor)

w Fig. 15 GIRT BAR ENGAGEMENT DETAIL

Study figures 16 and 17 re the flight deck indications. Note the use of magnetically operated proximity switches using a 28V dc supply and using the switches on the earth side of the circuit. Note also that this signal will be sent to a Symbol Generator Unit (SGU) to produce a colour display on the flight deck CRT/flat screen display.

blank

ENTRY AND 1 ' ATTENDmTs LIGHTS -1

Fig. 16 GIRT BAR ENGAGEMENT INDICATION CIRCUIT

RIGHT BUS

GIRT BAR

\

Fig. 17 A 3 3 0 DOOR INDICATION SYSTEM

n r p I1L'I PASS. DR I-', PASS. OR

PROX SNSR A PROX SNSR B PASS. OR, GlRT BAR ANN. LIGHT

Maintenance

TO OTHER PASS. DR GlRT BAR ANN. LIGHTS

In-situ checks will include checking for correct operation of the armldisarm mechanism; correct operation of the door opening mechanism (ONLY in the dis-armed condition); correct clearance between chute container and fuselage when opening the door; correct bottle pressure; bottle transportation pin removed; any flight deck indications; damage and security of floor mounted girt bar locking brackets; security of pack to door; security of any lanyards and correct lacing/Velcro attachment on chute packing.

Maintenance tasks will include removal for inspection and periodic lifed item changes. Whenever this is the case the slide is deployed to prove serviceability. Care needs to be taken during these type of activities to prevent injury or damage as the door will open fast and the slide deploy violently.

The slidelraft is subject to vibration, creasing, and general deterioration and the nylon fabric should be inspected periodically for:

* Cuts, tears and damage. * Chaffing and hardening. * Legibility of markings. * Security of attachments, ropes, etc. * Security of seams and leaks. * Lights for correct operation.

Check all metal fittings for correct operation, corrosion, damage, cracks, and , security of attachment. Check all indicating systems (study figures 16 and 17 for

examples of the types of indicator systems fitted). - -

Check all internal stores (drinking water bottles, flares etc) (raft). Check water activated battery for security of attachment and signs of moisture ingress (raft).

At the periods stated in the maintenance schedule the slide should be removed for overhaul. This will involve a functiond test and pressure test. A full functional

I

test (automatic deployment) should be carried out at least every; 36 months and the test must be videoed (AN12) '(AN = Airworthiness Notice). ,

Sample testing of escape chutes may be carried out to an agreed program provided --- - all . slides are tested withinthe above period. All failures must be subject to an MOR to the CAA.

Each type of slide must have its own servicing manual and maintenance bays should have the correct facilities including smooth topped work benches. Each slide should have a record.

The bottle is removed by disconnecting it from its operating mechanism carefilly as inadvertent operation could cause serious injury and very little force is required to set it off. After the bottle has been disconnected and the slide disarmed the safety pin should be fitted to the bottle and the bottle stored correctly in its container.

The Fabric

Originally neoprene proofed nylon coloured yellow. To comply with FAA TSO-C69A chutes are now made from polyurethane proofed nylon coloured silver on the outside.

Used Escape Slides

These should be returned to the manufacturer with a report on why they were deployed. Training escape slides (for cabin staff and air crew) should be marked accordingly and not used for anything else.

Storage

Slides should be stored deflated in their original containers with transit pins fitted, and lightly dusted with french chalk. Do not store more than 3 on top of each other. Temperature should be between 15 to 2 1°C; dry atmosphere; free from dust and fumes and out of direct sunlight.

The stored slide should be inspected every 18 months and inspected prior to fitment to the aircraft.

-

Repairis

Compgnents may be replaced and the fabric may be repaired by an approved organigation in accordance with the manual. Most repairs are carried out by patching - using adhesives, and're-cementing -- - the seams. I I

I \

1 -

Some parts of the, slide are not repairable. Refer to the manufacturer's literature. Repair: kits are available.

I I

I

All markings must be replaced using copper free ink. - - - - -

Escape Slide Requirements

When the aircraft is on the ground with any sill more than 1.8m from the groun" in any state of landing gear collapse then an assisted means of escape must be provided to allow all passengers to evacuate within 90 seconds. For passengers this is an inflated slide which must:

* Inflate in 10 seconds. * Inflate in a 25kt wind and be held stable by no more than one person. * Be a double slide at type A exits.

Where overwing emergency exists are provided then inflatable wing walkways must be provided and inflate within the same time period.

Where the passenger first lands on the ground that area must be illuminated to 3 lux (3 lumens per square metre) or more. The lights are normally fitted to the slide.

rnoodull l A-864

SAQ I

Prior to refitting a slide pack, what inspections should you carry out?

LIFE RAFT/ESCAPE SLIDE

Requirements

The CS25 structural specifications of the aircraft include the requirement that, provided the aircraft lands reasonably intact in the water, it should stay afloat long enough to ensure sufficient time to evacuate everybody to life rafts.

The A N 0 states they should be fitted to public transport when-ever the aircraft flies over water more than a certain gliding distance from an airfield. It also states that equipment is to include:

- -

x Weather protection. * A sea anchor. Prevents the raft being blown too quickly over the water

with a wind blowing. * Attached life lines and tow lines. * Paddles, water proof torch, pyrotechnic distress signal flares, a

quantity of fresh water (or provision to convert sea water to fresh water), some glucose tablets and radio survival beacons - (numbers depending on the raft numbers).

The quantity (specified in CS25) must be sufficient for the seating capacity of the aircraft with sufficient spare to cope with the lose of the largekt raft. They should be stored near the exits and if s to r~ou t_s ide and released remotely must be attached by a line to the aircraft. The raft line to be easily detached.

CS25 further states that, if the aircraft is not certified for ditching, then there is to be sufficient approved floatation gear (approved seat cushions for example) for the seating capacity.

If the escape chute is deployed with the aircraft in the water then it will convert automatically to a life raft. It can be detached from the aircraft and, to comply with the ANO, it must be fitted with additional items of equipment.

If the normal exit cannot be used the chutelraft must be capable of being moved to another exit, attached to the aircraft then lowered into the water and inflated. When evacuation into the raft is complete the mooring line to the aircraft is releasedlcut (a tool is provided on the raft).

In normal emergency operation the slidelraft (as a raft) is armed, the door is opened and the slidelraft deployed. Once deployed two handles are pulled, marked RAFT DETACH - PULL. This action extracts two locking pins which detaches the chute from the aircraft - except for the mooring line.

LIFE RAFTS

Figure 18 shows a typical life raft. The raft should be packed in a stowage such that it can be inflated easily by the crew or the passengers, and is accessible from an exit.

The life raft pack is inflated by pulling a lanyard which operates the valve to allow the COz bottle to discharge into the raft buoyancy chambers. The life raft, which is attached by a line to the aircraft, will inflate together with the canopy. Lines and rope ladders will allow passengers to get inside the raft from the water.

7-- The floor is double lined for warmth -- arid - axrain water catchment system is . provided.

I , I \

i By regulation life rafts should: I

1 % I /

x Be stored in containers-/c6mp&tments free from sharp edges and extremes of temperatux- -\ ' , ,

\

* Be ?ttached to the aircraft by line that prevents i the inflated raft frorntbeing blown a&ay by the br,keze but at the $ m e time is so

I - - designed that when the-drcraft sinks the line should not cause the _ raft to capsize (sharp knife or-detachment system).

* Be capable of operation at its stowage location and, if fitted outside, from a remote location or automatically.

Maintenance

In general this is not too unlike that required for slides.

Additionally check all internal stores such as drinking water bottles, flares etc. Check water activated battery for security of attachment and signs of moisture ingress. Check voltage output - it should be zero.

RAINWATER CATCHMENT INTERIOR LIGHT TUBE(S)

/ DOOR CLOSURE HANDLEIS) /

RAINWATER CATCHMENTE)

1 &QUOIT \ \ .lsE INFLATON CYLINDER LIFELINE (inside)

PAINTER PATCH - -

\ CANOPY DOOR(S)

Fig. 18 TYPICAL 10 MAN LIFE RAFT

Identification of Emergency Equipment (Placarding)

All emergency equipment must be clearly identified and located correctly. The normal method of doing this is through placards showing clear directions to the equipment and directions for use if required. These placards need to be of a specific material and design.

Any defects of cabin emergency equipment are serious and are covered under the Mandatory Occurrence Reporting (MOR) scheme.

A mandatory requirement is the provision of passenger emergency information in the appropriate language/languages. It is also a requirement that doors, life rafts and escape chutes should be easy enough to be operated by passengers.

Study figures 19 and 20 and note the operation of fitting on oxygen masks (more about oxygen systems in the book entitled "Cabin Air Conditioning and Pressurisation"), life jackets, opening of doors and the general deployment of slide rafts and rafts.

Fig. 19 PASSENGER SAFETY PLACARD - 1

Note. A n aircraft can fly with a door slide inoperative provided it is allowed in thr MEL; that the exit it is clearly marked NO EXIT; that the pilot is informed and th, passenger capacity reduced.

blank

S~artund L.Munp ~~blrraum.r( Nol.usgang 4 ~~y,;:;bdlno NO m k l n o E-,p.ncl e l l ,

~ . p l r ium, EXIT ~ o n ~ a ~ s e c o ~ , ~ NOIYIIIII S*lld.d. ulgencia H1olurn.r 5.161 a ~ r r n e ~ p & n ~ . ~

h ~ c o l l g e r n . NWI tumara UICIII a, ~m(cOlnr.

atartisu~psm ../mksa m x a t t v n Decollo. an.n.ggio H . r r ~ ~ b A ~ p i m u i o.um m.m/#mnaa ~ i o wlnt w~<cm.zapav~-t Ow.--- Zabr.nlano Dutilt 1 z I u I l n u l a ~ sun I l,ao*mi. A"-rGnot 'Uobo<c~rburou SI.r l lsp~t l lnje t0-v T . h l i k ~ c ~ h ~ ~ kap,,, P,c&vqm SIQM ~gilmal .PflU CI++ - +='ti. Kalkq YC ini* +A'J t.Yn

1 EXIT 1,2,4

EXIT 3

Fig. 20 PASSENGER SAFETY PLACARD - 2

Emergency Lighting

The A N 0 scale Z requires that sufficient lighting be provided both inside and outside the cabin to facilitate the safe evacuation of the aircraft in an emergency. This to be operable in the event of normal electrical supply failure.

rnoodull I A-869

It (and CS25) requires emergency floor path lighting in the passenger cabin sufficient to allow evacuation of the aircraft in the dark or in smoke conditions. These lights may be fixed to the bottom part of the seat structure next to the aisle or fixed to the floor. On some aircraft they are also fitted around escape hatches.

CS25 emergency lighting requirements include:

The emergency electrical system must be independent of the main aircraft electrical system except that the batteries may be charged by the main system. The emergency batteries must be able to supply the emergency system for a t least 10 minutes. If the fuselage was to split in two in a crash landing then no more than 25% of the emergency lights should fail. The floor path and emergency exit floor area must be illuminated and illuminated (back light) signs placed at specified locations within the cabin. Sign letter sizes are specified as is their colour (red) and back ground colour (white). ErrGergency signs are to iniiCate locations of

- --

emergency exits. \ \

Floor path lighting b u n t be so if all illurnihation Fft above floor level is obscured' the escape route can still be kollowed by the passengers : Switching by the pilot or cabin staff with switch gefectionb ON, OFF and ARM. When selecf~d-to '~~h$ the emergency lights - - will come on automatically if the 1 normal electrical system fails. ~ x t $ A a l lighting to idclude o ~ k r - ~ i n ~ exits, escabe! slides and ground areas around the bbttom of slides. I I I

[ I ! ,

, C - '

The specifications are comprehensive to include minimum lighting values and sizes for emergency signs and minimum lighting values and coverage areas for emergency lights.

ELECTROCUHINESCENT STRIP (BONDED TO FLOOR PANEL)

FLOMI PANEL

AFT ELECTROLUMINESCENT CROSS AISLE EMERGENCY LIGHT

EMERGENCY ESCAPE P A M LIGHTS

Fig. 2 1 TYPICAL FLOOR MOUNTED EMERGENCY ESCAPE LIGHTS

rnoodull l A-870

Figure 21 shows typical floor mounted emergency lights, powered from the essential dc bus with a dedicated emergency battery back-up.

For more information on lights you should refer to the book in this series entitled Lights.

Other Emergency Equipment

Other emergency equipment on aircraft includes:

* Torches. One for the use of each flight crew and one for each cabin crew member affied near each floor level exit (normal of emergency).

* A locator beacon. Capable of transmitting on 12 1.5MHz and 406MHz.

* - Oxygen systems. The requirements in the A N 0 are comprehensive and detailed. The individual requirements depend on the flight level of the aircraft, the ability of the aircraft to descend quickly and the persons for whom the oxygen is intended.

For example, if the aircraft is flying at 30,000ft and is capable of descending to 15,000ft in 4 minutes then the oxygen supply should be sufficient for 30 minutes supply for 15% of the passengers. If it cannot descend a t this rate then the supply should be enhanced to include a 10 minute supply for all the passengers.

That concludes this section on emergency equipment. Have a go at the following exercises to test your understanding. The answers are to be found in the text.

Exercises:

1. What is the minimum door sill height above which emergency slides are required?

2. Must an aircraft have a separate life raft? 3. Do all life jackets have whistles? If not why not? 4. What are the requirements of a megaphone? 5. Where are fire gloves kept? 6. How are hand held fire extinguishers prevented from inadvertent operation? 7. Look at the emergency equipment on your aircraft. Make sure you

understand how it works - using the manual if necessary. 8. Check the scales of equipment in the Air Navigation Order (if you have

access) for equipment to be carried and see how your aircraft compares.

CABIN GENERAL EQUIPMENT

SEATS

There are essentially two types of seat:

* Crew seats * Passenger seats

Crew seats can be further sub-divided into flight deck seats and attendant seats and passenger seats can be sub-divided into main cabin, business class and first class seats.

Flight Deck Seats

Figures -- 22 & 23 show typical flight deck seats. Equipment includes: -- - -

' . -

* A five-point harness (&-figure 22 the upper two points are not actually show, but will1 be their). I t has an inertia reel style of belt and has the option to lock the inertia reel in position., (Inertia reel belts use a small lead mass to move a locking pawl intp gear. This may be operated by the forward--movement of the pilot andlor the

' deceleration of the $reaft). ' --

I ' i i * ~ a t e r d and verticalsqat adjustment to ensure the crew member is

conifortable at the qontrols. , I I

-

* - Adjustable lumber-support and thigh pads.

* Arm rests that can be stowed out of the way for access and emergencies.

* Seat cushions and covers that meet fire safety tests and regulations defined in Airworthiness Notices, and do not impede flying control movement.

* Electrically powered adjustments (figure 23). With most older types of aircraft seat adjustment is manual.

Maintenance

Belts. Check belts for security, wear, damage and correct operation of quick release mechanism. Sit in the seat, strap yourself in and lean forward - the release mechanism should release without too much force.

On inertia reel type belts check the inertia reel by giving the belt a quick tug and it should lock. Check the operation of any manual locking and any ratchet mechanism. Check belt webbing for fraying - include all the webbing down to the inertia reel unit.

BACK CUSHION

OVERRIDE (HORIZONTAL)

Fig. 22 FLIGHT DECK CREW SEAT

Initial reel assembly. Check for security, damage and correct operation. If the seat does not have an inertial reel assembly check the quick release mechanism and ensure that it allows the belt to ratchet back into place. The shoulder straps are fitted to the initial reel assembly sometimes by a cable - check it is not frayed and runs correctly over the seat top pulley or roller.

Structure. May be made of aluminium alloy tubing with riveted, bolted or welded joints or machined sections bolted together. Check for damage, corrosion and security. Check seat security to the floor. If the seat slides in floor runners check these for damage, corrosion, wear and cleanliness. Check all seat adjustments for full and free range of adjustment and check that in each position it locks securely in place.

blank

m, HEADREST

SHOULDER HARNESS

BUCKLE

LLRIBAR REST

ABOOHINAL BELT

S I D E S T I C K AiVII1EST

SLEEVE

COLUIW

Fig. 23 ~ 3 3 0 FLIGHT DECK SEAT I ~ 1 - /'

1- HEADREST CUSHION

SEAT CUSHIO

BACKREST CUSHION

SEAT STRUCTURE

STOWAGE COMPARTMENT DOOR

SEAT TRACK F I T T I N G

Fig. 24 FLOOR MOUNTED ATTENDANT SEAT

Seat electrics. If the seat is electrically adjustable check that all adjustments work correctly. Check motors and cables for security, damage, signs of overheating and contamination.

Cushions. Check for damage, cleanliness, security and check that they are dry. Check with the seat in the forward most position, that the cushions do not foul the controls.

Cabin Crew Seats

These are usually located near the main aircraft exits (by regulation the seated attendant must be able to see the passengers for which he/she is responsible and must face forwards or backwards) and will fold up out of the way automatically as soon as the person gets up. They will be either forward or aft facing and depending on this depends on how strong they are, the forward facing ones being stronger.

Fig. 25 BULKHEAD MOUNTED ATTENDANT SEAT . -

- - -

When fitted to a bulkhead the structure must be capable of withstanding the normal static loads plus the acceleration (positive and negative) forces. Floor mounted seats are located and locked into the seat track in a similar way to passenger seats (note the seat track fitting in the drawing).

In figure 24 the area under the seat is used for stowage of emergency equipment. There may be a 'phone attached to allow the attendant to communicate with the flight deck or another flight attendant station.

Inspection of these seats follows a similar pattern to other seats. One additional test is carried out to check that the seat pan folds correctly and unaided (normally sprung loaded). If the seat pan was to remain down it would cause an obstruction to an emergency exit.

Passenger Seats

For larger aircraft these generally have more equipment fitted to them and can vary considerable between operators and between the class of the seat - main cabin (economy class), business class and first class. Seats may face in any direction and can fold forward except those at exit aisles.

Figure 26 shows a seat assembly that you might expect to find fitted in first class cabins/ business class cabins.

ADJUSTABLE L I D I N G HEADREST HEADREST WINGS

BACKREST WITH LITERATURE PACKET

ADJUSTABLE LUMBAR SUPPORT

F IBER-OPTIC TELEPHONE, PCU READING L I G H T S AND GAMES HANDSET

--

ARMREST WITH JACK PLUG AND

ARMSTOW TABLE

ASHTRAY----, V I D E O PLAYER

RUBSTRIP AND L ITERATURE POCKET

T I L T I N G SEAT

P I V O T OUT COCKTAIL TRAYS

I

I

a VIDEO MONITORS

' LIFEJACKET STOWAGE

SELF STOWING FOOTREST

Fig. 26 BUSINESS CLASS OR FIRST CLASS SEAT - EXAMPLE

INFL IGHT ' m A Y (folded under A

FOLDING L E G SUPPORT

\ FOLDED F O O T

SUPPORT \

under drinks table) /

BARS

Fig. 27 SEAT RECLINE - EXAMPLE

rnoodull l A-876

The seat is clamped to the floor seat tracks (figures 31 and 32). Each track is a hollow extrusion with a series of holes in it a t 1" (25mm) pitch. The seat locating lugs are placed in through the holes in the track then moved %" (12.5mm) forward to allow the feet of the lug to lock into place. A locking tab is provided which prevents the seat from moving forwards or backwards once locked.

Fig. 28 UNFOLDING A FLAT SCREEN FROM AN ARM-REST

Fig. 29 PULL-OUT CONTROL HANDSET

FOLDING BACKREST

FDLOING TABLE

Fig. 30 TYPICAL ECONOMY CLASS SEAT

L6EVLUOwwme=

PASSENGER SEAT ~ ~ L E ~

m mrm

( R - - -

m SEATLOCKAELEGE LEVER

-\ I '

\

/ YWRUaW I nw

! -- -' /

Fig. 3 1 TYPICAL SEAT -.SEAT TRACK CONAECTION

I

~ ~ I

I -

-- - -

UNIT (TYPICAL)

SEAT ELECTRONICS UNIT (REF)

ZONE 1 SEAT GROUP (TYPICAL)

SEAT TRACK COVER /

INTERSEAT CABLES I N RACEWAY

Fig. 32 SEAT TRACK CABLING

Fitted equipment can vary considerable, but some equipment is mandatory. Listed below are most of the items that can be found fitted to passenger seats, with some only on first class and/or business class seats.

* Seat belts - It is a requirement that seat belts be provided for all seats. This has to be a minimum of a lap strap but you could also see a lap and shoulder strap (similar to those fitted to most cars) or a full harness on some seats. Some seat belts, fitted to seats immediately behind bulkheads, may incorporate an inflatable air bag - operated automatically if the aircraft ,decelerates too quickly as in a crash landing.

* Cushions covered with fire resistant material. CAP747 gives details (was published in ANs) .

* Centre head rest - may be adjustable. * Side head rests - again may be adjustable. * In Flight Entertainment Equipment (IFE) and reading lights. * Games control unit - connected to the seat by a cable. * Telephone - may be part of the games control unit - credit card operated. * A method of gaining the crew'sattention without leaving theseat. --

* Folding trays - in the back of the seat or in the arm-reqt. * Individual reading lights in the side head rests. I

* Adjustable reclining backs - in general main cabin seats will only recline a little way whilst business class will recline almost to a 1;ecumbent position and first class will go all the way down to make a bed. Arm rests may fold up out of the way on main cabin seats. --

* Adjustable leg supports/foot supports. I

* Life jacket. * Document pocket - containing the mandatory safety noticels. * Chemical oxygen generator - fitted in the backs of seats for the passenger

behind (rare). - - ,

* Flat screen displays for IFE positioned in the back of the seat for the passenger behind or on adjustable arms folded into the arm-rest or spaces between the seats. Folded screens come on automatically when unfolded.

* Floor path lighting. Some seats have emergency lights fitted low to the floor illuminating emergency escape routes.

* Ashtray. Mandatory that it has a lid. * Electrically powered adjustments for tilt, recline, leg support etc. M a n seats

are manually adjusted. * Folding drinks tray. * Floor level luggage rail - fitted near to floor level to prevent any luggage

under the seat moving forward in severe braking conditions. * Power supply for a laptop computer.

Berths

Fitted to long distance aircraft for the use of a "slip crew" or fitted for medical reasons. Of simple construction meeting current EASA fire and safety regulations.

The fire regulations are the same as seats and class B and E cargo compartments (refer to the book in this series entitled "Structures - 1"). A test flame is applied for a period of time and when removed the flame on the test material should go out after 15 seconds. For seats etc three test samples should be used.

The berths may be contained in a small 'bedroom' on the aircraft. On large there may be one bedroom for flight crew, often behind the flight deck , and another for cabin staff, sometimes a t the rear of larger aircraft above the cabin ceiling up a short set of steps.

SEAT BELTS

Seat belts (and seats and berths) have to comply with the strength requirements ( A N 0 and CS25/23) of holding a person of 1701bs (77. lkg) safely in various directions of aircraft movement with a safety factor of 1.33. For example, the seat and belt should be capable of holding a 1701b person at a 9g forward acceleratir of the body (as would happen in a crash landing) with a 1.313 safety-factor. F r o n ~ row seats 16g as are all new aircr8fT.-

-

I \

A N 0 ( ~ c $ e B) states that a safety belt/hAness is required for all crew' and passenger seats and special chiid restraiqits for each child under 2 years of age. EASA !25/23 deals more in terms of the design features of the belt, and includes

---

such things as: ~ ' \ 1 -- I 1 I

A The belt strength. 1 I 1 I

* light deck seat belts must not be in such a as to foul the ' cdntrols - or they musl have,special stowage provision.

* . .- They must have metal to metal'latching devices. * They are tested using dummies of 170 1bs (77.1 kg).

Crew Seat Harnesses (Figure 33)

These vary in design but must conform to CAA/EASA regulations to restrain the pilot in all foreseeable emergency situations. The drawing shows a five point harness with the fifth strap between the pilot's legs. The two shoulder harnesses pass up over rollers at the top of the seat back to inertial reel assemblies either near the head restraint or at the bottom of the seat back.

On some seat assemblies the two straps are stitched to a single strap which passes around the inertial reel, on others each strap passes over the reel assembly separately.

The harness consists of a shoulder harness assembly, right and left hand abdominal belts, centre strap, release buckle, a harness reel and an operator unit assembly.

- 38 -

rnoodull l A-880

The straps are made of nylon webbing. The reel strap or shoulder harness is attached at one end to the harness reel, and at the other end to a metal fitting. Plug in type buckle fittings are stitched to the ends of all the straps except one. This has a permanently secured release buckle attached.

Abdominal and centre straps are fitted with body pads and adjustable buckles.

SHOULDER HARNESS CAN BE REMOVED UITHOUTREHOVING ABDOMINAL BELTS

--

ABDOMINAL --

BELT

SHOULDER HARNESS CONTROL

LEVER

Fig. 33 CREW SEAT HARNESS

The release buckle is a cylindrical housing in which the plug-in fittings of the other straps are inserted and automatically secured.

The belt mechanism will allow the pilot to move forward (to reach a switch on the instrument panel). On older assemblies this is achieved by the operation of a release handle on the seat arm. When the pilot moves back the harness reel mechanism will "ratchet" the pilot back into position.

On inertial reel assemblies the pilot is allowed to move forward without the need to operate a lever. The inertial reel will lock (depending on type) when either the aircraft slows down too quickly or the pilot accelerates forward too quickly. (This can be tested by pulling the shoulder straps forward sharply - and the belt should lock).

Passenger Seat Belt

Consists of two lap straps (but may have other straps also) with metal snap hooks fitted to one end of each to attach to the seat structure. By regulation must be suitable locked in.

\*\-&3 QUICK RELEASE ATTACHMENT

RELEASE BUCKLE

- - - . -

Fig. 34 PASSENGER SEAT BELT DETAILS

At the other end of one is a quick release mechanism and the other one has a snap-in fitting. One of the straps is adjustable for length. They are made of nylon webbing. Figure 34 shows a typical passenger seat belt. Note the quick release connection a t the seat end of the belt - which should be inspected for correct engagement and locking.

Servicing

The release mechanism and the adjusting mechanism should not slip under load, and the release mechanism should release the harness with the minimum of effort against a moderate load. These tests can be made by sitting in the seat with the release mechanism locked in the normal way and loading the straps by straining against them. (The release buckle on passenger lap belts should not release before the 90" position is reached).

Note. When the release mechanism consists of a spring down flap and cam mechanism (passenger belts), there may be a tendency for some slip to occur if the flap is only lightly closed. The belt should not slip significantly before it is gripped firmly.

The cables and straps of forward release mechanisms should be inspected for kinks, entanglement, and fraying. Control lock mechanisms, pulleys, Bowden cables, inertia reel assemblies etc, should be examined for security, wear, freedom of movement, corrosion, functioning and be correctly lubricated.

Check that the webbing is intact and that no stitching has come lose. Lose stitching may be re-stitched at user unit level. It is important to check the entire length of the webbing that runs over the back roller on the seat, on inertia reel type belts.

If any member of a belt is found to be defective the complete assembly should be I replaced.

- -

When not in use the harness should be properly stowed. he best way to do this is to fasten the belts in the normal way, take up all the adjustment and lay them on the seat.

Check that all operating instructions for legibility, eg "Pull to Release".

Rep air

Generally it is better to return the complete belt or harness to the manufacturers but some-operators may have approval to carry out their own repairs. This is usually confined to stitching and fitment replacement.

Cleaning

The material should not be cleaned using solvents. If it is necessary to clean the material an acid free soap and warm water solution should be used. After washing rinse well with clean water and dry with warm air.

Colour Change

If the material is to by dyed (to match the cabin decor) permission must be obtained from the belt manufacturer.

Proof Testing

After repair, and whenever specified in the maintenance schedule, that part of the belt using material in a load bearing capacity should be proof tested to a specific load. The load and the method of testing is specified by the manufacturer, but in general the test load should be '/z of the certified strength of the member under test. The release should work at 1.5g.

Storage

Safety belts and harnesses should be stored in clean, dry conditions at normal room temperature and away from direct sunlight. Harnesses, other than new articles in their original packages, which have not been tested in the last 12 months, should be tested before installation. The storage period for new original packaged equipment before they require testing is specified by the manufacturer.

Records

A record should be kept in the bay of all work carried out on the harness. ~ a r n e s s k s should be serial numbered to identify them against their record card.

IN FLIGHT ENTERTAINMENT (IFE) , l 1 I I

IFE is becoming big business arid important factor in detefmining who flies with whom pdi-titularly on m e d i u ~ a n d long haul flights. It is also a cause for concern-amongst designers as electrical generating systems are being pushed to their limits for current capacity. This means that the slightest damage could cause overheating and a fire risk.

It is a legal requirement that IFE (and other equipment such as lap-top electrical supplies, in-flight telephones etc) should be so designed that no adverse affects are felt by the aircraft or aircraft systems - due to such things as electrical supply system overloading, electromagnetic interference etc.

SAQ 2

What do you feel constitutes IFE and what are your responsibilities towards it as a B 1 EASA66 licence holder?

The systems fitted to aircraft can vary considerable depending on the operator, whether the aircraft is short haul or long haul, whether it is a budget carrier and what class of seat is being considered - and the manufacturer of the IFE equipment.

rnoodull l A-884

They can broadly be divided into:

* Analogue systems. * Split analogue digital systems. 2 Digital systems - real time. x Digital systems - on demand.

Analogue Systems

This is usually some form of background music. It will come from a tape deck or cd player and the analogue signal sent direct to speakers in the ceiling/walls of the aircraft. The pilot/crew can interrupt the normal service to talk to the passengers on the same system - again all analogue.

, Analogue - Digital Systems

Some aircraft are fitted with fixed conventional television sets spaced at regular intervals down the centre of the aircraft in the ceiling (some may be retractable manually). Some aircraft have retractable flat screens above the passenger's heads that extend automatically when the system is powered - spaced at every 2"d

or 3 r d seat row. The television screens receive an analogue vision signal from a - -

video tape recorder.

Because some passengers do not what to watchllisten to the film the sound is sent to each individual seat so the person can select sound if1require.

MULTl CHANNEL AUDIO TAPE

ANALOGUE TO DIGITAL

CONVERTER

SPEAKERS SEAT

4

1 . 1 CHANNEL SELECT

DIGITAL TO ANALOGUE

t f f ! CONVERTER I

+ MULTIPLEXOR

'TO ALL OTHER SEAT GROUPS

Fig. 35 SIMPLIFIED MULTIPLEXED SYSTEM

To get the sound to the seats the analogue sound from the video tape is put through an analogue to digital (A to D) converter and then through a multiplexor to be sent to the individual seat electronic units (fitted under each seat). After demultiplexing and conversion back to analogue in the seat electronics unit, the signal is sent to speakers in the seat arm-rest or to a head phone connection in the arm rest (for speakers in the head phones). Also multiplexed into the system, usually, are sound channels from a multi-channel tape recorder.

Digital Systems - Real Time

This may use a single projector (three gun primary colour projector) to put the image on a single screen for each cabin. The sound is transmitted to the seats in a similar way to that described above.

On some aircraft each seat is fitted with a flat screen. This is housed within the seat arm rest or in the back of the seat in front. The channel and volume contro' selectors &e housed within the seat --or -- on an attached hand-set.

I 1 I

A s many as 15 vision channels ark provided by a video reprodu~ing unit. These are multiplexed and sent down a common line that runs within1 the seat tracks fore and aft. The channels are de-multiplhkd by the seat electtonics unit, and one channel (as selected on the seat~ontrol - unit) is shown on the flat screen.

I / ,

The sound signal is multiplexed in a similar way. All sound ahd vision is in real time, it cannot bgplayed back or @ut on hold! I

, - --

Digital Systems-- On Demand ---PA ..

These are vision and sound channels sent to the seat as digital signals from a bank of cd players. There is a cd player for each film channel and for each sound channel. Each cd player has a number of laser beams for signal pick-off - one beam for each seat.

This means that the passenger can select any of the channels (any one of the cd players) and have control of his/her own laser beam. This means that a film can be viewed a t anytime. It can be stopped at any time (eg for lunch) and restarted at the same place later. During this process the particular cd continues to play, it is just the laser beam that is controlled.

The system will also have games available from a computer.

A Simple Multiplexed System

Here I will explain the basic principles of a digital multiplexed system - starting with the signal inputs to the sound/vision output. Remember the whole idea of multiplexing is to send many signals down one transmission line with a consequential weight saving on a multi-line system.

Using figure 36 with a limited number of channels (4 audio signals and 1 vision) signal as an example.

The audio tape deck produces four separate entertainment channels. These four signals are converted from their present analogue form to a digital form by an Analogue to Digitdl Converter (AID Converter) - not needed if the original signal is digital.

These four signals now take turns to be transmitted down the one line and this process is carried out by a multiplexor. This is similar to many trains travelling downone main line, one behind the other, but leaving the station from different platforms'. I n this case, however, it is carried out at very high speed with electronic devices and no moving parts.

analogue vis~on signal to projector

VlDlO

AID CONVERTER

AUDIO TAPE

Fig. 36 SIGNAL INPUTS

AID CONVERTER digital

The video for the film produces two analogue signals - one vision and one sound. The vision signal is sent direct to the television set or the three gun colour projector (both fitted in the cabin ceiling). The sound signal is put through an A/D converter and then a sub-multiplexor to wait its turn for transmission (a nanosecond or two).

7

-

These sound signals are now sent to all seats (normally from front to rear) to the seat electronics units housed under the seats (figure 37).

Within the seat unit the signal is first de-multiplexed. This is the reverse of multiplexing and puts each separate signal into its own separate "platform at the station".

MULTIPLEXOR

- - -

analwue , - d~gital sound signal

to seal units

- 4

- - SUB-MULTIPLEXOR

The signal is then converted to an analogue signal by a digital to analogue converter (D/A converter) ready to be understood by the human ear. Which signal is actually processed depends on the selection of the seat selector unit - made by the passenger.

Once the signal is processed it is sent to a speaker either:

(a) In the arm rest where air phones (small rubber tubes) channel the noise to the ears.

OR (b) To speakers in a head set connected to the system via a jack in

the arm rest.

analogue wund to seat selector

O/A CONVERTER P

1 2 3 4 , film sound

- l(-)LTIPLWoR -

1st SEAT 2nd SEAT 3rd etc

I I I

I ,' Fig. 37 SIGNAL RECEPTION AT THE SEAT I

What follows - - - - - is - - a description ofa system - as used in some B747-300 series (remember, the system actually fitted to a n aircraft depends = the operator). It uses similar principles to those explained above.

Note. Video recorders, tape reproducers, cd units etc are housed in the passenger cabin in a locker (in some cases an overhead locker). Also in the locker will be a selection of videos, cds etc. There is often a monitor so the attendant can check that all players are working correctly. The attendant will pick a cd or video with the appropriate film on depending on the menu - which is distributed to the passengers.

A Typical System

You need not remember the details of this system but you should understand the principles and be able to test a typical system.

The system is a digital time-division multiplexing system. Seat controls are provided to allow each passenger to individually select an audio channel and control audio volume.

rnoodull l A-888

One audio channel is reserved for passenger address by the flight crew. When transmitted, the passenger address audio overrides the program selected at the passenger seat, so that passenger address is heard at all seats regardless of the channel selected by the passenger. This signal is also received by the cabin speakers (including those in the galleys and toilets).

I - P A I I ~ W I I

I 1111 r US1 I U M l I I t l I X f I A W I O - C I L A N N f L

Z O N E A Z O N f C S l B M U l I I - S M M I I I - C I l B h U J l l l -

P L I X f I I L E X I " r L I X f l

r------- I I, C. A N D 0 I T

I Y P I C A L Z O N E L ----------- --------,-----,-- -- - - -

Fig. 38 TYPICAL MULTIPLEX SYSTEM

The system consists of three major assemblies: Main Multiplexor; Zone Sub- multiplexor, and, depending on the seating arrangement, a 2, 3 or 4 seat Electronics Unit for each seat group.

Audio inputs to the Main Multiplexor are provided by a multi-channel magnetic tape reproducer or similar source.

In a typical system, 10 channels of a multi-channel stereo/monaural tape reproducer are connected to the Main Multiplexor audio input connector by twisted pairs of wires.

In turn, the Main Multiplexor is connected with the zone sub-multiplexors via a single coaxial cable. Up to four additional zone-generated movie audio channels and a passenger address audio channel are connected to each zone sub- multiplexor.

Each zone sub-multiplexor is interconnected to all associated seat electronics assemblies within the zone via a single coaxial cable.

The system is a digital time-division multiplexing system that uses Pulse Coded Modulation (PCM) to transmit audio over a single coaxial cable.

Prior to transmission, each audio channel is sequentially switched into a sample and hold circuit a t a rapid rate. The sampling rate (25.6Kh) is high enough so that no information is lost.

The amplitude of each sample is then measured and compared with a scale of discrete values. A digital word is generated (eg 0 1 100 1 1000) which represents the amplitude of each pulse, and a digital pulse train is produced. This process is called PCM.

Seat Electronics --- - Unit I -

-- --

The Seat Electronics Unit (SEU) is involved in the operation of both the Passenger ' I ~nte r ta ibment System and the Passenger Service System. These systems are

independent of each other.

The basic objective of the SEU is to deEiultip1ex the digital signal and convert it into its original analogue audio form. , I 1

I \

I ! I

1 1 I 1 I

MOSFET Module I ~ I

- -

This module is the heart of the SEU, wher-e data is demultiplexed and stereo entertainment, passenger address override, and channel selection functions are controlled.

The analogue output from the D/A converter is directed to the output audio processor modules according to the channel selection at each passenger seat. When a passenger selects a channel it is remembered by the MOSFET. Each time that channel appears in the multiplexed scheme, it is switched to the audio output lines for that seat.

When a passenger changes channels, new selection information is generated and stored in the MOSFET memory circuit.

Self Test Modules

These are two distinct modules that operate as a unit to accomplish the self test function. They are used to check transducer continuity and audio power output circuitry common to passenger seat groups. A successful self-test will light the appropriate attendant call lamp.

Fig. 39 TYPICAL SEAT ARRANGEMENT FOR A MULTIPLEXED SYSTEM

-- - ,

The passenger Service System provides a means of multiplexing control signals in a manner such that the passenger-to-attendant call and reading light on/off are controlled from each passenger seat. The system also provides attendant stations with test and reset control functions.

The system comprises the following components: The TimerIDecoder; Passenger Service Unit (PSU) Decoders, and a Seat Electronics Unit.

The components continuously monitor the state of each passenger's seat control unit switches. This task is accomplished by a Seat-to-Seat and a Seat-Group-to- Seat-Group sampling of each zone. The sampling is under control of a TimerIDecoder (T/D), that provides timing, synchronisation, and acts as an interface between each seat group and its associated passenger service unit.

(Note that all data transmission on computer systems is strictly controlled using an "electronic clock producing a clock pulse.)

Each seat group has a Seat Electronics Unit containing, in addition to the entertainment de-multiplexor, and a seat coder for the PSS. The seat coder accepts passenger control function inputs from each seat in the seat group and timing and data from the T/D. The seat coder encodes the passenger control function, and during a specified time period, the encoded command is transmitted to the T/D.

rnoodull I A-891

Upon being received a t the T/D, the circuits immediately transmits the received data to the PSU Decoder. The PSU Decoder, which has been synchronised with the transmitting Seat Electronics coder by T/D timing, accepts the data, and decodes it, and depending upon the original command, will control the state of a reading lamp or the attendant call lamp (ie switch it on or off). Two TimerlDecoders are utilised to service each passenger cabin zone. The T/D1s in a zone interface with the attendant's station for that zone.

SAQ 3

You are called to a n aircraft that is about to depart by the Cabin Services Director (the person in charge of all the cabin services on a large aircraft - different operators may use different titles), who reports that the PA system is not functioning - list your actions.

CABIN LAYOUT

SAQ 4

Aircraft cabins have to be laid out in a particular way. Why?

, Operators have some latitude in! the design of the cabin interior and will often have their own cabin interior manuals which need to be cons'ulted in specific cases. Here we will cover, mostly,-the-regulitions that apply to cabin interiors.

In general the cabin staff need to man the emergency exits during a n evacuation procedure, therefore, you would expect to find the attendants seat located by the door in the zones they are responsible for. Seats, galleys and toilets will need to he positioned for ease of access both in normal situations and in emergencies.

Galleys are located in clusters around the cabin in such a way as to be able to cater for the whole of the aircraft. There are no legal restrictions on galley location, but by default they are usually placed at the front and/or the back of t h e cabin for medium sized aircraft. On larger aircraft they are placed a t mid positions in the fuselage.

Toilets tend to be spaced at the ends of cabin areas on medium sized aircraft with additional "toilet blocks'' being positioned a mid points along the cabins. Each class zone will have a t least one toilet. Also, a t various locations around the cabin there will be various closets, some for emergency equipment, some containing IFE equipment and some for coat, hat and luggage storage etc.

Minimum Seat Pitch

CAP747 (was previously in ANs) shows the minimum distance between the seat cushion and any other seat back or fixed structure is 26" (66cm). This is shown on figure 40 as dimension A. Dimension B must be a minimum of 7" (18cm), this is the distance the passengers leg has to move from thigh support to seat back.

The vertical separation between seats must not reduce less than 3" (7.5cm). This includes the seat in the reclined position.

Around some emergency exits (typically overwing small and slightly larger exits) the clearance has to be increased. The vertical separation between seats at these locations must not reduce below 10" (25.5cm). In addition the seat backs at these locations must not %re& forward' as other seat backs do.

D~menslan A datum polnl haghl abow

lz:7wb -

Range lor dlrnension A datum mlnl

Fig. 40 SEAT DIMENSIONS

The reason for the break forward action of the seat back is to prevent head injury of passengers as they impact their head on the seat in front under heavy brakmg etc. The seat back breaks forward to move out of the way.

Around overwing emergency exit points, studies have shown that, in an emergency, people will climb over the seats to get out. If the seats were to fold a crush would ensue which would hamper the evacuation, so these seat backs will not fold forward to prevent this sort of action.

- 51 -

rnoodull l A-893

Note. These dimensions are not the seat pitch as the seat pitch is usually measured from the feet of the seat and so different dimensions might be found.

CS25 specifies aisle width which will depend on seat numbers. Maximum numbers of seats abreast in a single aisle aircraft is 3.

Cabin Emergency Exits

The minimum number and size is specified in CS25 and are related to the seating capacity of the aircraft:

Passenger seats 1 to 9 10 to 19 20 to 39 4@to 79 80,,to 109 110 to 139 140 to 179

Number Type 1 (iv) 1 (iii) 2 (ii) and (iii) 2 -- (i) and (iii)

-3 "(i) and (iii) -

31 (i') and (iii) 4 (i) and (iii) 1

' I

As s e b numbers continue to rise o exit numbers go up to include larger exits until, aircraft with seating capacitiFs%bove 299, have all exits a s either class (i) or type A. \ I /

, - - - I

I

I I N For techbical details of doors and exits pleise see the books in this series entitled Structures. I I

, -

- - - - . .

CABIN DECOR AND FURNISHINGS

Flame Proofing

All materials must be flame proof to the standards laid down in CS25.

CAP747 GR 13. (Was AN58). States that furnishing materials (seats, berths etc) should be flame resistant to CS25 standards. The material must be flame resistant after 3 representative cleaning processes - using three samples with the flame on the material going out after a specified time.

CAP747 GR 14. (Was AN61). Applies to public transport aircraft over 5700kg with 20 or more passengers. Deals with wall and ceiling panels. This states they must meet flammability and smoke emission tests as laid down in BCARs or EASA25.

Note. CAP = Civil Air Publication published by the CAA and GR = Generic (general) Requirement.

Flame Proofing Testing

CS25 states that 3 specimens have to be tested and details of flame areas, propagation rates, longevity rates and specimen orientations are comprehensively covered. CS23 specifications are simpler and states that for compartment panels the flame length should be no longer than 15cm and maximum duration should be 15 seconds. The figures for upholstery are 20cm and 15 seconds.

Cabin Linings and Partitions

Designed to provide comfort, aesthetic quality, soundproofing, cabin insulation, acoustic and thermal insulation, and mu st meet fire resistant regulations. They are fitted direct to the cabin structure and will cover service wiring, ducting, pipelines, components and insulation.

The partitions separate the various compartments, and also serve as mountings for various items of equipment. They may be transferable to alternative positions in the aircraft to provide for variations of the seating arrangements.

CEILING PANELS REF 25-23-41

I CLOSlNG COVERS R E F 25-23-41

I DADO PANELS R E F 25-23-44

Fig. 41 CABIN COMPARTMENT LININGS - A330

Provision is made in the construction of the partitions to carry stewards' seats, oxygen equipment, passenger life-saving jackets, fire extinguishers etc.

The partitions are located and secured in position by foot mounting brackets which pick u p on the seat rails, or shoe-mountings, and attachments which pick up with the aircraft side-wall or roof structure.

Disconnecting the foot, ceiling and wall attachments enables the partitions to be moved to their alternative position.

Ceilings

These are fitted to a framework suspended from the top of the fuselage structure, or on some aircraft, fitted to the top of the fuselage structure direct. The ceiling may contain:

k Projection equipment. * Passenger Service Units (PSUs) . * Luggage lockers. A Television sets (CRTs/ flat screens). -k Lighting. * Insulation.

I

PANEL CONNECTIOH

SECTION

B- 8 EXAMPLE OF UIHOOU PANEL COHHECTION

Fig. 42 SIDE WALL PANELS - A 3 3 0

Side Wall Panels

May be of similar construction to partitions/ceilings and are attached to the sides of the fuselage. Attachment may be by the use of special clips or slide-in fittings. Some panels may be fitted into position by the use of screws fitting into captive nuts on brackets fitted to the fuselage. Gaps between panels are covered by decor strips. Panel removal is normally by removing the decor strips, releasing the quick release fasteners or screws and pulling the panel out.

Panelled areas may contain lighting, window reveals, kick-boards, decompression blow out panels etc. May be made of one or more of the following materials/ techniques:

All Metal Construction. These normally comprise aluminium alloy sheet parts, formed and riveted to a framework of light-alloy angle members, channel sections, brackets and cleats.

Honeycomb Construction. The honeycomb-material is sandwiched and Reduxed (bonded) between aluminium alloy sheets riveted to light-alloy boundary members.

Glasscloth Construction. Laminated glasscloth skins are mounted on each side of a framework. The cavities formed between the sluns are filled with Dufaylite, or similar material such as composite honeycomb. --

Composite Construction. Composite materials in either solid, sandwich, or honeycomb construction.

-

Window Reveals

Fitted to the sides of the cabin at each window location and may be fitted with:

(a) A transparent decor panel - to protect the actual aircraft window. (b) A pull down blind - for the convenience of passengers. (c) A lamp assembly - on most aircraft. (d) Attachment fittings. (e) Curtains in some locations.

The reveal can be clipped into position and so removal is straight forward. It will need to be removed for maintenance on the transparent panel, lamp fixture shade etc.

INNER

Fig. 43 ' ~ P I C A L ~ N D O W REVEAL I I I

-- / I \

.. I /

Cabin saggage Stowage I --

I I i I

Cabin baggage may be stowed uhder the seats and in cabin blaggage lockers. These are usually overhead lockers (on large aircraft) but ma$ de beside the seats or in front of them where seats are at the front of a cabin.

- -- -

Overhead Lockers

These must conform to strength, fire and smoke requirements as laid down in C" 25. They are normally of composite or plastic construction.

The locker should be checked:

1. For cleanliness and damage. 2. That it will shut and lock securely. Check that it will open. 3. That when lowered (for those where the container comes

down) it lowers slowly with a reasonable load applied. Change the snubber unit if it lowers too quickly.

4. That the travel stops function correct correctly. 5 . For corrosion (metal parts), securiv of attachment and correct fitting

of all component parts.

LATCH

Fig. 44 OVERHEAD LOCKER

Passenger Service Units (PSUs)

These are usually situated in the ceiling above each row of seats.

Equipmept, for passenger use, is fitted to the panel. It is supported by a catch (operable by maintenance staff) and is hinged on one side. I t is latched shut by a catch assembly which allows for easy access for maintenance.

OXYGEN SUPPLY l l H E 6ASrER AIR SUPPLY

READING LIGHT ASSY

USE HEX HtAD WlInCH

Fig. 45 TYPICAL PSU

The complete unit can be removed by releasing the catch mechanism (on some units this is carried out by inserting a small screwdriver into a slot or turning a small hexagon headed release screw). This will allow the complete unit to swing down on its hinges.

After disconnecting the service connections the unit complete can be lifted off its hinges and removed.

The PSU may contain some or all of the following equipment:

Fresh air supply (Gasper Air). Cold air supply to the passenger louvres. On some aircraft these may not be fitted. Oxygen supply line. Supplies gaseous oxygen to the passenger masks from an airframe mounted system in an emergency. This system is not fitted if a chemical oxygen generator is fitted. Electrical power supply - for passenger reading lights etc. Chemical oxygen generator. Fitted on some aircraft to supply oxyge* to the passenger masks -- i n an emergency. (See bok-on Cabin Conditioning and qre,ssurisation). I p

Oxygen masks. ~hesefa l l out automatically if cabin altitucle rises above a certain value (say 10,000ft). Reading lights. On some aircraft may be operate+ directly by a switch on or near the passenger seatcon many aircraft 'the - passenger --

reading light switch ( ~ K e n operated) sends a code down a common signal line to the en'd of the cabin then back to the IPSU. The PSU decoder recognises itS own code and switches thk light on or off as

I appropriate. I

Attendant call light. Lights when the attendant call button is pressed on the seat arm. Also lights a lamp at the attendant's position. It may also be connected to the inflight entertainment system so that on TEST MODE it will light if the seat is receiving the audio signal. When IFE TEST is pressed all attendant lamps should light. Speakers. Information signs (toilet engaged, fasten seat belts etc)

CARGO

'Cargo' may be divided into:

* Passenger hand baggage. 'Carry on' baggage taken into the cabin. * Passenger hold baggage. Put in the aircraft hold or in sealed areas of

the cabin either in specialised containers or individually. * Cargo - with consignment documents containerised or crated. Put in

specialised cargo aircraft or in the hold of passenger flights or in the cabin of 'combi' aircraft. A combi aircraft has part of the cabin partitioned off for cargo only.

* Dangerous goods - where special requirements have to be met with regards to cargo handling, type of aircraft, special equipment, route flown, no passengers etc.

Passenger Luggage

Apart from cabin baggage, luggage is handed over with the tickets at the check-in desk. After weighing and taggng (and security checking) it is placed on a conveyor belt from where it is loaded into containers or loaded individually into the aircraft using conveyer belts.

Loaded containers are towed or driven on trucks to the aircraft in the ramp area where they are raised to aircraft sill height for loading.

Cargo Consignments

These are assembled on pallets/cont&ners and bound by netting to prevent movement in transit. Mechanical rollers carry the containers from the make-up floor in the cargo shop to the aircraft via a truck or trailer with a mechanical lifting roller platform.

CARGO CONT REAR CARGO DOOR

AIR CARGO LOADING SCISSOR LD CONTAINERS TERMINAL BRIDGE PLATFORM

Fig. 46 CARGO LOADING

Loading - Freight

Pallets can be loaded from ground level by fork-lift trucks or 'scissor' trucks. The loading doors may be on the side of the aircraft or the front or rear. Once the pallets are inside the aircraft the loader/operator can manoeuvre them into position either manually using rollers or mechanically using motorised wheels in the floor. These motorised wheels retract into the floor until required for use.

- 59 -

moodull l A-901

The B747F for example, has a mechanised cargo handling system on the main deck. Pallets or containers are moved in the aircraft on a motor driven system by two men, who can load the whole aircraft in about 30 minutes.

Some aircraft are interchangeable as passengerlfreighter and are usually called a "combi". They are designed so that blocks of seats plus galleys can be removed, and roller equipped freight floors fitted.

Loading - Baggage

The system will vary with the aircraft, eg on the B747 the system is similar to the above and the pallets are loaded from special trucks equipped with elevating platforms with motorised wheels in the floor.

On the B757, for example, each cargo hold is fitted with a number of telescoping f la t -bedodules . To load the aircraft -- they are moved by an electrically operated rotary, actuator forward to the door.J'heyXare all telescoped into each other opposite the cargo door. I , 1

,

I

BULK CARW M Y -

DOOR v

Fig. 47 TYPICAL BAGGAGE CONTAINER LOADING

Each flat-bed module is mounted on rollers and after it is loaded manually a t the door entrance it is moved back by the screwjack - exposing the next empty flat- bed. The next flat-bed i s loaded and so on.

On large aircraft loose baggage is secured using traditional netting and the netting secured using the tie down fittings.

On small aircraft the cargolbaggage is loaded by hand into the hold and secured.

Notes:

1. For aircraft of 19 seats or less cargo and baggage must be weighed and placed in the aircraft in such a way as not to upset the C of G. This may be done by working out the moment arms, or on small aircraft, placing the baggage on the C of G. Aircraft must not exceed their MTWA. With most operators a standard form simplifies the procedure.

2. For aircraft with 20 or more seats the weight and C of G location may be calculated using the pro-rata standard masses stated in EASA OPS. When packing luggage items into standard containers - when full the total weight of the container can be ascertained.

3 . All baggage/cargo must be properly secured using the securing points on the airframe.

4. In general - and especially on small aircraft - heavy items of baggage / cargo are loaded a s close to the C of G as possible with lighter items loaded progressively further away - fore and aft.

5. On small commercial aircraft passengers as well as baggage are weighed, and the passengers are seated according to weight with the heaviest nearest the C of G and the ligHtest further away - all recorded on -

the aircraft loading record.

Powered cargo handling systems is a luxury not all aircraft have. Most aircraft of moderate size have some form of astsistmce for the loading and unloading of cargo (a powered conveyer belt for example). Describe below is a typical powered system. Check your aircraft and see what system it has.

A cargo bay will be split into a number of areas such as a Containerised Cargo area and a Bulk (or loose) Cargo area.

With containerised cargo all the goods are loaded either onto pallets or more often into specially shaped containers. If these containers are not standard containers then they will have to get CAA approval separate from that of the aircraft.

Bulk cargo is usually located at the back of the aircraft and will have oversize and lose articles restrained by nets.

Figures 47 and 48 show standard LD containers as used on the B747. Containers are made from aluminium alloy, sometimes with side tie-down covers and designed to lock into the appropriate aircraft hold.

- 61 -

rnoodull l A-903

Fig. 48 STANDARD LD CONTAINERS IN THE B747

Fig. 49 CARGO HANDLING

A Containerised Cargo System

The cargo handling system has the following components (figure 49 and 50):

* Guide rollers * Sill rollers * Roll out stops * Lateral and longitudinal power drive units * Ball transfer mats * Lateral guides

* Roller trays * Centre guides * Pallet locks * Control panel

System Operation (figures 50 and 5 1)

Power is provided for the cargo handling system when the cargo door is fully open and the system is switched on a t the control panel. The supply is usually via the ground handling bus, which is live as soon as ground power is plugged in. This is useful for cargo handlers who don't need to provide full aircraft power to load the bays.

LATCHES,

w LATERAL GUIDE - SMALL

POWERED D R I V E UNIT

Fig. 50 COMPONENTS OF A TYPICAL POWERD SYSTEM

rnoodull 'I A-905

With power on the cargo loader is driven into position in front of the aircraft cargo door and is adjusted for correct height. The containers are rolled off the loader (using the in-built drive wheels) and into the aircraft hold onto the ball transfer mats. Small sill rollers ensure that any slight height error doesn't result in the container hitting the sill of the cargo door.

With the container now on the ball transfer panel the operator will move a joystick on the control panel (fitted to the side of the fuselage just inside the door) to power first the lateral Power Drive Units (PDUs) to bring the container into the aircraft. Then, still using the joystick, the longitudinal PDUs will be powered to take the container down the cargo bay, along the roller trays.

Fig. 5 1 CARGO HANDLING CIRCUIT

When the container is in the correct location (normally at the furthest end to one side of the hold) power will be removed and the pallet locks installed to prevent it moving.

The lateral guides prevent the container from skewing down the bay as they enter the aircraft. They are normally up and hence straighten the container. A s the longitudinal PDUs are selected the lateral guides momentarily drop to allow the container to pass over the top.

The roll-out stops prevent the cargo from rolling out of the bay. As the container enters the aircraft the roll-out stop is automatically (by the weight of the container) pushed down. As soon as the container has passed the rollout stops spring back u p and prevent the container from leaving' the aircraft. The mechanic can select them down when unloading.

Various rollers are placed in the system to keep the cargo straight and stop it from impacting the door, the aircraft or its equipment.

The PDUs can have their power removed individually to prevent them from turning when cargo is already installed in them. The furthest bay from the door have containers in them whilst the rest are still being loaded for example. In addition all the PDU's have a manual disconnect facility to allow for manual cargo loading operations.

Figure 51 shows the power supply and logic circuits for the system. Take a little i

time and study the AND gates and how they work; note the power supplies and the location of the PDUs.

Maintenance of the cargo handling system is fairly straight forward and trouble shooting guides are given in the maintenance manual. Most prdblems with this system are usually through misuse. The PDU tyres need checking for wear and condition. The door sill and seal depressor will often become damaged - as will parts of the cargo hold liner. If any powered item fails to work then the problem is tackled the same way as any electrical problem - visual first followed by a logical approach to rectifying the fault.

-

Bulk Cargo

Bulk cargo is handled in the old fashioned way by using manual labour. Loose and oversize iterns are put in this area as well as any livestock. If livestock has been loaded a n entry in the Tech Log is required to inform the crew to keep the bay heated. All the items are restrained using nets and ties (see figure 5 1).

Dangerous Goods

Almost all operators now have specialist cargo loaders that deal with the packing and loading of cargo, but as a licensed engineer you still need to be aware of all cargo loading including any 'dangerous goods'.

The Air Navigation Order and Regulations (the ANO) gives details of goods that are not permitted on-board aircraft.

I I

Fig. 521 BULK CARGO AREA i

The list is extensive and includes anything that might harm the aircraft, its occupants, or those beneath the-flight-path, or those a t or near-the take-off and landing airfields. Examples include:

* Munitions of war * Weapons * Flammable liquids * Mercury * Chemicals * Radio active materials * Bacteriological agents

There may be times when dangerous goods have to be transported and where this is the case the consignment needs to be packed in the specified manner and the shipping information must be presented to the aircraft commander at the same time as his load sheet. (Usually airlines do actually combine the dangerous goods form and the load sheet and the captain must sign this before departure).

rnoodull lA-908

UPPER ATTACHMENT F I T T I N G

LOWER ATTACHHENT FITTING (DOOR NET)

Fig. 53 BULK CARGO AREA TIE DOWN - EXAMPLE

The secretary of state is required by the A N 0 to specify certain articles and substances as dangerous goods and what may and may not be carried. He/she will also specify the exact conditions t6 be complied with if the dangerous goods are to b e carried - labelling - packaging - equipment - passengers? - loading - route flown etc.

Cargo Holds .

For the classification of cargo holds please refer to the book in this series entitled "Structures - I".

AIR STAIRS

These are generally considered to be stairs for passenger use that are carried on the aircraft. They may be built into the design of the door so when the door folds down the steps can be used. With this type the hand-railslhand-ropes deploy automatically. The door may be fitted conventionally to the side of the aircraft or beneath the fuselage at the rear. The door is designed not to touch the ground.

Another type of air-stair is pulled out from the side of the fuselage complete with handrails and these are fitted to the sides of the door aperture before the passengers use the stairs. These air-stairs may be stowed in a separate compartment under the cabin floor with its own separate fuselage closure panel or may be pulled out from under the cabin floor through the door aperture once the door is opened.

rnoodull l A-909

HAND RAILS UNFOLDED 8 ATTACHED TO

21 n) I I\ FUSELAGE \ n\

/ HAND RAILS UNFOLD WHEN DOOR IS OPENED

- - Fig. 54 AIRSTAIRS -

L \ --

'1

\

' I

That concludes this section andithe book. Take some time now ko recap and consolidate your knowledge with the following exercises. The adswers can be

1 found lin the text. - - /

, I -- I - -

\

~xe rc i s e s , I I I

' I

1. &here is information fourid on minimum seat pitch? 2. What t@es of cargo holds are there? 3. -Explain the operation of a containeri6ed rnotorised cargo bay system. 4. What is the purpose of the roll-out stop? 5. What are dangerous goods? 6. Can the aircraft carry them and how?

ANSWERS TO SELF ASSESSMENT QUESTIONS

SAQ 1. The following checks should be carried out. Inspect attachment lugs, fittings, push/pull rods and door/slide operating mechanisms for security, damage, corrosion and wear. The slide pack itself should be checked for signs of chafing, damage and security of attachment. The transportation pin should be checked that it is removed. Check that the girt is securely attached to the girt bar.

Check all lacings and Velcro strips for correct fitting. Close door and check the ' girt bar for correct operation and that it latches correctly to the floor fittings.

Check also for security and damage.

Check the bottle pressure and date life. With the decor panel fitted check that bottle pressure indicator is visible through inspection window and all handles operate freely. Check the door opens and closes correctly with correct clearances and that arm/disarrn indicators show. Check that all electrical indicators work.

SAQ 2. IFE is basically all the electrical equipment that goes to the seat, including video and audio signals, electrical power, audio and vision signal equipment and data handling equipment. In other words all the system from the box that reproduces the sound and/or vision to the speaker at the other end. Provided you are B 1 qualified with appropriate certification authorisation from a EASA145 company then it is your responsibility to look after these systems and clear the CRS. (AN 3 refers). Responsibility covers avionic LRUs and electrical power supply systems where simple BIT tests or using simple test equipment can verify the s e ~ c e a b l e of the system. Of course the seat it-self is 'all yours7 anyway.

j ," SAQ 3. Gain information from the crew as to the details of the defect. Confirm the defectandsee if it can be isolated to one zone. Check the AMM fault finding section. Carry out a system BITE check - using the equipment BITE or any aircraft centralised fault location equipment. Check the boarding music is heard from all speakers (if it uses the same speakers).

If all speakers work then it is a fault in the dedicated part of the PA system.

Check that power is available to the system, if it is not find the fault and rectify. Check the wiring, replace if open circuit. Check the individual microphone handsets - change any that do not work. Check the PA switching units and any other electronic units - change if necessary. Check the system multiplexors (if fitted), change if defective. If the boarding music is not heard from all the speakers then c-heck the main multiplexor (if fitted), change if defective.

If only some speakers do not work then it is almost certainly an open circuit in the line downstream of the last 'good7 speaker - locate and rectlfy. If the fault cannot be cleared in a reasonable time scale check the MEL (Minimum Equipment List) to see if the aircraft can fly (it probably cannot). Ifit can the crew must be made aware of the problem and it must be recorded. If it cannot be fmed and the MEL does not allow flight then the passengers will have to disembark.

During all this the pilot will be wanting to use his/her take-off slot and might request that there is no time to complete all the checks with the passengers on- board - there might not be access because of passengers anyway and they might be getting restless as well.

SAQ 4. First and foremost it is to allow the quick and successful evacuation of an aircraft in an emergency. Secondly it is for the comfort and convenience of the passengers. Standard layouts within a fleet also allow the cabin staff to move from aircraft to aircraft without the need for re-training.

LBP Dec 03 - Feb 04 - April04

Addendums module 1 1A book FIRE PROTECTION SYSTEMS pending amendment action in response to student feedback after taking the CAA examinations.

*** The integrity monitor of the Systron Donner fire detection system monitors the pressure holding capability of the system.

*****

NOTE: It is possible that some of the above statements may not be too meaningful when read out of context, so it is suggested that the appropriate book/subject be read first then the information above be checked against that topic.

moodull I p i 1 2

CONTENTS

Page

Fire extinguishing systems Extinguishants Fire extinguishing system - fixed

Maintenance Fire and overheat detection

Unit type detectors \ , -Continuous type detectors -

Smoke detectors \ .. \

Cargo hold system Wheel well system 1

Pneumatic duct leak detqction system Toilet smoke and fire detection syste'ms

Maintenance - general - -- . Your activity

-,

AIRCRAFT FIRE, SMOKE AND OVERHEAT DETECTION SYSTEMS

Types Of Fire Extinguishing Systems

Fire extinguishing systems can be divided into fzed, portable, and mixed systems.

The term "fmed" refers to a permanently installed system of extinguisher bottles, distribution pipes and controls provided for the protection of powerplants, APUs, landing gear wheel bays and cargo compartments. A self- contained system is fitted to paper towel waste bins in toilets.

J A R 25 specifies the minimum capacity of each fixed system and that the contents must not be hazardous to personnel or the structure.

A portable system refers to hand held fire extinguishers provided in flight-crew and passenger compartments. (Refer to the book in this series entitled ''Safety" for details-on hand held extinguishers). -

?- - -

-, I !

A mixed system is used in some &craft baggage and service compartments. The distribution pipelines are fixed in thei appropriate comp'atment and coupled to adapter points within the crew drea to which a p,oqtable extinguisher is plugged into when required. Not very'cpmmon. 1 ,

-- - - -

\

I ; ! Extinguishants On Fixed systeks

I / 1 ~ ethyl Bromide IMB1. Boils at about S C a i d is used for the protection - of power plants: -Highly toxic andmust-not-be used in confined spaces such as cabins. The effects of breathing the vapours may not be immediately apparent, but serious or even fatal after-effects may be sustained.

Not very common, though might be fitted to engines of older aircraft where the engines are away from the fuselage.

Bromochlorodifluoromethane (BCF1. This is semi toxic and is particularly effective against electrical and liquid fires. It is used in powerplant and APU fire zones and portable extinguishers. It is gaseous at normal temperatures and condenses to a liquid at -4OC.

It has little or no corrosive effect, although halogen acids will form if its products, which have been decomposed by fire, come into contact with water. In contact with fire BCF volatilises instantly, giving rapid flame extinction.

Bromotrifluoromethane (BTM). Similar to BCF. Semi-toxic and is used in power plant, APU, and cargo compartment fire zones.

BCF and BTM are very effective extinguishants, however, they are halogenated hydrocarbons, and their discharge into the atmosphere seriously affkcts the ozone layer. They are being withdrawn and research is being carried out to find an effective replacement.

Chemical Dm Powder. Produces a dry powder discharge. Non toxic though it can cause choking. Is available from extinguisher systems manufacturers though not yet common on aircraft.

FIRE EXTINGUISHING

FACILITIES

APU EXTINGUISHER

TOILET EXTINGUISHERS

CARGO I I rp - - I

COMPARTMENT I -..-. .,--, J AL 1 '

Fig. 1 ?IF& PREVENTION .--- SYSTEMS - GENERAL A ~ N G E M E N T - I

Fire Zones

Fire zones are designated in an aircraft where there is a potential fire risk. Each fire zone will have a fire and/or smoke detector system and a fire extinguisher system. Certain fire zones will have fire proof bulkheads (engines and APUs), and fireproof linings (cargo bays). Most engines will have more than one fire zone (as specified by JAR 25).

The following are usually designated fire zones and/or fire potential hazards.

* Engines. * APUs. * Fuel tanks. * Cargo bays. * Wheel bays. * Toilets (paper towel disposal bins)

Each designated fire zone will have a fire/overheat detection system and an extinguishing system. Fire detection systems may be used for overheat warning along hot air ducts etc.

Passenger compartment linings and furniture materials, and cargo compartment interior linings must meet the fire resistant specifications as laid down in JAR 25 and ANs 58, 61 and 83. Cargo compartments are classified primarily on their location in relation to how a fire can be tackled:

* Class A Easily accessible in flight where a fire would be quickly detected by a crew member a t his/her station.

* Class B Must have a smoke/fire detection system and access can be made to any part of the compartment by a person with a hand held fire extinguisher.

-k Class C Must have a smoke/fire detection system and a built- in fire extinguisher system controllable from the flight deck.

t Class D Any fire occurring within the compartment will be completely contained and compartment volume is 1000 cubic ft or less.

~r Class E For cargo aircraft only. It must have a smoke/fire detection system and means to shut-off all ventilation

I - airflow-to/from the compartmentT- -

- - -- ,

I

WINDSCREEN DEFOG

Fig. 2 DETECTOR POSITIONS - EXAMPLE

FIRE EXTINGUISHING SYSTEMS - FIXED

Regulations laid down in JAR 25 state that aircraft in the transport category must have provision to tackle a fire at least twice in any engine fire zone (except for combustion heater fire zones). In some cases two bottles per fire zone with attendant pipework are provided (figure 7) whilst in others a "two shot" system is used with one bottle per zone and a "cross-over" pipe-work system (figure 3).

With reference to figures 3 and 4. Each fire bottle has 2 discharge heads and the fire switches in the flight deck are supplied with 28V dc from the essential or hot bus bar. When the extinguisher switch is operated current will flow to the appropriate fire bottle head detonator unitlcartridge unit/squib and to the warning lamp / indicator fuse.

The detonator unit will operate allowing extinguishant to go to the fire zone via the directional flow valve.

dc BUS BAR

Fig 3 A TWO SHOT SYSTEM (ONE BOTTLE PER FIRE ZONE)

Should the fire persist or re-start then the bottle from the other fire zone can be discharged by moving the switch to the "shot 2" position. This will cause one of the detonator units on the other bottle to operate allowing extinguishant to flow via the directional flow valve to the fire.

In an emergency most aircraft have systems that will operate all the fwe bottles automatically (except "toilet" and portable bottles).

In the drawing an inertia switch is fitted that operates if the aircraft is put through more than say 3g (as would happen in a crash landing).

INDICATOR PIN OUT

Fig. 4 BOTTLE DISCHARGE INDICATOR

On some aircraft a rubber covered crash strip may be fitted beneath the aircraft. If the fuselage touches the ground the two elements of the strip come into contact to complete a circuit - operating "all" the extinguishers.

When operated the pressure in the pipelines (on some aircraft) pushes out a small plastic discharge indicator disc on the outside of the fuselagelengine nacelle (figure 4). This allows outside verification that the system has been operated. If the disc is found missing then a check should be made on the rubber seal inside the unit.

If it is still their then it is possible that the disc has become detached by the airflow and needs replacing. If the rubber seal is missing then there is a good chance that the system has been operated - of course there should be other indications also - Tech Log entry - signs of a fire - flight deck fire handle position - fault computer etc.

Should the bottle over-pressurise due to high ambient temperature conditions then a disc will burst in the extinguisher head (to prevent the bottle exploding) and the extinguishant will be piped over-board via a discharge indicator disc. In this case all the extinguishant will be lost.

Note. All fixed fire systems are connected to the essential or hot dc bus bar.

PIN

Fig. 5 TYPICAL TWIN HEADED EXTINGUISHER

Concorde, for example, is fitted with a "two shot" (one fire bottle per engine) (figure 8) system whilst aircraft like the Boeing 747 (figure 6) are fitted with a two shot system with two bottles per engine. Study both these figures and note the general layout and operation.

QUESTION Can you think why they are different? (2 mins).

ANSWER If the engines are placed far apart on the airframe or wing, then on a "two shot" (one bottle per engine) system, too much extinguishant would be required just to fill u p the cross-over pipeline. So engines that are close together - and you can't get much closer than those on Concorde - can use the "two shot" (one bottle per engine) system, whilst engines that are far apart usually have 2 bottles per engine.

Directional Flow Valves

These are a type of non-return valve which only allows extinguishant to the fire zone irrespective of which bottle has been fired. An arrow on the valve body allows for correct alignment when fitting.

8. SWIVEL NUT 9. PRESSURE SWITCH

Fig. 6 SINGLE HEAD EXTINGUISHER OF THE B747

blank

rnoodul11~~990

BOTTLE 2 /

BOTTLE I BOTTLE 2

BOTTLE I BOTTLE 2

- --

- - LHK 1 KIUUC

CARTRIDGE

TO ENGINE -

- / -

-- - --

Fig. 7 A TWO SHOT SYSTEM (TWO BOTTLES PER ENGINE)

MAINTENANCE

Extinguishers

Check that the bottle is of the correct type as laid down in the AMM and/or the IPC, that all markings on the container are legible and that the bottle is securely attached. Check the bottle is within life (bottle date stamped). Check for signs of leakage, corrosion, dents, scores, and damage. The state of charge should be checked by reference to the flight deck indicator (if fitted) and/or the gauge on the bottle (if fitted) and by carrying out a weight check.

The fully charged weight of the bottle is stamped on the bottle neck or on the bottle itself. On some aircraft this weight includes the blanking caps but excludes the cartridge units, on other aircraft (Boeing for example) the weight includes the cartridge units (check the AMM). Typical weight tolerances are Boeing 4%, and BAe 0.1 lbs.

EXTINGUISHER FIRING BUTTONS

DIRECTIONAL FLOW

AUTOhWTIC EXTINGUISHERS EXTINCTEURS, AUTOMATIQUES

1 ,

i 1 Fig. 8 EXTINGUISHING ~YSTEM - CONCORDE

- - / I

Check the discharge indicator. ~Fo~example, with the pin type~thzit the pin is flush with the cap. Check the threads f&,security, damagei ana 16&ing. On

1 assembly where specified use the approved grease, eg barium chromate grease. 1 ,

1 l '

I I

~etof ia tor units/ Squabs -

-- --

These are dangerous so handle with care. Never put then in your pocket. Always point the discharge end away from you and other people. Never torch the electrical connector ends - you could set it off.

Check detonator heads/cartridge units (sometimes called squibs) for corrosion, damage, and security of attachment. Check electrical cables for security of attachment and correct fitting. Check date of manufacture stamped on the unit and check the life of the unit as laid down in the AMM.

Check detonators for continuity and insulation. Remember to use safety test meters for this purpose with the detonator off the aircraft, removed from the bottle and facing away from any personnel.

Note. Early detonators were wired to the aircraft using a terminal block on the detonator head and compression fittings. Double check - particularly on the "two shot one bottle per fire zone" system that detonators are correctly wired.

The wires should have identification codes which must be correlated to the aircraft wiring diagram. If they have no codes or they have been obliterated then an electrical system function check must be carried out - with the detonator wires disconnected.

The reason for this check is to make sure that the correct pair of wires are fitted to the correct detonator - so that when the pilot selects "shot one" he gets "shot one" and not "shot two" - which would mean the extinguishant going to the wrong fire zone.

On most modem fire systems the detonators are connected by a plug and socket and the pins are so arranged that incorrect connection cannot be made.

QUESTION How would you check that the correct cartridge unit is being fired when the appropriate switch is selected in the flight deck? (5 mins)

ANSWER Consult the AMM. Disconnect all cartridge units. Connect a - -

voltmeter (set to dc and the-correct range) to the wires leading to the unit in question. ~risXire\that power is on. Operate the "fire '

one" switch for that ynit. The voltmeter should sqow 28V (or there- abouts). If it does no? check Ghich unit is beind signalled. Ensure that all wiring is correctly cqnnected.

/

Reconfigure the a i r L r @ ~ s the AMM. 1 I ',

I On iome aircraft B I T C ~ ~ C ~ S h a y be available to i-erify correct cartridge unit signallkg. I , I I

-- - L --

-

- - --

Cables

Inspect cables for chafing, moisture ingress, fraying and condition of insulation. Check for correct support and clearance. If necessary carry out continuity and insulation tests.

Pipelines

Before installation the pipes should be blown through with clean dry air or nitrogen. After installation the pipe system should be pressure tested in accordance with the AMM. Check pipes, spray rings, threads, and unions for damage, corrosion, and security of attachment.

Valves

Should be checked for corrosion, damage, security of attachment, cleanliness, freedom of moving parts, and correct location with respect to extinguishant flow. When fitting new valves always fit new seals.

Discharge Indicator Disc

Check for corrosion, damage and security of attachment. Check that the plastic disc is secure and in position and the sealing plug is in the pipe.

Inertia Switches

Always set/activate after fitting to the aircraft and before putting power on. If you put power on before pressing the SET button in then all bottles will be discharged. Refer to the AMM.

General 1 Check that all lockink is correct &d secure. 1

STOFWGE I

Extinguishers I \

Should be stored in a room, on shelves, in their packaging with blanking caps fitted. he^ should be out of directsunlight, in a corrosive free atmosphere at normal room temperature. he^ should-be inspected annbally and at the end of their life (normally 5 years) they must be returned to the manufacturers. Should be issued in strict rotation - first in first out - and records kept.

Detonators

Conditions are similar to above.

FIRE AND OVERHEAT DETECTION

The fitting of fire and overheat detectors is laid down in JAR 25 (for large aircraft). They must be fitted in Designated Fire zones of all power plants above 12,5001b.

The system should be able to:

* Show when a fire starts and when it stops. ~r Not give spurious warnings when it fails.

~r Be capable of being checked by the flight crew from the flight deck. * Provide both audible and visual warnings, with cancellation of the

audible warning with automatic reset.

There are several fire and overheat detection systems and the following pages gives a brief description of some of them. The detectors can be divided into two main groups - unit type and continuous type (fire wire).

DETECTORS

UNIT TYPE CONTINUOUS TYPE

MELTING THERMO DIFFERENTIAL LIGHT GAS PYROTECHNIC ELECT LIQUID LINK COUPLE EXPANSION DETECTOR

SWITCH

-- - --

- RESISTIVE CAPACINE RESICAP

-

TYFES OF DETECTORS

I UNIT TYPE DETECTORS 1 -- -

I - - _ . These are used in fire zones singly or more than one to give; better coverage. In some cases they are used in cohjunction with a continuous' detector.

I I ,

Melting - - Link - Switch - -. - - - - - I

Consists of a pair of contacts held apart by a mechanism controlled by a fusible plug. At a known temperature the plug will melt allowing the contacts together and completing a circuit to a warning light in the flight deck. Not very common.

Thermo-Couple Detector

A thermo-couple principle is used in jet pipe temperature measurement as well as fire detection. When two different metals are held in contact with each other and are heated they will produce a small pd (Potential Difference - a voltage that can be measured). The higher the temperature the higher the pd. When a particular pd value is reached in a fire detection circuit a sensitive relay will operate or an electronic circuit will operate to cause a visual/aural warning in the flight deck.

When the fire goes out, the pd drops and the relay/electronic circuit will cancel the warning.

rnoodull IA-~% 12 -

Differential Expansion Switch

This consists of a switch in the fire zone connected between the dc essential bus bar and the warning systems in the flight deck. It consists of two contacts held apart on a spring-bow assembly which is secured a t either end of an alloy steel barrel or expansion tube. When the unit experiences heat the barrel expands length-wise. It's coefficient of linear expansion is greater than that of the spring-bow assembly, thus the spring-bow is caused to straighten and the contacts to close - causing a flight deck warning to come on.

Fig. 9

SPRING BOW

/ / - TERMINAL

QUESTION What happens when the unit cools down? (2 mins).

ANSWER Of course the barrel will contract (more than the spring-bow), and cause the contacts to open, thus switching off the warning.

The unit is adjustable by the manufacturers only and is adjusted to operate at different temperatures. The units look very similar but their part numbers relate to their operating temperature, it is therefore most important to check that the correct unit is being fitted in the correct location by reference to it's part number and the IPC/AMM.

Besides the usually visual inspection for security, corrosion, damage etc, the unit is tested by the use of special heated tongs that are clamped to the expansion tube. These will cause the unit to expand and give a warning light in the flightdeck. When the tongs are removed the unit will reset causing the light to go out.

Light Detectors

These use photo-electric light detectors that converts visible light into an electric current. Their electronic circuitry analyses the light spectrum and will only give a warning when it detects light coming from a hydrocarbon fire. Are self resetable.

CONTINUOUS TYPE DETECTORS (FIRE WIRE)

For large fire zones a number of unit type detectors would have to be used, so it is often better to use a single continuous type.

Gas Operated Firewire

Sometimes called the Systron-Donner system it consists of a sealed firewire connected at one end to a pressure operated responder. The construction of the firewirFco_nsists of a small bore-tube Inside of which is housed-a titanium hydGde core. Outside the core, but within the firewire is helium-gas.

, ' I I I

RESPONDER

Fig. 10 GAS OPERATED FIREWIRE - GENERAL LAYOUT

HELIUM GAS (AVERAGE GAS)

TITANIUM H~DRIDE (GIVES OFF HYDROGEN GAS AVELANCHE FOR LOCAL HEATING)

Fig. 11 CROSS SECTION OF FIREWIRE

Operation

When the firewire is subjected to a n increase in temperature the pressure of the helium gas increases in proportion. This pressure is sensed by the responder and when it reaches the responder setting value the switch closes and connects a 28V dc supply to the aircraft warning circuits.

When the temperature decreases, the pressure drops and the warning circuits are de-activated. This may be used for overheat and fire detection.

When the firewire is subjected to a local high temperature, such as a small flame, the increase in pressure of the helium gas alone may be insufficient to operate the responder. In this condition the central titanium hydride core will give off a considerable amount of hydrogen gas which increases the helium gas pressure. This operates the responder.

- - OVERHEAT EXPANDS -r .. - .. lNERT.AVERAGlNG GAS

ALARM SIGNAL

POWER SUPPLY

: RE~PONDER ALARM SWITCH ,

(NORMALLY OYEN) I

ALARM SIGNAL

POWER SUPPLY

FAULT SIGNAL

POWER SUPPLY

- AVERAGE OVERHEAT - 1 I I

FIRE RELEASES ACTIVE GAS FROM yYDRlDE CORE

-

.-

f

RESPONDER ALARM SWITCH (NORMALLY OPEN)

LOCAL OVERHEAT

INTEGRITY SWITCH (HELD CLOSED BY NORMAL PRESSURE)

FAULT STATE

Fig. 12 OPERATION O F GAS OPERATED FIREWIRE

As soon as the temperature drops below a certain value the titanium hydride core re-absorbs the hydrogen gas, the pressure drops in the helium gas and the warning signal is cancelled. This is used as the fire detection mode.

The above cycle is repeatable indefinitely.

The fire wire is pre-pressurised so that a pressure integrity switch will operate should the firewire develop a leak. This is incorporated in the responder and linked in with the element self monitoring circuits.

The fire wire may be a twin fire wire with the two wires running in parallel in rubber lined clips.

CONTROLS AND INDICATORS: CONTROLS AND INDICATORS ON OVERHEAD PANEL:

- 1 - ---

L

I 1

AGENT I ?H-J-,r 5 El

I R

- -

1 f -- - ENG, 1 FIRE

' TEST 8 i

I -

I

qUsH

I

Fig. 13 EXAMPLE OF A FLIGHT DECK FIRE PANEL - A 3 2 0

Example - A320 System (figure 13)

Two control panels are provided - one for each engine. The fire push button (1) is guarded and released out. When this happens the following occurs for this engine:

-k Aural warning is cancelled. * Squib (cartridge) is armed. A The following valves close - LP fuel - hydraulic - air bleed -

pack valve. * Generator deactivated.

After operation of button (1) the Squib button (2) is pressed to cause the fire bottle to discharge. The squib button lights when button (1) is pressed to indicate which squib button to push. The DISCH area lights amber when the bottle is discharged.

The test button (3) tests the system. When pressed:

* Warning chimes sound and master warn lamp flashes. * ECAM engine fire warning is activated. * Engine fire button (1) lights red. * Squib lights illuminate white. * DISCH lights illuminate amber. * On engine centre pedestal panel fire light illuminate.

Pyrotechnic Fire Wire

Not used on civil aircraft. A pyrotechnic cord is housed within a small bore steel tube (similar in size to ordinary fire wire). When an over-heat situation arises the cord ignites, creating pressure and operating a pressure switch within the control module.

Electrical Firewire - -

I

These form a continuous stainless steel (or'inconel) loop a k u h d the fire zone with both ends of this small di-eter tube entering a control box. , I I

, i I I

Modem technology exists for the single loop to pass around several'fire zones with the electronic control box ablFto detect where, in the loop, theloverheat has occurred. Most aircraft ha?e one loop - usually duplicated - for each fire

I zone. , l 1

The element has one or two central electrodes which are inkulated from the outer tube by a temperature sensitive material. In some aiicraft the outer tube element is supported on special metal supports and may be protected within perforated stainless steel tubes. It is more usually supported in rubber covered "P" clips to prevent chafing.

Fig 14 CROSS-SECTION OF FIREWIRE

blank

There are basically three types of electrical fire wire - resistive type, capacitive type and resistive/capacitor type.

(a) Resistance type. With this type the central electrode has a positive voltage applied to it and the outer case is grounded. As the temperature rises the insulation breaks down and current will start to flow from the central electrode to the outside of the fire wire - to earth.

This current will cause a relay to operate in the control box and cause warnings to come on in the flight deck.

When the fire is out and the temperature drops the insulation will regain its former properties, the current will cease to flow, the relay will open and the warnings will be cancelled.

When the 'press to test' switch is pressed in the cockpit a relay is operated within the detector unit which sends a dc current through the complete circuit. This will only happen if the circuit-is complete, and all electrical supplies are connected and on (usually botKadc and an ac supply is required). I

Fig. 15 RESISTANCE TYPE FIREWIRE DETECT1 ON CIRCUIT

The ac supply provides power for the fire wire, the dc supply for the relay, and the dc test for the test circuit. When the "press to test" button is operated (from the flight deck) a current is passed through the control box and the complete length of the central electrode of the fire wire to operate the warning lamp in the flight deck.

Should the electrode be broken or the control box not work correctly or any supply be missing then the warning lamp will not operate. With a break in the fire wire, fire detection is not affected but "press to test" is.

(b) Capacitance type. Capacitance is the ability to store charge, and the central electrode and shell of the fire wire form a capacitor with the insulation or dielectric between. The capacitance of a capacitor depends on several things and one is the value of the dielectric (dielectric constant). This changes with temperature. As temperature rises so does the capacitance of the fire wire.

The core is supplied with half wave ac current which the wire stores during the first half of the cycle and returns to the control box during the second half cycle.

With an increase in temperature the returned current becomes greater, and at a pre-set value operates a relay to trigger the warning systems in the flight deck.

A "press to test" facility is provided. If the wiring or fire wire are shorted to earth then a false warning does not occur.

!c) Resistive/capacitive type. With this type nf fir^ wire the impeds=ce as welldas the resistance ismonitored. 'with an increase i n temperature the resistance drops and the impedance becomes more rqactive. When this happens the detector unit registers this as a fire. A pvrk resistance will be registered as a fault.

I

/ I 1 --

I I - , With rhost aircraft the signals from-the above systems go t o a card within a rack mounted unit. This will prockss the signal and send w v i n g s signals etc to all other appropriate systems/ units. 1 I

1 I

I

--

- - - - - - - - - FIREWIRE ELEMENTS ELEMENTS FIREWIRE

CONNECTING CABLE CABLE DE CONNEXION

Fig. 16 FIRE DETECTION SYSTEM - CONCORDE

rnoodutl lkBg

Liquid Type Detector

A small bore stainless steel pipe is filled with liquid that expands on heating - the pressure causing a pressure operated switch to close an electrical contact. When the heat is removed the liquid contracts and the switch re-sets. Rare.

SMOKE DETECTORS

These are fitted to:

* Baggage holds (type B, C & E - others optional). * Freight bays. * Toilets (each one). * Equipment bays (E & E bays for example).

On some aircraft the smoke detector system might incorporate an air duct system with fans that draw air from various parts of the aircraft through a single-detector unit. With this system-there may be provision to "select" the area that-& is being drawn from andhence the ability to determine the action to be taken in the event of a s4oke warning. l 1

I I

Most detectors have their own electrically driven fan to draw in the air 'and this may come via a small plenum chamber. .,

- ,

Smoke detectors ,may be of the Ionisation type or Photoelectri~ type. I 1 1 I

REFLECTED ELECTRONIC PILOT LAMF LIGHT CIRCUIT & RELAY

.. . \ \- / / CELL

OUT

TEST LIGHT BEAM

LAMP TRAP

BLOWERS I/

Fig. 17 PHOTOELECTRIC SMOKE DETECTOR (LIGHT SCATTER TYPE)

- 20 - moodull 1.A-933

Photoelectric Smoke Detectors

These may work on the principal of Light Absorption or Light Scatter

Light Scatter Type. (Figure 17) These work on the principle that when light falls onto a photo-electric cell it will produce an electric current.

A bearn of light is caused to shine within a light proof container. The beam is a focused beam which shines into a light trap (painted mat black to absorb the light).

When clear air is being drawn through the unit no light is scattered and no light reaches the photoelectric cell. When smoke enters the unit it causes the light rays to reflect off the smoke particles and scatter in all directions.

The scattered light rays fall onto the photoelectric cell producing electricity (in basic terms, one photon of light will produce one electron of current). The

\ output is amplified, put through a relay, which causes a warning to come on in ~ 1 - - n:-i_L _1--1- - - - - - I - - -1 --: - - - - 1 ~ l l e III~IIL UCCK - a u l a iulu v l s u a . I

- - -

- - -

When the smoke ceases the reverse happens and the warninis cease. 1 ,

In some units two cells are used with on& being subject to smoke (if it is present)and the other not. Both have the same light sourcd apd ark connected by a Wheatstone bridge circuit. ~ k e n 3 m o k e is present one, cell produces a current which unbalances, an btherwise balanced bridge. Thq bridge output will operate a relay to cause a &-ing to come on in the flikht deck.

I

Of course, when there is no smoke-present everything reveqses and the flight deck warning goes out.

QUESTION All of these units are fitted with a "press to test" facility. In figure 17, the "press to test" button sends a 28V dc supply to the test lamp. When it's light falls on the photo-electric cell a warning is given. However, the pilot lamp may not be working so the test would be invalid. How could the test be made to be valid? (5 mins).

ANSWER If the test lamp is wired in series with the pilot lamp, the test lamp will not work unless the pilot light is also working.

IMPORTANT. If the pilot lamp fails it is important that it is replaced with the correct lamp. The wrong larnp may cause the sensitivity of the unit to drop, reducing it's effectiveness.

Light Absorption Tvpe. With this type of unit a photoelectric cell is placed at a distance from a light source (within the unit) so that the cell receives 100% of the light during "no smoke" conditions. When smoke is present the amount of light reaching the cell reduces and it's electrical output drops.

This reduced output is used to trigger a smoke warning on the flight deck.

When the smoke clears the unit reverts to it's 100% light condition and the warning is cancelled.

Alpha Particle Detector

This is a dc operated unit containing a double balanced, ionisation chamber. One part of the balanced chamber is protected while the other is open to atmosphere. When smoke is present an unbalance occurs because the alpha particles are absorbed which causes a decrease in the current in the ionisation chamber. This initiates a discharge of a cold cathode tube which activates a

-

~ ~ - ! - i n ~ When the smoke cle-qs the whole system will reset. - --

l 1

I 1 I EXAMPLE - A 3 2 0 CARGO HOLD SYSTEM

I 7--

L

The f o y a r d and aft holds are dasically deiigned as class 13' compartments requiring neithersmoke detectjon nor fire. extinguishing systems.

I I t I 1 1

However, if the optional cargo Gold ventilation is installed, smoke detector system must be installed. - ' - ,

- - --

This system is if the dual loop type. Two detectors of the ionisation type (one per loop) are located in the forward cargo hold, two in the parallel section of the aft hold and two in the non-parallel section (six detectors in total).

In each compartment, signalling from two detectors (loops) is required to produce a smoke warning (via the AND gate), except in the case of failure of one of the detectors in which case the respective loop is automatically deactivated.

A smoke warning will automatically shut-down the (optional) ventilation system.

A cargo smoke detection unit (CSDU) is installed to control signalling and monitor the system. The CSDU provides signals to the CFDS, FWS, ECAM and signals to illuminate appropriate smoke warnings in the flight deck.

Testing of the smoke detectors can also be done by a push button on the overhead panel.

--

Fig. 18 A320 CARGO BAY DETECTION & EXTINGUISHING SYSTEMS

i 1

In addition an optional cargo hold fire extin&ishing system is available. This system consists of a fire-extinguishing cylinder for both holds. The extinguishing agent is halon 130 1. The cylinder is equipped with two

--

independent discharge outlets.

The discharge cartridges are electrically ignited by guarded switches located on the flight-deck. Fire extinguishing agent may be supplied to either the forward or aft cargo holds.

The continuity of the wiring to the discharge cartridge including cartridges is permanently monitored by the CSDU. Once the cylinder content is discharged a pressure switch signals lamps in the cockpit to confirm that the agent has been discharged.

WHEEL WELL FIRE DETECTION SYSTEM

Fires can be caused by heat generated from the brakes. Provision exists on most later types of aircraft to monitor the temperature in the wheel well in terms of overheat conditions when the landing gear is up.

Temperature sensors are also fitted to the brakes to give flight-deck indication of actual brake temperatures at all times. (see the book in this series on Wheels, Tyres and Erdces).

The following description is based on a Boeing aircraft. There is no need to commit the details to memory, but you should know the principles and its operation.

Wheel Well System (Figure 19)

There is a single loop detector located in the each main wheel well. The detectors are made by Fenwall, and are sensitive to temperature (resistive type firewire). These detector elements operate by the reaction of their internal salts to temperature change - the resistance decreases with an increase in temperature. At a temperature greater than 400°F (204°C) the wire provides a ground warning signal to the control card.

The resistivity of the salt drop at this temperature and allows current to flow from the source element (in the centre of the firewire) to the grounded outer sheath. The elements automatically reset on temperature reduction.

Fig. 19 WHEEL WELL OVERHEAT DETECTION

The wiring is in two sections and connected in series to form each detector. The two detectors are in turn connected in series and wired to the duct leak and wheel well Fire Control Card.

The system will give a warning on EICAS (Engine Indicating and Crew Alerting System) and turn on the wheel well fire light. In addition, the flight compartment master warning lights illuminate, the discrete fire warning illuminates and the flight compartment fire audio sounds.

Test

A test switch is located on the fire test panel. The switch is labelled "WHL WELL" and when activated sends a signal to the control card. If the control card is operating normally, and the two detectors are continuous, the wheel well fire light illuminates along with the master warning lights, discrete fire light, and fire audio.

PNEUMATIC DUCT LEAK DETECTION SYSTEM

~ n e u m a t i c duct leak detection systems are designed to give the crew notification of a ruptured pneumatic duct. Duct air temperatures can be as high as 200°C and any leaks are a fire hazard. I

I

Sensors are fitted to ducts taking air from the jet engine for pressurisation purposes, de-icing, heating etc. 1

1 I

1

The following system is based on a Boeing aircraft. Do not try, to remember the detail but you should understand it and relate what you read to what is fitted to you'r +-craft. I / 1

-- A -

The system-iscomposed of dual loop ~ e n w a l detection elements wired with 'AND' logic. These loops are set to trigger an alarm at approximately 225°F (124°C).

The system is divided into left and right zones. The left zone has five elements wired in series: the left wing leading edge, the left air conditioning bay, the wheel well, the aft cargo compartment, and the pressure bulkhead aft to the APU.

The right system has two elements wired in series; the right wing leading edge and the right air conditioning pack bay.

Annunciation and Test

The two systems are wired to the duct leak and wheel well fire Electrical Systems Card. If the system senses an overheat, it is enunciated as a caution through the warning electronics unit. It also annunciates the appropriate duct leak caution light on the pneumatic control panel and a level "B" caution appears on EICAS.

To test the system for correct operation, a test switch is located in the flight deck. The switch is labelled 'DUCT LEAK'. A successful test illuminates the two duct leak lights.

PNEUMATIC CONTROL PANEL

------

RIGHT SYSTEM DETECTION

------

L (R) EICAS , COMPUTERS

I

EICAS DISPLAY)

I I ' ,

Fig. 20 BOEING PNEU~ATIC DUCT LEAK DETECTION SYSTEM - i- - ,

TOILET TOWEL BIN FIRE DETECTION

This fire extinguisher is a self-contained detection and extinguishing system located in each towel bin. The bottle is positioned above the paper towel bin, to the side of the slot used for discarding the paper towels. The two nozzles point through the bulkhead directly down into the removable waste container.

It is fitted to detect fires caused by people putting their cigarettes out in the waste paper bin. Smoking is not allowed in the toilets - nor incidentally when walking in the aisles - so it should not happen. Unfortunately it does, but not often.

The bottle has two nozzles. These have a heat sensitive solder covering their discharge ports. The bottle contains HALON 130 1 and weighs 2 pounds (0.9kg) when full. When either nozzle temperature reaches 175°F (80°C)' the solder on the discharge nozzle melts. The extinguisher agent is directed into the waste container. This operation is automatic and has no controls or indications on the flightdeck.

U(TINGUISHER INSTALLATION

Fig. 2 1 PAPER TOWEL-BIN FIRE BOTTLE l 1

Temperature Strip I

The dots on thisfstrip are progrks,sively sensitive to heat. The heat will turn the dots black in sequence. The strip is located /above the chute, behind the towel slot.,After a fire, this strip must be repl5ced. Fires that exceed a critical temperature will require a structural inspection of the chute compartment and immediate area.

TOILET CUBICAL

The requirements of JAR 25 (for large aircraft) state that a smoke detector is fitted in the toilet cubical to give a warning to the flight crew should anyone smoke in the toilet cubical. The smoke detection system senses the presence of smoke in the toilet cubical and activates an aural/visual alarm. The detector is located in the ceiling.

The assembly consists of a sensor unit, alarm, horn and electronic circuitry. The sensor incorporates a red alarm indicator to provide a visual indication when smoke is detected. A green power indicator is provided. A powerlreset button is fitted near the power indicator. The power supply is 28 volts dc.

blank

- 27 - rnoodul I I -940

POWER POWENRESET INDICATOR SWITCH (GREEN) \ /

J

DETECTOR FACE

FLUSH MOUNTING CONNECTOR BRACKET LAVATORY

SMOKE DETECTOR

, Fig. 22 TOILET CUBICAL SMOKE DETECTOR - -- -

I ' I

I

i

LAVATORY CALL LIGHTIRESET SWITCH

CABIN SYSTEM CONTROL PANEL

-. CABIN AREA CABIN MANAGEMENT CONTROL PANEL SYSTEM r CHIME - I A V MASTER CALL

LIGHT

LAVATORY SMOKE DETECTOR I

Fig. 23 BOEING 777 TOILET CUBICAL SMOKE DETECTOR SYSTEM

AIMS CABINETS (2)

blank

-- ElCAS MESSAGE

MAINTENANCE - GENERAL

In general d l components must:

* Have blanking caps fitted whilst in transit (also transit safety pins where necessary).

* Have correct washers and sealing glands fitted - new when replaced. * Be secured, locked, damage and corrosion free. * Have all connections correctly torque loaded. * Be correctly located - not too near heat shields or other hot

surfaces - not too near moving parts. * Be clean and free from oil, dirt, etc.

Wiring should be checked for:

* Correct support and routing. * Good insulation and continuity. * Security of attachment, plugs, etc.

- -

Control Units and Detector Heads \

I 1 These may have to be removed irorn the aircraft from time tb time for bench checks to be carried out. This will involve the use of special test equipment. These tests will normally be cmriexout when specified in the AMM-and as an acceptance check on a new component rece'ived from stores.

, , Fire Wire-- - - . ' - ~ - ., ,

i , -~

._,' ~ -~ ~- .- - - - ~ , -

Some fire wire is fitted within an armoured sheath to protect it from damage but where the fire wire is not so enclosed it must be correctly supported in rubber supports (to prevent chaffing) within special clips ("P" clips).

It is most important when fitting new lengths of fire wire to check that the part number is correct - as many fire wires are very similar and the fitting of an incorrect fire wire could have serious consequences.

Electrical insulation/continuity and other special checks should be carried out on the fire wire prior to fitment as per the AMM.

The attachment to the structure should be in accordance with the AMM - usually in rubber grommets to prevent chafing by vibration, but in general the fire wire should be supported:

* at 6" intervals along the length. * 4" from the end fittings.

Bend radii of the fire wire should be kept as large as possible.

System Testing

All systems have a "press to test" or BIT facility. There may be one test button to test the complete system or there be more than one, eg, one for the firewire, one for the cartridge unit and one for the rest of the system.

It is important to note that for a "press to test" to show a satisfactory result all the system that it is testing is working properly - for example, all power supplies are available - to lamp and control modules (possibly ac & dc); all appropriate control modules are working; continuity of firewires is satisfactory and bonding connections are serviceable.

O n the electrical fire wire the "press to test" button grounds the resistive type central electrode and causes current to flow operating a relay. O n the capacitive type the "press to test" button increases the capacitance of the system causing it to operate and give a warning.

On some aircraft the fire detection systems are monitored by on-board fault computer -- systems which can be accessed on the flight deck. -

1 -

I

YOUR ACTIVITY I '

I

If you are currently working on aircraft have a look a t your aircraft aqd answer -- - --

the following: -, \

'> I I

\ I

1. Are there fire detecfiqn/smoke detection systemslfitted to all those areas mentioned in this book? If not why not?

I

2. Any that are fitted, are they similar to those described in this book?

3 . If they are different can you describe their operation and maintenance?

4. Are there any fire extinguishing systems fitted? If so how do they work and what extinguishant is used?

CONTENTS

Page

Powered flying control units 1 Valve ram type PFCUs 4 Power assisted controls 5 Power operated controls 5 Manual reversion 7

BAC 1 - 1 1 PFCU 8 Self contained PFCUs 11 Artificid feel 13

Spring feel 13 Q feel 14

'Trimrmjn~~ - -.- -- , 16 F ' - --

The autopilot system 17 -

Servos 24 1~

Maintenance I ?7

Airbus aileron system I 28 - - -

1 I

Airbus elevator system ' 1' 30 I -

Ruddqr system \ 30 1

1 32 1

B767 system I I I I

B767 yaw damper servo , 35 I )

Flying c&ntrol centralised warning system 37 I Fly-by-wire -systems - - -/ 39 -

Airbus A330 system -- - - 40 --

Boeing B777 system 47 Fly-by-light systems 52 Mach trim 55 Stall warninglangle of attack indication 57

Airflow operated 58 Leading edge stall warning vane 58 Rotating A of A probe 58 Trailing A of A vane 60

rnoodull I A-944

POWER FLYING CONTROL UNITS (PFCUs)

QUESTION Why are PFCUs (PFUs) fitted to aircraft? (3 mins)

ANSWER Because of the speed of the aircraft and/or the size of the control surface the aerodynamic loads imposed on the control surfaces are too great for the pilot to overcome manually. So the controls are powered on lagelfast aircraft - usually hydraulically.

The control surfaces of large/modern high performance aircraft are subject to

high aerodynamic loads. These loads are related to the formula i p ? s , where 2

p = density (1.2 kg/m3 a t sea level) v = velocity of aircraft s = is related to control surface size/angle of deflection

- -. -- - -

As aircrait have got larger and/or faster,so,v and/or s in the tormula have got greater with a resultant increase in aerodynamic loading. ' I These'lohds are likely to be in ~ x c e s s of thoke which could be ~ornfoytably accommodated by the pilot. To assist-the pilot in overcoming these loads some

L.- 1 form of powered operation of the Fontrol&rface is essential. W ~ t h fully power- operated controls the normal forms of aerpdynamic assis take, ie aerodynamic balance, spring balance etc are not effectike.,

1 I 1

herd a control system is fully,pbwered thd pilot has no feei of the aerodynamic loads on the control surface:'~herefore he/she c s n o t instinctively position the control surface in relation to the speed and altitude of the aircraft. If through the lack of such control, rapid movement is applied to a control surface, the surface may be damaged, the airframe may be over- stressed or the aircraft may become unmanageable. The pilot must therefore be given some indication of the aerodynamic forces on the control surface by an artificial feel system incorporated into the control system.

If the system is a fly-by-wire computer controlled system such as on the A320 then the pilot has no feel at all - either artificial or otherwise. In this case when the pilot makes a selection the computers judge whether the selection can be made by that control surface - the speed of movement and the range. If the pilot over-selects - which would normally cause the aircraft to go into a violent manoeuvre or cause damage to the airframe/control surface - then the computer gives the pilot the maximum selection of the control surface under those circumstances. Thus the aircraft/control surface will not be over stressed.

On aircraft such as the B777 (which is fly-by-wire) the flight deck controls are of the conventional type and provided with artificial feel. On the A320 the control column is replaced by a side-stick.

There are two types of PFCUs:

(a) Valve Ram Type. This consists of a servo valve and ram assembly supplied with hydraulic power from the aircraft's hydraulic system.

(b) Self Contained Wpe. This consists of a self contained unit with its own hydraulic system with an electrical supply from the aircraft's electrical system to operate the hydraulic pumps. Also contains servo valves and a ram assembly.

General Principle

The pilot moves the controls and that movement is transmitted to the PFCU b 2 111eckl~6i~d system (pi--ish/p~=~l_l Fcl_s, ca-hl~s etr) to n p ~ r a t e the sprvo valve; or a fly-by-wire system (an electrical sign& to an electrically obeiated servo valve), or a fly-by-light system (the signdl transmitted by light to be c,onverted:to an electrical signal to operate an electrically operated servo valye).

1 I I

When ;the pilot moves the controlkihYdr<ulic servo valve i$ selected in the PFCU add the PFCU output jack \ram) 'st&ts to move. As soon as the output jack staits to move its movementis felt bick at the selectorservo valve - and it tries to de-selecf {he servo valve. This is called "feedback.

, 1 - I -

QUESTION Thei-e are two types-of feedback - negative and positive. Do you happen to know what sort this is? (1 minute)

ANSWER This is negative feedback - as indeed it is in almost all control systems. The negative feedback is fed back from the output (the jack) to cancel the input (the pilot's input to the servo valve).

SERVO VALVE CONTROL SYSTEM eg PFCU

P l L o r s INPUT SIGNAL ACTUATOR1 JACK

SUMMATION DEVICE . NEGATIVE FEEDBACK SIGNAL

Fig. 1 NEGATIVE FEEDBACK

The fundamental requirements of a powered control system are laid down in J A R 2 5 and include:

(a) Sensitivity - The system must respond as rapidly as possible to input signals initiated by the pilot. To ensure this the servo valve must be sensitive to small input movements.

(b) Stability - The system must remain stable and uninfluenced by any signals which do not originate from the pilot. To maintain stability it is essential to ensure that the linkage is free from backlash, that the dampers function correctly and that the hydraulic system is free from air.

(c) R i ~ d i t ~ - Flexibility in links and anchorage's must be eliminated.

(d) Irreversibility. - Once a control surface has been selected to take u p a particular position it must not be deflected by the airflow.

(' (e) Pilot's feel - In power operated systems feel must be provided by artificial feel units. On power assisted-systeqs the arrangement I-, of-the linkage provides feel (proportion& feedback). \-

I I

(f) Emergency Measures - There must be provision for alt$Aative methods of control so that in the evenf df a rnephdnical failure of the PFCUlthe pilot will! still be able to control 'the-&-craft. There can includk: ,

/ 1 I

\

(1) ' Manual Reversion. Where the pidots input lever is connected mechanicall$ to the PFCU output lever) in the event of

I I , I hydraulic failure. , I ' I

' I

-- /, / _,

(2) Duplication of ~ o m L o n e n t s ~ - ~ l i & - e there are more th& one PFCUs operating each control surface, each with its own hydraulic supply. Interconnected via spring rods.

(3) Split Surfaces. Where each control surface is divided into two or more portions. Each having it's own PFCU. Therefore in the event of failure of one unit the other portion or portions of the control surface will continue to be operated by their own particular unit and the portion which is operated by the failed unit will be isolated. Failure of a PFCU is indicated to the pilot by warning lights or magnetic indicators. A failed PFCU should go automatically into the trail position during flight.

(4) Duplication of Hydraulic Supplies. Most large aircraft have three independent hydraulic systems all connected to the PFCUs.

(5) Duplicate control runs with a disconnect mechanism should one become jammed.

(6) Fly-by-Wire Systems. For 'Fly-by-Wire' aircraft with computer controlled PFCUs there is duplication or triplication of the PFCUs and of the computers. It is also usual to have each computer working with software that is supplied by a different software house than that used by the other computer. Manual reversion is also available - albeit in a limited form - should all the electronics 'go down'. Some data buses may also be duplicated.

VALVE RAM TYPE PFCUs (figures 2 to 9)

With reference to figure 2. The unit comprises a jack body connected to the control surface, an equal area jack ram connected to the structure, and an integral servo valve connected to the pilot's control. When the servo valve is neutral the delivery ports are closed and the fluid is trapped in the jack body so that the jack is hydraulically locked and there will be no movement. When an input signal is applied the servo valve moves to open one port to pressure and

7

t h e Qther, te ret~.!.m, n d the jzrk ~ Q & T J_---- mnves - - alnng the ram i i n r l ~ ~ hydra-cilic pressure deflecting the control surface. A s soon as the input ceases, ie the pilot's control is held station& in it's new position, the jack body moves relative to the servo valve slide until the slide is in the neutral1 position, ie closing the delivery and return borts and ensuring a hydradlic lock within the jack. The control surface is nod in the coAect position reladive to the pilot's control. 1 rppp- 1

RETURN 1 PRESSURE /

~- ~ SERVOVALVE - - - -

- PILOTS INPUT

OUTPUT TO CONTROL SURFACE -

S TRUCTURE

JACK RAM - Fig. 2 VALVE RAM TYPE PFCU

This 'follow up' of the jack body is the negative feed back of the system. So long as the pilot keeps the input going so the selection will be maintained and the control surface will continue to move. But as soon as he/she stops the jack body will catch the pilot's input servo spindle up and cancel the selection. This amount of movement is very small and takes a fraction of a second.

The same PFCU can be mounted in two ways so as to give either fully powered operation or power-assisted operation.

Power Assisted Control (figure 3)

In this type of control the input link and the output link are connected in such a way that some of the loading felt by the control surface is felt by the pilot. Thus the pilot has some feel.

When the pilot's control is moved it causes selection of the servo valve which causes the PFCU to move. The loading felt by the control surface is also felt back through the jack ram onto the pilot's input lever. The pilot will feel a proportion of this loading this in the ratio a:b. To enable the system to work there is usually a lost-motion bush (a bush with a small amount of play in it) a t the connection of the jack ram to the pilot's input lever.

-

TO CONTROL SURFACE

, Fig. 3 POWER ASS1;STED PFCU 1 i

l 1 ,

/ /

A s the --- control - - surface is linked to the-control column in su'ch =way that part of the control surface load is imposed-on the control columnpwhen power is 'ON' proportional feed back provides feel to the pilot.

Power Operated Controls (figures 2 & 4 to 11)

With controls fully power operated the whole of the force needed to operate the control surface is provided by the power system (hydraulic). Movement of the pilot's control column moves the control rod of the servo valve. The servo valve then allows fluid under pressure to operate the hydraulic jack and so move the control surface. As all the effort to move the control surface is supplied by the hydraulic jack, the pilot has no feel of the loads on the control surface. Feel therefore must be provided artificially, eg a spring or Q feel unit.

The systems shown in figures 2, 3, 4, 6, 7 & 8 use a PFCU where the body of the unit moves the control surface while the jack ram is attached to the structure. With some PFCUs this is the other way around (figures 5, 9 and 10).

When the jack ram moves the PFCU there is a "summing link" between the pilot's input and the PFCU output which is in fact, the negative feed back link.

CONTROL COLUMN

fl HYDRAULIC RETURN PRESSURE

PIVOT 1 PFCU

. VOT

ELEVATOR

U U - FLYING CONTROL SYSTEM - Fig. 4 FULLY POWERED PFCU

- SERVO VALVE RETURN -LSUPPLY INPUT LINK PILOT INPUT

I C 9 C----L

S~RUCTURE /SUMMING 1 1 LINK

B

, , I

1 b - -

I \ A 1 PFCU OUTPUT

PFCU I TO CONTROLS I

-

--- - Fig. 5 P ~ C U WITH SUMMING LINK'

With reference to figure 5. When the pilot puts an input into the PFCU the summing link will initially pivot above point A on the ram. This will cause the servo valve to select to move the ram. The movement of the ram will be in suck, a direction as to try to de-select the servo-valve, so as soon as the pilot stops his/her input the jack ram will de-select the servo valve via the summing link pivoting about point C. Example:

Pilot inputs to the left. Summing link rotates about A in an anticlockwise direction. Servo valve selects. Jack ram moves to the right. Pilot ceases input. For a fraction of a second jack ram continues to move. With pilot holding the control stationary the summing link rotates anti-clockwise about point C. This movement will de-select the servo valve setting it into the neutral position and holding the PFCU in a hydraulic lock in its new position.


Note. The simple summing link shown here is normally a set of links so arranged as to allow for the incompatibility of the range of movements between the jack room which can be, say 8 to 12 inches (50mm to 305mm), and the servo valve which is in the range of 0.030" (0.7mm).

Manual Reversion

Some PFCUs have a manual reversion facility. This allows the pilot to operate the controls manually should the PFCU fail. Manual operation will be heavy with reduced control authority but it is a reliable emergency standby measure.

PILOT'S INPU

HYDRAULIC FAILURE (PILOT'S INPUT LEVER LOCKED, FLUID BYPASS VALVE OPEN)

-

I - ,- L

I I

' 1

Fig. 6 MANUAL REVERSION TYPE PFCU

When manual reversion occurs the pilot will move the control surface directly via the PFCU. In this case the unit just acts as another link in the system. The PFCU goes into manual mode by:

(a) Disconnecting the jack ram from the structure. The jack ram is connected to the structure by a hydraulically operated lock mechanism which disengages automatically to allow the jack ram to slide freely back and forth. The pilot's input, via the servo valve, moves the complete PFCU including the ram to move the control surface direct.

(b) Allowing fluid to transfer freely from one side of the jack ram piston to the other by a hydraulically operated transfer valve. Both sides of the jack ram piston are connected together hydraulically by a valve which is normally closed. When normal supply pressure fails the valve opens and allows free movement of hydraulic fluid from one side of the piston to the other (figure 6). The pilot's input causes the servo valve to move the jack body direct and it slides u p and down the jack ram.

In both cases the pilot's input is via the servo valve input and the PFCU moves in response to the pilot's force thus moving the control surface. In other words the PFCU acts as a control link between the control system and the control surface, and has no other function.

THE BAC 1 - 1 1 UNIT

The following is a brief description of the PFCU as fitted to the BAC 1 - 1 1. There is no need to commit the details to memory but you should be able to understand the operation and-relate this to what has alregdy been said.

I \ I

The uAit shown in figure 7 is ohe of a pair khich together operate the rudder. Each unit is fed by a separate dydraulic supply from the aircraft.

! ' / I

Individual flying control input <ohne~tion$ via spring strut4 pbrmit bperation of each unit even though the other has failed and each unit isi capable of operating the rudder on its own. /

I

Operation of Rudder Unit

Pressure Off. The by-pass valve provides a fluid way to both sides of the jack ram via restrictors so the PFCU body will move, operating as a virtual control link with gust dampers. To prevent cavitation under such conditions, the suction NRV (Non Return Valve or Check Valve) opens to permit fluid flow frc the return line.

Pressure On. Pressure action on the by-pass valve closes it against its spring, shutting off the by-pass flow. The supply NRV then opens and delivers fluid to the rotary control valve moved by pilot input.

Input selection, as shown, opens up one side of the jack to pressure and the other to return and the jack body moves in the direction shown.

Once input ceases, the body will catch-up with the input, returning the rotary valve to the neutral 'off position and a hydraulic lock occurs providing irreversibility.

Cross leak restriction provides stability within the unit.

r - ---- PFCll WITH NO HYDRAULIC SUPPLY PRESSURE

! -\

PFCU WITH PILOT'S INPUT AND HYDRAULIC PRESSURE SUPPLY

Fig. 7 EXAMPLE - BAC 1- 11 RUDDER PFCU

Emergency

Hydraulic System Failure. The failed unit reverts to by-pass condition and is driven by the remaining operative unit.

-. g -

rnoodull 1 A-953

Stuck Valve. On pilot input the spring st,rut will collapse in the appropriate direction permitting full travel of the remaining PFCU rotary control valve.

At the same time the micro switch is activated by the spring and roller bearing in the groove. This switches on a cockpit warning light and closes a pressure shut-off valve in the supply to the failed unit, which then reverts to the by-pass condition (as in hydraulic system failure).

Operation of Elevator Unit

The unit shown in figure 8 is one of the pair of units which operate the elevators. It is virtually two rudder units in a common case with two rotary valves and two input spring struts. Its operation is basically similar to the rudder units described in both normal and emergency.

However, the left hand by-pass valve has three positions:

11 Pressure on - fully closed. \

2. Pressure off on g o u n h - fully open. 3. Pressure off failure - &ual reversion.

1 I 1

%STEM I [' m 2

SUPPLY RENRN -- \ SUPPLY RENRN - .

SUPPLYPRESSURE O F F - m u s l - - BYPASS W U I I L ClRCUrr OPEWnDNl - -

Fig. 8 EXAMPLE - BAC 1 - 1 1 ELEVATOR PFCU

Pressure Off - On Ground. Both by-pass valves to by-pass condition - but left hand flow is via restrictor valve which acts as gust damper.

Two System Failure in Flight. Controls revert to manual. Left hand by-pass valve held in intermediate position by undercarriage controlled solenoid, provides unrestricted by-pass as does right hand one. Input signals continue until the input connection abuts the input stops and so connects input to output.

Most modern PFCUs are similar in principle with the majority of servo valves being of the slide type. With a Fly By Wire PFCU (A330) as shown in figure 9 the slide valve is moved by a solenoid operated hydraulic jet type valve (servo valve). The solenoid gets its signals from a flying control computer. Note also that the PFCU has a futed body with the ram moving the control surface and feedback being provided by the feedback linkage from the ram to the servo valve.

--- 7 - I SUPPLY RETURN i

\ A

I

LECTOR

ATTACHMENT TO FEED BACK LINKAGE

RAM THE STRUCTURE

Fig. 9 PFCU - A330

CONTROL SURFACE

SELF CONTAINED TYPE PFCU

These are less common than the valve ram type. They are self contained in that they have no external hydraulic power supplies. Internally they have their own complete hydraulic system to include: pumps; valves; reservoirs; pipelines; jacks, etc built into the one case. The only external connections are electrical and the control rods (input and output).

An electric motor continuously drives a bank of hydraulic pumps.

Movement of the pilot's input causes a servo valve to move the main bank of pumps, which causes fluid to be pumped to one side of the jack. The other side of the jack is connected to suction. Movement of the jack moves the control surface and a feedback link mechanism.

SPRING BOX

INPUT SPRING

- CONTROL SURFACE - - \ 1

I I \ Fig. 10 SELF I CONTAIN,ED TYPE PFCU ~

SERVO 8t MAIN PUMPS

I

PILOT'S INPUT JACK / - TO CONTROL SURFACE

Fig. 11 SELF CONTAINED PFCU - GENERAL VIEW

With reference to figure 10. All the components except for most of the control links and the control surface are all housed in the one unit - which is shown in figure 11. The servo pump provides pressure for the main servo valve which is moved by the pilot. Movement of the valve will cause the main pump to pivot up or down providing pressure to the left or right hand side of the jack. The jack will move left or right with feedback linkage providing negative feedback to the control input. Servo and main pumps are on a common drive shaft continuously driven by the electric motor.

Figure 11 shows a general view of the PFCU as fitted to the VC 10. Note the attachment of the body to the structure; the pilot's input link; the jack ram connection to the control surface and the feedback link to cancel the pilot's input.

ARTIFICIAL FEEL

/' Since the control load is now taken from the pilot and all he/she has to do is move-the yalves of the various PFCUs; there is no "feel" of;the loads on the control surfaces, and it is thereforeplikely that the pilot will'ovler-control and overstress the aircraft and/or the control sjistem.

I I

Various types of synthetic feel devices ar,e used to simulate control surface loadings. They vary in effect aria comp16@y depending on the aircrddt 'type. Sometimes feel is not provided onpt= -aileron system.

-

I , --, I

I 1 \\ ' I I

1

Spring Feel I I 1 I ' I 1

,' L--/'' I I ' .I \

with-; spring-feel unit the pilLt pulls (or-pbshes) on a sp&ng.- he spring may be placed directly in line with the control linkage (as shown in figure 12) or it may be placed "outside" the control run as shown in figure 13.

SPRING BOX CONTROL SYSTEM

\ SPRINGS MOVEMENT

/ \ - SCREW TRIM WHEEL MOVEMENT

STRUCTURE \ SCREWTHREAD

Fig. 12 SPRING FEEL UNIT - 1

A s the unit moves so it will move the complete system - via the springs - to a new neutral setting - thus trimming the aircraft.

Figure 13 shows a more common method of connecting the trim actuator to the flying control system.


CONTROL COLUMN TO FLYING CONTROL

SPRING FEEL UNIT

Fig. 13 SPRING FEEL UNIT WITH TRIM ACTUATOR

"Q" Feel (figures 14, 15 & 16), - - ' ip -' , \ \I I

QUESTION Can you think of any advanthgqs/disadvantages of the spring feel system? (5 mins). 1 I I

I

I I

ANSWER Advantages: * I G h ' e a p . I I

ANSWER Advantages: * I G h ' e a p . ,

f -Light. , r - , * Maintenance free. 1 1

I 1

I UlSllL. \ r - * Maintenance free. 1 1

I 1

~ i s e d v a n t a ~ e s (one ar least). The force is constant in relation to I - airspeed and therefore_t&es'no i account of how fast the aircraft is

-- - --flying. For this reason Q1 feel'is often used. --

DIAPHRAGM

Fig. 14 SIMPLIFIED "Q" FEEL SYSTEM (Q POT)

QUESTION How does it get it's name "Q"? (1 min).

ANSWER It got it's name "Q" from the formula q=%pv2. I t is the dynamic equation and capital "Q" is used for "Q feel".

- 14- rnoodull l A-958

A device which gives the pilot feel proportional to "q" is the "Q pot". Used on a number of aircraft, its popularity arises from its simplicity. A disadvantage it possesses is that because of the small pressures involved, large areas of diaphragm are necessary to extract any useful force from the unit directly.

The component consists of a sealed cylinder containing a diaphragm separating two chambers. Attached to the diaphragm is a sliding rod operating on rollers which is pulled out of the casing (against Pitot pressure) with movement of the control away from neutral. A base pressure is usually provided by a spring. The base pressure provides for a minimum feel and Q feel is added to this.

On some systems a double pot is used to increase the feel force.

Hydraulic Feel Simulator (figure 15)

This is similar in principle to the Q Pot except that the dynamic pressure is increased hydrau~callywhich then allows for the construction of a smaller unit; -

I

PlTOT PRESSURE

Fig. 15 HYDRAULIC "Q" FEEL SIMULATOR

Construction

A loading jack, connected to the pilot's control in such a way that it extends whenever the control is moved from neutral, is fed with hydraulic pressure in direct proportion to dynamic pressure. Pilot stick force is thus also proportional to "Q".

Operation

The hydraulic spool valve is a t all times balanced between a down-wards force from the diaphragm (Pitot - Static = Dynamic) and an up-wards force due to signal pressure on the spool valve, which is trapped between it and the loading jack. As velocity increases so does dynamic pressure and the spool valve is pushed downwards, admitting supply fluid so that signal pressure increases.

However, an increase in signal pressure also increases the up-load on the spool valve, until it is returned to the neutral position. Signal pressure is now again proportional to the new dynamic pressure.

A reduction in velocity unbalances the spool valve upwards opening signal pressure to return thus reducing signal pressure until it is again proportional to dynamic pressure.

A control movement causes an increase in the signal pressure line (the links are so -- arranged that a control push or pull always causes the loading jack to be pulled_out). unbalancing the spoolvalve upwards opening the signal line tc, return, so that signal pressure f d s until the spool v&e is agaiibalanced, at which time it will close. I I 1

I

I i I I

Signal pressure, against which khe pilot is always having to lwork to move the controls, is therefore always prdcisely proportional to the dynamic pressure acting on the control surfaces.

I

Trimming

Since the control surfaces are now made irreversible by the-PFCUs normal trimming by means of tabs is not possible.

The spring feel unit is often used as the trim mechanism.

Trimming of powered flying controls is usually achieved by setting the whole system to a "new neutral". This means, for example that to trim the aircraft longitudinally nose up the complete elevator system is trimmed so the elevators are slightly up when the pilot flies "hands-off". The neutral position of the links to the unit can be changed or the unit itself can be moved to a new neutral position on the structure.

Note. Longitudinal trim systems on most large aircraft make use of a trimmable tail plane (stabiliser) which works independently of the normal flying controls - for most aircraft.

In figure 12 the whole spring unit can be moved to the left or to the right by the operation of the screw trim wheel-connected by trim cables from the flight deck or electrically operated. The systems in figures 13 and 16 use electrically operated trim actuators. Study the drawings and make sure you understand how the systems work.

QUESTION Where is the feel unit located in the flying control system? (2 mins)

ANSWER Between the pilot's input and the next unit down stream in the system - on most aircraft this would be the auto pilot servo - but on some aircraft it is the PFCU.

STATIC PlTOT

STRUCTURE

CONTROL SURFACE I

-1 PFCU

@- -,

CONTROL COLUMN

I Fig. 16 TRIM ACTUATOR/ WITH Q FEEL UqIT

I 1

- - / / / ,

I -- -.. \ I \

r 1' THE AUTO PILOT SYSTEM I / I

The auto pilot (Alp) system is capable of holding the aircraft qn any selected headirig bd.the aircraft will retuqn_t thhat heading if displdcdd from it. The A/P keeps the aircraft stabili&d-in-flight,&d accepts navigation commands for course headings etc. It will also fly the aircraft to an automatic landing. On some aircraft will also control nose wheel steering.

Principle

With reference to figure 17. In principle the system is made up of an outer loop and an inner loop control system. The outer loop control is a non-feed back loop and is made up of inputs from the pilot and other systems such as navigation systems etc. These signals are sent for computation before going to the inner loop computer.

The inner loop system consists of two negative feed-back loops. The aerodynamic coupling is connected with the aircraft movement and is sensed by rate and displacement gyros. Both gyro inputs send signals to try to negate or cancel the effect of the original command signal.

The other feed-back signal is servo actuator position (or control surface position) being fed back to the A/P computer - trying to negate the original servo signal input.

AERODYNAMIC COUPLING d

I - I FEEDBACK SIGNAL

C

INNER LOOP

I

DISPLACEMENT GYRO

- MECHANICAL CONNECTIONS - ELECTRICAL SIGNALS

INNER LOOP COMPUTATION

AIP SERVO PFCU

CONTROL SURFACE

OUTER LOOP MODES

OUTER LOOP COMMAND COMPUTATION

CONTROL

OUTER LOOP

I Fig. 17 INNER & OUTER LOOP CONTROL

I I ~ I

The displacement gyro senses the actual amount of aircraft angular displacement and produces a signal proportional to this displacement. The rate gyro produces a signal proportional to the rate of movemen( (@d senses direction).

, , I I I /

1 I 1 1

On many aircraft one cornputef (bften duplicated) will cany; out all the computations-for both loops. - -- --

General

The purpose of the A/P is to reduce the fatigue of the crew on long flights by taking over the routine flight control of the aircraft. The pilot can engage and disengage the A/P at any time. If he/she wishes to change the attitude/ course/altitude of the aircraft whilst the autopilot is connected he/she may disconnect the A/P and fly the aircraft manually to the new attitude/course/ heading and re-engage the A/P (allowing for A/P input corrections), or he/she can select the A/P to do the task via the A/P mode control panel.

Auto pilot control can provide for one, two (usually elevator and ailerons), or three axis control (lateral - ailerons, longtudinal - elevator and directional - rudder).

There are normally 2 autopilot/flight director systems fitted to the aircraft (some have 3) and the pilot can select either A/P 1 or A/P 2. With glide slope approach/go-around mode engaged both are selected with A / P 1 active and A/P 2 on standby. In all other modes only one A/P is selected.

- 1 8 - rnoodull l A-962

One A / P can be selected on the ground provided the engines are not running. Once an engine is started the A/P is automatically selected out. Once airborne the A / P can be engaged 5 seconds after take-off.

GYROS AUTOPILOT - COMPUTER SERVO PFCU CONTROL

CONTROL PANEL

CONRTOL SURFACE POSITIONAL FEEDBACK A Fig. 18 BASIC A/P SYSTEM

Gyros (mechanical or laser) to sense the displacement and ra te of change of the aircraft. , 1 I

Signal conditioning e4uipment (amplifiers, A to D Lonverteks, etc). Computers. I - - _ -\ ' \--_/ , Servos to move the cohtrol surfaces (electric or hydraulic). Mode controller, situdte'd on theIlflightdeck.

I

I ' I I

; i

The gyros and other sensing equipment develop signals which are processed in a computer and the resulting signal is sent to operate the autopilot servos. The output from these servos moves the appropriate flying control system PFCU to move the control surface. There is feedback from the servos to the computer (not shown in the drawing).

Most systems can be described in terms of their major channels - rudder, aileron, and elevator. Inputs from the pilot's mode control panel, navigation inputs etc go to the computer where they are digitised and summed with other inputs. The output is an analogue signal, either to a n electric servo motor or a solenoid to select a hydraulic servo. This servo will move the control system to input a signal into the PFCU servo valve.

The PFCU will move the control surface and a (negative) feed back signal is sent to the computer. At the same time, as the aircraft moves, so the gyros will pick up this movement. This aerodynamic coupling will provide another negative feedback signal to the computer to control the rate and displacement of the control surface.

The Three Channels (figure 19)

This shows in more detail the various inputs to a typical A / P system. The summation devices (shown as a cross within a circle) are parts of a computer where the various input signals are worked on and summed before being outputted. The output signal is amplified before being sent to the A / P servo motor (which could be a hydraulic actuator).

There are two main signals for the rudder channel which will determine when and how much the rudder will move. The first signal is a course signal obtained from a gyro compass system. A s long as the aircraft remains on the heading that it was on when the auto pilot was engaged no signal will be produced. Any deviation will cause the gyro compass to send a signal to the rudder that is proportional to the angular displacement of the aircraft.

The second signal received by the rudder channel is the rate signal. This signal is proportional to the rate of turn. The faster the aircraft is turning the stronger will be the signal. This comes from a rate gyro.

--

- -

PITCH, POSITIONAL FEEDBACK ~ CONTROL WHEEL

I

PITCH RATE

I I

--

GYRO

I l i VG 1 1 PITCH

'1 I I /

ALTITUDE I

&SPEED I

C ADC , 1 I

-- - ILS GLIDE SLOPE I- -= - - -

NAV W

E

ROLL SERVO

YAW YAW SERVO

SENSOR

Fig. 19 THREE AXIS SYSTEM


These two signals are combined, processed, and amplified (in a computer), and the resulting output is sent to the rudder servo. This will turn the rudder in the correct direction to bring the aircraft to the correct heading.

As the rudder moves a feed back signal is sent back to the amplifier. This is a negative feed back signal and is summed with the input signal (from the gyros). When the two signals are equal the rudder has reached the desired position, the output to the rudder servo stops and the rudder movement ceases. The rudder will stay in this selected position until the aircraft is in the new heading when the reverse will happen.

There is a cross feed to the aileron channel to produce the correct amount of bank for any given rudder deflection.

The aileron channel receives it's signals from a roll rate gyro which may be the gyro horizon indicator (on older aircraft), the turn control (on the flight deck) and navigational computers. There is cross-feed to the rudder for co-ordinated turns.

-- --

I -- . - -- -- - -

- L When .commanded the ailerons m6-d a feed back signal iCproduced. This signalis summed, and when equal to thd input signal the obiput s i h a l to the servo Leases and the ailerons stab moving. To keep the airckaft in a bank some aileroh $dl have to be maintained which'is carried out by signal summation.

I /' 1 1 / I i 1

On return to level flight the inpLtsignd\td the servo becomesismallei and the feed back signal begins to movk +e ailerons back to the trdl position. As the aircraft reaches the lateral position so the1 ailerons should reach their trail

' I I t

1 i

\ -- - ,-' I I

-- - 1 \

The 'elevator- channel works similar-to-the aileron channel7-

Figure 20 shows the general arrangement of an A/P system. Study the drawing and note the various data links and that on most systems there would be duplicate A/P servos.

Flight Controller (Mode Control Panel)

This is manually operated to set the various parameters and cause the aircraft to perform manoeuvres. Additional command signals can be sent to the A/P Flight Director computer from the INS, VOR, ILS, and DADC etc. The A/P system may be co-ordinated with Stab Trim, Mach Trim, Yaw Damper and Thrust Management systems. The auto pilot can be engaged and disengaged electrically or mechanically depending on the system.

While the autopilot is engaged, operation of the controls on the flight controller will cause the aircraft to climb, dive, or carry out a co-ordinated turn. The engage switch is used to engage or disengage the autopilot, and there is also a disengage switch on the control column.

COMMAND SIGNALS

Fig. 20 TYPICAL A/P SYSTEM

Sensing Elements

The directional gyro, turn and bank gyro, attitude gyro, rate gyro, and altitude control are sensing elements. Though signals are generated from other gyros on most modem aircraft. The gyros are usually laser operated, housed within units in the aircraft and are duplicated or sometimes triplicated.

Note. For details of the operation of the mechanical and laser gyro refer to the book entitled "Instruments" in this series.

The Computer

Much data is analogue (a continuous variable - eg the pressure transducer giving a continuously variable output on the dc ratiometer system as described in the Instruments book). For a digital computer to understand this, the signal must be converted to a digital signal. This is done by a analogue to digital converter (AID).

The computer will then "work" on this digital information. But for most systems outside the computer system this digital information is useless. So for the digital information to, say, drive a servo motor in a control system it must be converted back to an analogue signal and amplified. This is carried by a D/A converter. We then have an ordinary analogue voltage/current signal that can be used to drive motors, relays etc.

- - - 1

- --

I '_ compute& on aircraft pertor&functlons such as:

I I

* Air Data Storage (Air Data Computer - DADC) I l I I I * Automatic Flight control (FCS or AFCS) I

* Inertial Navigation (INS)' etc-- ' / , - - _ . -

I ' , I I \ /'

\ I - A basic A/P computer , will have: l 1 1 ' I I

* A power supply. I 1 I / I

-* A data bus - from which it--il/r?ceive information -) (from L other - c o m p u t e r s ) - and send-information to other computers. * Inputs from the pilot's mode control panel/control column buttons. * Sensor inputs. Gyros, altitude sensors etc. * Control surface position monitoring inputs. * Control system servo output signals - analogue electrical signal to

solenoid operated hydraulic valves - usually.

The inputs to a modern autopilot/flight control computer will include:

Mode control panel. Cross-feed from "other" A/P FCC (Flight Control Computer) computer. IRS - Inertial Reference System - ground speed, attitude etc DAD - Digital Air Data - altitude - airspeed etc. VHF navigation - VOR - G / S etc. Rad alt. FMC - BITE display etc. Position sensors - for all the controls. Autothrottle.

* Discreet inputs including: - Weight switches. - Landing gear micro switches. - Control system force limiters. - Flap micro switches. - Hydraulic pressure switches.

Outputs to other components / systems include:

Flight deck panel displays. Audio visual warnings. AD1 (Attitude Director Indicator). The other A / P FCC computer. Servos - elevator, aileron, rudder. Mach trim. Stab trim. Speed trim. Yaw damper.

- - - * _Nav systems. - - -

I -

Servos

I I

Within the flying control system th~y-are -- fitted in front of the PFCU or combined within the PFCU. I

I

Note that when the servo is connLcted in such a way that its h p u t moves the whole flying control system (including the controls in the flight deck) it is called a parallel system. When it is connected in such a way that its input only affects the flying-control system "downstreamn-(the pilot's controls do not move) it is called a series system.

The same principle apples to trim actuators and yaw dampers but yaw dampers are always connected in series.

ARTIFICIAL FEEL AUTOPILOT SERVO PFCU - n m

Fig. 21 BASIC FLYING CONTROL SYSTEM LAYOUT

__C

Operation

In the electric type the motor of the servo is connected to the output shaft through reduction gears. The motor starts, stops, and reverses in response to command signals from the computer.

PILOT INPUT CONTROL SURFACE 1 I I L

In the electro hydraulic type a valve is caused to operate by signals from the computer. This will cause fluid under pressure to pass to a jack (part of the servo) which will cause the controls to move.

The servo may be a separate unit or combined with the PFCU and is usually duplicated with each unit being connected to a different hydraulic supply. In many cases when one unit fails the other will continue but with half the authority.

Hydraulic Servo Operation (separate unit)

With reference to figure 22. The control system is connected to the servo via the output crank and internal crank. When the A/P is disconnected the two detent pistons are withdrawn by spring action from the roller and these cranks move freely when the system is moved by the pilot.

When the A / P is selected the detent pistons move in to engage with the roller. In thts condition hydraulic pressure i supply is fed to the servo-piston to move the control system via the cranks. .T?iis-pressure suppiy iscpntroiied by the EHSV under command from the A/P codputer. The output ipgsition'LVDT (Linear Variable Differential ~rdnsducer) se$ds positional feed back data to the

i 1 i compute^. I ,

SOLENOID

- \ -. 1 ELECTRO-HYDRAULIC

SERVO V F V E (EHSV)

REGULATOR

Alp SERVO POSITION LVDT

SERVO PISTON CENTRING SPRING

POSITION I V n T 1 VALVl I2

. - -. . . - . . - - - . -..,-- --- -. .--, ,- --.-.., (RUDDER DETENT ENGAGE)

\ lNTLw fi IN ICtXfMP

ELECTRICAL CRANK OUTPUT CRANK SERVO V A V E /CONNECTOR TO PARALLEL

RUDDER SYSTEM SOLENOID VALVE 1 (RUDDER HYD ARM)

\Alp SERVO

FILTER COVER POSITION LVDT

OUTPUT CRANK TO PARALLEL RUDDER CONTROL LINKAGE

LVDT

1 rlvul run uu I r u I wwNn " ...-.--.,. LL CRANK

Fig. 22 HYDRAULIC A/P SERVO

Servo Operation (combined type) (figure 23)

The PFCU and the A/P servo are both housed in the same unit.

A/P disengaged. The pilot's input goes in a t point B of the input link which pivots about the (at present) fmed pivot A. This causes the push/pull link to move thus selecting the main servo valve. This causes the main jack to move and a t the same time cancelling (or trying to) the main servo selection. When the pilot stops his/her input movement the main jack will continue to move, thus moving the link about pivot B. This will cancel the main servo selection and the main jack will stop. This follow-up movement takes but a fraction of a second. (negative feed back).

A/P engaged. With auto pilot on, hydraulic fluid pressure is fed to the autoflight control jack under the control of the autoflight servo. This will move the link (about point A) via the summing links to select the main servo as before. Feed back to cancel the signal is similar to that described above.

- .

RETURN HYD SUPPLY

i - ! INPUT

I a ,

I I \

MAIN JACK I

FEEDBACK LINK

\ OUTPUT TO CONTROL SURFACE

PFCU HOUSING 4 FLUID FLOW

-Ct MOVEMENT

Fig. 23 COMBINED A/P SERVO & PFCU

Note that the drawing does no show a by-pass valve (omitted for clarity). In the event of hydraulic failure this would open automatically and allow fluid to pass freely from one side of the main jack to the other. The pilot would then move the control surface direct by moving the main jack.

QUESTION When the auto flight control jack moves what happens to the pilots input? (5 mins) (figure 23)

ANSWER The pilot's controls will move as well. So when the auto pilot is in operation the cockpit controls will move.

For most aircraft the A/P will disengage when:

* The pilot selects A/P OFF on the mode control panel. * The pilot presses the disconnect button on the control column. * The pilot moves the rudder pedals/control column with a force

greater than a certain threshold. * The other A/P is engaged (other than in glide-slope or go-around

mode) - for most aircraft. * The associated PFCU fails (on the B747). * Certain modes fail (eg B777 when all air data is lost).

SYSTEM MAINTENANCE

At specified times the system-willrequire a visual inspection-; cleaning; 1 1 . \ > . /- i.

~ U U ~ K ~ L ~ U I ~ , a id fcii~ti~iid C I ~ C ~ S . lt-i~$d~~ ~ q i i ~ C G ~ X Q G ~ I C G ~ X ~ ~ Z C C Z I C Z ~

and modifications. The followink is general jnformation andi rgferende must always be made to the AMM. I

I 1 1 i I , ' ' I

\ /

r -- . -/ 1

System Functioning -1 ' , / , ", ~ I- -

1. Refer td the AMM. 1 I

2 . Configure aircraft for functiona(tq include: - (i)' Pitot static ~ ~ s t e m ~ ~ u m ~ e d up". f 1 1

- 4

- ( i i ) Hydraulic po~er-on.(~/ i or 3 systems) r'-- -- 1

(iii) Electrical power on - check flight deck indications. 4. Ensure all personnel are clear of control surfaces and position

warning signs. 5. Auto pilot OFF. Check the control system for correct functioning,

ie a normal flying control system rigging check. (refer to the book on Flying Control Systems).

6. Select auto pilot ON. Check flightdeck indications and ensure it engages smoothly.

7. Move the controls on the auto pilot mode controller and check: (i) That the control surfaces move in the correct sense. (ii) That the pilot's controls move instinctively. (iii) That all the controls are co-ordinated in a turn.

8. Check pilot override by moving flight deck controls and noting disconnection of A/P - together with warning signals (a spring balance may be required).

9. Check disconnect switch on control column (note warning signals). 10. Check other system inputs as per the AMM. 1 1. Check pitch trim system. 12. Check duplicate or standby systems.

rnoodull l A-971

13. Check any BITE systems and any faults stored in the fault computer . 14. Check that all visual and aural warning systems work correctly. 15. Switch off system, check freedom of controls, and check for leaks. 16. Carry out duplicate inspections if required. 17. Re-configure aircraft. 18. Record all work in the Tech Log or work cards and sign.

Note

Most large aircraft have Flight Director Systems, Inertial Navigation Systems, Inertial Reference Systems (INS / IRS) , and various other navigation systems. These will all have test procedures stated in the AMM.

These are compatible with the Auto Pilot System and may control:

(a) Altitude hold. (b) Vertical speed. (c) Airspeed hold. (d) Glideslope. ----

I I

(e) ~utoianci. \ \ I

(0 Heading and Tracking on VOR,'ILS, INS, etc I I '

I /

Fig. 24 EXAMPLE OF A THREE AXIS RATE TRANSMITTER

EXAMPLE - AIRBUS A300 AILERON SYSTEM

Figure 25 shows a schematic of the Airbus A300 aileron and roll control spoiler operating system. Only one side is shown as the other is the same, except for the fitment of an A/P servo. While there is no need to memorise the system you should have a good understanding of it's operation and should be able to relate the principles to your own aircraft/aircraft experience.

Note the following:

JC The selection (on the down going wing - up going aileron) of the roll control spoilers by the operation of the pilot's handwheel via the EFCU (Electronic Flight Control Unit). The handwheel also operates the ailerons of course.

* The inputs and outputs to the EFCU - in particular the input from the ADC (Air Data Computer) and the output to the ECAM (Electronic Centralised Aircraft Monitor - for flight deck displays).

Fig. 25 A300 AILERON SYSTEM

The spring feel artificial feel units (2)

The Droop Actuator (2). Sets the ailerons to the droop position for take-off and landing. Gives extra lift similar to a trailing edge flap. Automatically set when the leading edge slats are selected. Does not prevent the ailerons from working in the normal manner, but biases the whole system so the ailerons on both wings are set lower than the trailing edge of the mainplane.

It does this by changing the position of the pivot a t the trim screw jack position.

* The Differential and Droop Unit (2), which sums the inputs from the flight deck handwheel and Droop Actuator.

* Three servo control actuators or PFCUs per aileron - each with its own hydraulic supply.

* Five roll control spoilers sharing 3 hydraulic systems.

* Aileron Position Transmitter to ECAM.

* Autopilot Roll Actuator (A/P servo). One unit with two hydraulic supplies (Green and Yellow). When engaged the right-hand side of the system is moved via the right-hand Differential and Droop Unit and the left-hand side is moved via the flight deck controls and the

- -

left-hand side Differential and Droop Unit. I , I \

J; Operation of the tAmXvsteem; Operation of the switch in the flight I

deck causes the motor to operate screw jacks via a cable system to set the whole systemto a new neutral.

1 1 1

With reference toifigure 26 note the following: ! ~ 1. The relative positions of-the major components. Pilots input - Stick

shaker =Artificial feel unit - Auto pilot servo - PFCU - control surface - Power supplies - Position transmitter to ECAM.

2. The forward detent bellcrank and the pitch-uncoupling unit - allowing 'one side' operation should a 'single side'jam occur.

EXAMPLE RUDDER SYSTEM

Figure 27 shows a schematic of a rudder system. Note the:

* Q feel. * Trim input into the system. * Duplicate autopilot and yaw damper actuators connected in series

with the main system. * Single PFCU - for clarity - normally there would be 3. * Control linkage movement from rudder pedals to rudder.

f !--.-.-.-.- .-.-. FEEL & 1

I L ~ M ~ T A T ~ O N THS POSITION

L.-.-.-.-.-.-.-.J TRANSMITTER

Fig. 26 EXAMPLE OF A POWERED FLYING CONTROL ELEVATOR SYSTEM - A300

blank

- \

Fig. 27 1 RUDDER SYSTEM I I

i I I

I I

I

I I I

~ 7 6 7 SYSTEM 1

1 - I

-- I

Figure 29 shows the general &,a&ement ,of the 767 rudder1 system and figure 28 shows a simplified schematic arrangement based on the ,same system.

It is a fully powered system us i&3 - - PFCUs - each with it's own hydraulic supply. ,

- - - -- --

ELECTRICAL COMMAND 0 = PNOTS FIXED TO AIRFRAME &FEEDBACK SIGNALS

I PlVOTSlLlNK CONNECTIONS ' H ~ R A U L ~ C SUPPLY,RETtJRN - = PUSHIPULL RODS BELLCRANKS

CENTRING UNIT

STRUCTURE

REGULATOR

i HYDRAULIC SUPPLYIRETURN

t ELECTRICAL COMhWND 8 FEEDBACK SIGNALS

Fig. 28 SIMPLIFIED SCHEMATIC RUDDER SYSTEM - B 7 6 7

Input is via the pilot's and co-pilot's rudder pedals. A cross linking system ensures that the right-hand side and left-hand side move together, but should one side become jammed, then a detent link or crushable link will allow operation of the other side of the system.

Most units throughout the system are either duplicated or triplicated (on some aircraft there are two rudders - which move together normally, but one will move on its own should the other one become jammed or inoperative for any reason).

Artificial Feel

Provided by a spring unit. With reference to figure 28, when the pilot moves the rudder pedals the cables will move and cause the push/pull rod to move link 1 about pivot B to cause the rod to move into or out of the feel unit - pushing against a spring and providing feel for the pilot.

Movement-of the rod will alsomove link3 about pivot D. This will cause push jpuii roci ZF .to move iink 4 aboui H. Tnis wiii pui an-inpui inio ihe servo valve of the PFCU causing it to move. ' I

I

/ I

I / /

1 1 -

LOWER ElCAS DISPLAY

YAW DAMPER ACTUATORS (UPPER L LOWER)

RATIO CHANGER FEEDBACK TRANSM17TER

RATIO CHANGER SERVO

ELECTRIC TRlM ACTUATOR 8 TRlM POSITION TRANSMllTER

FEEL 8 CENTRING DEVICE

\ UDDER AUTOPILOT SERVOS

(LEFT, CENTRE 8 RIGHT)

Fig. 29 THE B767 RUDDER SYSTEM

Trim

This is provided by an electric actuator attached between the structure and the artificial feel unit. Movement of the actuator will cause link 2 to pivot about A moving the feel unit to the left or to the right. This will set the whole of the system to a new neutral, change the position of the control surface and hence put a trim input to the aircraft.

Autopilot Servo

When engaged (using an analogue signal) detent pistons move to lock onto' the main servo piston using hydraulic pressure. Signal commands (analogue) from the A/P computer selects hydraulic fluid to one side or the other of the main servo piston and it will move. This will move the push/pull rod and cause link 3 to pivot about D and cause the whole of the system to move. This movement will put an input into the PFCU servo causing the rudder to move.

-

Yaw D,arnper

To try to combat adverse yaw associated with Dutch Roll, y a y dampers are fitted. When Dutch Roll occurs: sensors (&os etc) will send signals to a computer which will send a n analO@e -- - signal to the yaw d h p e r servo. This will put a small corrective input into the,rudder system. It clauses link 4 to pivot abbut F and put an input i i to the PFCU servo. 1 ~ 1 I

1

Thus, during Dutch Roll conditions the rudder will be gently moved continuously from side to side by a-small amount. This actionl will significantly reduce-the-effects of Dutch Roll.- - -

-

QUESTION Describe Dutch Roll? (10 mins)

ANSWER It starts from an aerodynamic disturbance causing the aircraft to yaw. This causes the wing on the outside of the yaw (turn) to speed up and hence gain lift. Thus the aircraft is both in a yaw and a bank.

If the aircraft is passively stable enough the combined directional and lateral stability's will cause the aircraft to move back to its original position about the vertical and longitudinal axes. If the aircraft is dynamically unstable then it will overshoot the middle position and end up in the same attitude but on the opposite side of the flight path.

The cycle will now repeat itself with the aircraft gently rolling and yawing from side to side. This 'falling leaf' type of motion has a frequency of about 1 to l/z Hz. There is generally no loss of altitude during this condition.

So yaw dampers are fitted as a form of active stability - using sensors, computers and powered servos - to counteract this oscillating action.

Ratio Changer

This device progressively reduces the range of the rudder with increasing airspeed. Some aircraft have variable stops for the rudder that are related to airspeed, some have fixed stops that are adjustable on the ground only.

Control Surface Position Indicator

Shows the position of the rudder on a CRT screen EICAS (Engine Indicating and Crew Alerting System). Most large aircraft have a system to show the position of the control surfaces.

-- - - -- -- . , I 767 YAW DAMPER SERVO -

I i ,

\

The yawdarnper system connehed to the rudder powered Qiflg control system provides! commands to damp ahy'undesired yaw and to provide turn coordination (figure 3 0). I -

I -- \

-- , /

With most aircraft two yaw damper servos are provided and they work in parallel, each with it's own indqpendent input using data fro& it's own yaw damper module. When workingi together their inputs are summed and if one works on it's own then the range - of movement is usually h&d.

I - - ;- --

Control Panel

Situated on the flight deck, it provides for system disengagement and inop status indication.

Press To Test Switch

Provides for press-to-test facility for ground testing of the system.

Yaw Damper Module (Figure 30)

This is usually duplicated and takes data from the ADC and the inertial reference unit to derive rudder commands. Signals are sent as analogue electrical signals to the yaw damper servo. The module monitors system performance, provides for manual selection and allows for automatic system testing. Displays are on the front of this module, in addition to the control panel, show test results.

PRESS TO TEST SIW v CONTROL PANEL

AIR DATA COMPUTER I 1 ELECTRO

WEIGHT SWITCH - POWER

115V ac SUPPLY -----c MODULE b L

I GYRO ROLL RATE & I I

TAS

ATTITUDE. YAW RATE. LATERAL ACCEL GROUND SPEED

-

INERTIAL REFERENCE UNIT

-.

PlTOT PRESSURE A OF A r

pc -

HYDRAULIC SOLENOID I SERVO

YAW DAMPER MODULE

FILTER VALVE 1 ~&LG

I HYDRAULIC

SUPPLY --)

RETURN -c .

- -- I - - I 1-1 I I

POSIT1 LVDT

I 1

I DETENT &CENTRING 1 SERVO PISTON MECHANISM

i , I I I

Fig. 30 767 I ' YAW DAMPER SYSTEM ' i 1

Yaw Damper Servo

Analogue electrical commands from the yaw damper modules control hydraulic pressure to the servo piston (ports C1 and C2) and movement of the piston, which is in series with the rudder control system.

OUTPUT PISTON ROD \

SERVOVALVE

Fig. 31

OUTPUT PISTON ROD

ELECTROHYD SERVOVALVE

VALVE

POSITION LVDT

GENERAL VIEW - B767 YAW DAMPER SERVO

FLIGHT CONTROLS - CENTRALISED WARNING SYSTEM (an example)

Remember - you should be able to read through this example of a centralised warning system and relate it to your own aircraftlexperience. You will not be asked questions on it specifically but you should know in general how the system works and be able to describe the operation.

The centralised warning system provides warning, by means of a red warning lamp on the pilot's centre panel, in the event of a powered flying control unit (PFCU) failure. Warning of an individual PFCU failure is simultaneously presented on the flight engineer's panel. An aural warning is provided when the aircraft is on the ground under take-off conditions.

This warning, which is intermittently operated, provides warning on selection of take-off rpm if:

1. Any PFCU pump motor, or the No 1 or No 2 feel simulator pump is in -

a failed condition and, , - r--- C

- -

2. Certain surfaces are-inc:orrectly positioned. ' \

I I

I '

The warning is also interconne{t<d with the weight switch ainh I the aititude switch circuits. I /

e l i 1 I -,

I .\\ I r

~ e s c r i ~ t i o n I I I

' I A failure warning lamp on the flight engineer's panel is associated with each PFCU aridpis dperated by a pre&ure-switch within the uni t , '~hese lamps are parallel-connected to the centralised warhing lamp and the--aural warning.

A failure warning lamp is also associated with each of the two feel simulator pump motors. These lamps, which are mounted next to the centralised warning lamp on the pilot's centre panel, are parallel - connected to the aural warning circuit only.

The centralised warning lamp contains a double-pole switch which is operated by depressing the lamp unit cap. The switch provides the means of cancelling the warning.

A micro-switch, within the engineer's pedestal, is operated when take-off rpm is selected, and connects the aural warning to the centralised warning system and to certain surface position switches. Micro-switches are operated by the following surfaces or their associated controls, as follows:

1. Landing flaps - one micro-switch. 2. Slats - one micro-switch. 3. Spoilers - one micro-switch. 4. Tailplane - two micro-switches.

The micro-switches, which are connected in parallel, are part of the aural warning circuit and thus provide warning if any of the associated control surfaces are incorrectly positioned when take-off rpm is selected.

The aural warning is isolated from the centralised warning system under flight conditions by the action of the weight switches.

Operation

In each warning unit there are three separate channels, each comprising relays, and rectifiers. Two PFCU low-pressure switches connect to each channel and as aLl channels are similar only one is described.

When the low-pressure switch within a PFCU closes, a supply is passed to its associated warning lamp. The same supply is paralleled to:

1. The appropriate channel within the associated warning unit. The ligb t ,_remaining ON untilcZiGelled by the lamp cap being pressed, or the

r -7 -- I fault rectified. The centralisedtvahming lamp goes out but the ' individual PFCU lamp itays ON.

2. The rectifier unit. This allows a supply to the throttle operated micro switch. When take-off rprn isselected the supply 4s sent to the weight switch operated relay- and if the aircraft is on the ground the aural warning and lamps operate.

I

I

The shpply cirduit (for the PFCU) low preisure warning swiich operates in a similar way.' .- - -

- -- - -

Control Surface Position Switches (Config Warning)

If the control surfaces are incorrectly positioned for the take-off configuration when take-off rpm is selected, micro-switches operated by the control surfaces, or their associated controls, operate the aural warning.

When the micro-switch is operated, a supply energses a relay. The relay contacts close and a circuit is made to the throttle-operated micro-switch which, if take-off rprn is selected, connects the supply to the weight switch.

If the aircraft is on the ground the supply is connected to the aural warning. Thus the aural warning can only operate when a primary flying control surface or a secondary flying control surface is incorrectly positioned and take-off rprn is selected whilst the aircraft is on the ground,

FLY-BY-WIRE SYSTEMS

The flying control system, like any system, has an input and an output. In fact it can have several inputs. The pilot is one of then - and normally the main one - and another is the auto pilot. The output of the system is the movement of the control surface - and hence the movement of the aircraft.

For the majority of aircraft the input is transmitted to the control surface via a mechanical linkage - push/pull rods, cables, chains etc. It is a simple system using technology that is easily understood. However it has its disadvantages.

The disadvantages of a mechanical system over an electrical/electronic system are:

* Heavy. * Requires more maintenance. * Complex - with many different moving parts. * Non self testing. * --Not so precise - play - backlash~tc. 7 I - -- , - - * Difficult to interface-~th-Cith3r,el~ctronic equipment, eg auto pilot. *

I Less responsive to actual aircrqt needs. I I I * : Less safe - a computerised system1 can have builtjn safety re: stall,

windshear, overspeed, ahd overload. 1 1 ' - ' ,

,' // 1

The advantages of a ~ l ~ - b ~ - ~ i r k s y i ~ ~ ~ > e , the reverse of disadvantages listed abbve, with one addition4 advantage. : I n electronic system allows the introduction of active controls dnd active stability.

, 1 I / 1 1 1

Note. Active controls are cont~ols that-operate automaticalJy to alleviate usually gust loads; Active stability is stability acliieved by the use of control surfaces and not by the design of the airframe. Some modem aircraft have no inherent stability and are kept stable only by computer control via the flying control surfaces - these are mostly military.

Operation - General

The pilot's input is transduced into an analogue or digital electrical signal at the control column/rudder bar end, and the signal is sent to a computer. If analogue the signals will be digitised and then compared to other parameters to check that it is satisfied that the aircraft's response will be within its limitations and any other pre-existing conditions.

If the computer is satisfied then it will send an appropriate signal to operate the PFCU.

The aircraft will only do what the computer allows - what is safe and what the aircraft is configured to do, eg the aircraft cannot be stalled.

Signal Transmission

In a Fly-by-Wire system the signal may be analogue or, more usually, digital electric/electronic. In a Fly-by-Light system the signal is visible or laser light pulses digitised down wires which are made of high purity silica glass.

Safety

Most systems will have a level of integrity to make then at least as safe, or safer, than a mechanical system. This will usually include more than one power supply; more than one signal transmission path; more than one computer per signal path; and more than one software supplier for each of the duplicatedl triplicated computers.

Should all else fail many systems have a t least some mechanical back-up. For example, the A 3 2 0 and A 3 3 0 have a mechanical linkage from the rudder pedals to the rudder PFCUs and the trimrnable tailplane has a direct mechanical cennPFfG6 - tn the flightdeck trim - wEee1.- ,

-

I

I

EXAMPLE - THE AIRBUS ~ 3 3 0 I

, I

The 330 uses a system with the follow~g'computers: I ,

- - - . \

* Flight Control Primary Computers (FCPCs) - 3 of. * Flight Control secondary Computers (FCSCs) - 2 of. * Slat and Flap control Computers (SFCCs) - 2 of.

-- is-- -

The first two-control the aircraft-in-pitch, roll and yaw and-the SFCCs control the slats and flaps. All surfaces are actuated by electro-hydraulic PFCUs.

Flight deck indication and maintenance functions use Flight Control Data Concentrators (FCDCs).

The number of computing lanes is determined by the integrity requirements to achieve equivalent safety and response functions to a mechanical system and dissimilar hardware and software is used between control and monitor functions within each computer to ensure maximum integrity.

Safety features include:

* A high level of redundancy, ie 3 FCPCs backed up by 2 FCFCs. * Dissimilar redundancy - different computer types, different

microprocessors, and different suppliers between each computer. * Each computer is in two separate units. * Separate power supplies. * Segregation of signalling lanes. * Mechanical standby for rudder and trimmable tailplane.

The Side-stick

This replaces the conventional control column with a left hand side-stick for the pilot and a right-hand one for the co-pilot fitted to the left and right control consoles.

The side stick connects direct to a sealed unit under the console. Connections to and from the sealed unit are all electrical.

Within the sealed unit are torque tubes, levers, dampers (1 roll 1 pitch), spring feel units (for roll and pitch), control stops and transducers (four for each control axis).

The stick is held in the neutral position by the spring feel units.

Pitch Control

Signalled from the flight deck by the-use-of a side-stick tormove the two elevators. There is also a TrimmableHorizontal Stabiliser (?;HS)-moved by the centre console handwheel. I ,

I 1

Two efectro-hydraulic servocontrols actuate each elevator. A tbansducer unit mounted adjacent to the servocontrols~on~each elevator sends theposition of the elevdtors to the Flight Control Primary Computers ( F C P ~ ) a n d ~ l i b h t

1- Control Secondag Computers., : 1

I I

In manual mode elevator control is performid from the side siicks, which send electrical signals to the FCPCs and FCSCs. The computers elaborate command orders to-the-servocontrols, dkpending odthe different co6rol laws.

The position of the control surfaces is shown on the System Display (SD) via the Flight Control Data Concentrator (FCDC).

In A / P mode, the Flight Management, Guidance and Envelope Computers (FMGEC) :

* Send the command orders to the FCPCs; the FCPCs transmit them to the FCSCs.

* Supply the side stock solenoids in order to increase the feel force threshold - helping to prevent inadvertent operation of the side stick - though the pilot can override this if necessary.

Each elevator is moved by a servo in the active mode whilst the other is in damping mode. In normal configuration the active servo is the inboard one. When large deflections are called for both servos become active.

With the loss of all electrical signals the servos are centred hydraulically.

Pitch trim is provided by adjustment of the THS. The pitch trim is signalled automatically in normal operating mode, ie the electrical mode. In emergency mechanical mode operation of the tailplane is achieved using either of the two interconnecting handwheels on the centre console. Mechanical control will override the electrical control.

SIDE STICKS

Fig. 32 PITCH CONTROL SYSTEM - A330

Roll Control

The ailerons and spoilers are signalled by the lateral movement of the flight deck side stick.

The ailerons (and spoilers) allow the following functions:

* Roll control. Inboard/outboard ailerons and spoilers (2 to 6) associated with the rudder ensures roll/yaw coordination during turns and Dutch roll damping.

* Manoeuvre Load Alleviation (MLA) . Inboard/outboard ailerons and spoilers (4 to 6) allow for load alleviation.

- Two inboard and outboard ailerons are provided on each wing; two electro- hydraulic servo-controls actuate each aileron. In manual mode, the aileron control is performed from the side sticks, which send electrical signals to the FCPC and FCSC.

The computers elaborate command orders to the servo controls, depending on the different control laws.

At high speed (Vc higher than 190kts in CLEAN CONF), the outboard ailerons are sewoed to zero. In A/P mode and in certain failure cases the outboard ailerons are used u p to 300kts.

When the RAT is extended, the outboard ailerons are not used, all the servo controls being switched to the damping mode in order to minimise the hydraulic consumption.

Fig. 33 ROLL CONTROL - A330

r - - - - - i

1 1 , 2 , 3 1 Im I I I I I I I L - - - - - J

r - - - - - i LGClU 1,2, w

L - - - - - FCSC

FLIGHT CONTROL SECONDARY COMPUTER

C

FLIGHT CONTROL P 1,2,3= FCPC 1,2,3 PRIMARY S 1,2 = FCSC 1,2 COMPUTER 0, G, Y = BLUE. GREEN,

YELLOW HYD SYSTEMS

+ARROWS INDICATE ACTUATION RECONFIGURATION PRIORITIES

Note. There is no aileron trim control.

The position of the control surfaces is shown on the SD via the FCDC.

In A/P mode, the FMGEC:

* Sends the command orders to the FCPCs; the FCPCs transmit them to the FCSCs.

* Supply the side stick solenoid in order to increase the feel force threshold to help prevent inadvertent movement of the side stick when A/P is engaged.

Each aileron is actuated by one servo in the active mode while the other is in damping mode. In normal configuration the outer servo is in the active mode. The damping mode of the other servo helps reduce flutter.

Should electrical failure occur the active servo will revert to damping mode ar ' ,

Yaw Control

Uses a single piece rudder actuated-by thi-ee independently shpplied hydraulic servojacks signalled by means of inte?cbnnecting conventiobal rudder pedals.

The rudher control associated with the aihrons and spoilers permits automatic rolljyaw coordination during tGrns and the damping of Dutch Roll. It is also used for airci-aft guidance on the'ground.

-- --- -

In case of total loss of the electrical controls, the rudder also permits yaw control of the aircraft by direct mechanical linkage.

Three hydraulic servo controls with mechanical input, simultaneously active, actuate the rudder. The servo control input is actuated from the rudder pedals and yaw damper servo-actuators through a mechanical control to which are associated:

A Two transducer units, which send the position data of each set of pedals to the FCPC and the FCSC.

J; An artificial feel mechanism, which elaborates the feel force laws for the manual control and A/P control. Its zero force position is adjusted by a trim actuator. A solenoid pennits engagement of the A/P artificial feel.

* A travel limitation unit, which limits the linkage travel at the servo control input in relation to the aircraft airspeed.

In the A/P mode the FMGEC:

* Sends the control orders to the FCPCs, which send them to the yaw damper servo-actuators and via the FCSCs, to the trim actuator.

-k Activates the A/P artificial feel law through supply to the artificial feel solenoid.

Interface with Braking System. The braking is controlled by deflection of the pedals.

Fig. 34 YAW CONTROL SYSTEM - A330

Interface with Nosewheel Steering System. The pedal transducer units send position data to the BrAnglSteering Control Unit (BSCU) through the FCPCs for the nosewheel steering system.

-

CENTRING MECHANISM

- - 7 - - .

- L c, lrllr rn.,san, 8 .mu,,,--z,. ,..WL

I I SECONDARY\

I COMPUTER

b

FLIGHT CONTROL I

I PRIMARY COMPUTER

ROFL SSTU c

TRANSDUCER

.c MECHANICAL CONTROL

T P 1,2,3 = FCPC 1,2.3 +ARROWS INDICATE S 1,2 = FCSC 1,2 ACTUATION El, G, Y = BLUE, GREEN, RECONFIGURATION

YELLOW HYD PRIORITIES SYSTEMS

I

r---I- 1 FMGEC 1 , ~ -

L---,, J L -- --

r - - - - - I SFCC 1.2 m

L----, J LFCSC

r - - - - - - I 1

c

I ADlRU I

L-- , - -J

r - - - - - l * LGCIU 1,2

L - - - - - J

, i / - - / i- -

I

1 ., ,\ I

RATE I

GYRO--+ FCSC 1,2 ' s1

1 -- r --

CONTROL MECHANISM

1,2,3

SFCC -+

LGClU -+

SECONDARY COHPUTER

u --tALERONS

TRIM ART FEEL

ACTUATOR THANMM -7 1 s U P ART FEEL 52

SPRING ROD

> (

U P SOLENOID r - - - - 1

-BSCU 1.2 FMGEC L---- J

BACK-UP YAW DAMPER

B UNIT 1 F::r?l

The rudder system is centred by a spring mechanism in the fin should all electrical power fail.

Yaw damping is operative throughout the whole flight.

Figure 35 shows a schematic of the complete A 3 3 0 control system - you should note all the inputs to the flight control computers. Figure 36 shows the location of the control surfaces.

FLT CON DATA EIVMU CONCENTRATOR

(2)

INDICATIONS- EFIS. ECAM. FLT RECORDER

WARNINGS-* NVC

MAINTENANCE - CMC

FCMC

SIDE STICKS

PEDALS c

Fig. 35 A330 FBW CONTROL SYSTEM

- - a A -- - - 2 S?C!IE!?S 5

i - -

2 VERTICAL ACCELEROMETERS-

' RATE GYRO UNIT - I 2 SPOILERS 4

HYD PRESSURE SENSORS c

TACHOMETERS 2 SPOILERS 3

2 LATERAL ACCELEROMETERS - FMGEC

SF,CC y

7 HSMU

ADlRU 2 ELEVATORS

RA -/,' b i-- '

LGClU + - - --

ECU P,. YAW DAMPER SERVO ACTR

TRV LIMIT UNlT

PEDALS - MECHANICAL CONTROL SERVO CTL - - ELECTIHYD

TRlM ACTUATOR - MECHANICAL CONTROL HANDWHEEL -

FLIGHT CONTROL COMPUTERS

,RUDDER TRlM IND

blank

RUDDER TRlM CONTROL SWITCH ----C

SPEEDBRAKE CONTROL LRlER-

THROTTLE CONTRIOL LEVER PRIMARY COMPUTERS

FAULTIOFF PUSH BUTTONS - --- - -

- 46 -

rnoodull lA_990

- E L E . c T R 0

H Y D "

- .! OUTBOARD AILERONS (2)

INBOARD 7 AILERONS (2)

* 2 SPOILERS 6

Electrical control: Hydraul~c actuation' ' A~lerons ' All surfaces ' Roll spollers ' Rudder trlm Mechanical control. ' Elevators ' Rudder ' THS THS ' Slats &flaps " Speedbrakeslground

spo~lers ' Rudder

SLATS

HORIZONTAL STABlLlSER

7-- - - - - -

-

- -1

I ! '1 I - -

I Fig. 36 A330 FLIGHT ~ ~ N T R O L SURFACES \

Electrical control: Hydraulic actuation: 'Ailerons ' All surfaces ' Roll spoilers ' Rudder trim Mechanical control: ' Elevators ' Rudder ' THS THS ' Slats &flaps " Speedbrakeslground

spoilers ' Rudder

I /

THE BOEING 777 FBW SYSTEM ' /'

I , \ I I \ 1

The Boeing 777, like the A330, lhCs a highli\,integrated flyink control system. Unlike the A330 it uses convenkional fligGt-deck controls. s\&alling between comput<rs is via ARINC 629 data bus with ~ a l o g u e signale corning from flight deck transducers and analogue ~ i ~ n a l s ~ ~ o i n ~ to the hydraulically powered PCUs.

> , 1 L-- - I-- I

Primary flying control surfaces on each wing are one aileron (outboard), one flaperon (inboard) and 7 spoilers (one spoiler being mechanically selected all others are FBW). The other primary flying control surfaces are the rudder, the elevator (2) and the stabiliser.

The system has three modes:

* Normal. * Secondary. * Direct.

In normal mode the primary flight computers compute control surface position from the following inputs:

A Pilot. * Autopilot. JC Windspeed and direction. * Phase of flight. * Attitude. A Altitude, etc.

Secondary mode is initiated if sensory information is lost. This means that the Air Data Unit, IRS and Secondary Heading Reference units have failed. The autopilot will drop out. The Primary Flight Computer (PCU) will still operate using pilot inputs but some of the protection devices such a s bank angle protection are lost.

In Direct Mode some of the computers are not used and the Actuator Control Electronics unit (ACE) controls the control surfaces direct (still using FBW). The pilot can select Direct Mode or it can be selected automatically under some failure conditions.

The flight deck controls consists of a conventional control column for control of the elevators with a hand wheel for control of the ailerons, flaperons and roll control spoilers. The rudder bar controls the rudder. These controls are provided with artificial feel and back-drive motors to move them in the correct sense when the system is in A/P mode. An aileron trim actuator is fitted to the aileron system.

There is a-direct control cable run fro-m the pilot's hand wheel to the PCU of spoilers 4 (left wing) and 11 (right wing) and there is direct kable control of the stabiliser trim from the flight deck pitch trim lever to the stabiliser trim module.

I

ACTUATOR CONTROL ELECTRONICS ADM - AIR DATA MODULE AIR DATA INERTIAL REFERENCE UNIT PFC - PRIMARY FLIGHT COMPUTER 2FRU 1 AIRPLANE INFQRHATION MANAGEMENT SYSTEM SAARU - SECONDARY ATTITUDE AND

- AUTOPILOT FLIGHT DIRECTOR COUPUTER AIR DATA REFERENCE UNIT

AFDC - Fig. 37 ARINC 629 DATA BUS SYSTEM

FLIGHT CONTROLS - BOEING 777


The ARINC 629 data bus is a twisted pair of wires transmitting data in both directions to all cornputers/LRUs (Line Replaceable Units). Each cornputer/LRU is connected to the bus by untwisting the twisted pair locally and clamping on a n Inductive Couple Unit (which does not cut the insulation of the bus). In operation each computer listens to the bus and waits for a quiet period before it transmits. It then waits its turn until all the other computers/LRUs have transmitted before transmitting again (the system of listening and transmitting is called 'protocol').

Transmitting and receiving is carried out on the same bus. Any computer/LRU can listen to any data on the bus and receive the data according to how its personality PROM (Programmable Read Only Memory) is programmed. In other words the computer's permanent memory knows what information on the data bus is for its use.

-- -

WHEEL POSITION XDCR'S

' WHEEL POSITION XDCRS

WHEEL BACKDRIVE ACTUATOR - RC ---

WHEEL FORCE XDCR

WHEEL BACKD ACTUATOR - L SPOILERS 4 & 11

ORCE LIMITER

FEEUCENTRING ASSEMBLY

Fig. 38 AILERONIFLAPERON FLIGHT DECK CONTROLS - B777

Having recognised that the information on the bus at that instant is to be 'read', it will take it in, put it in temporary store (RAM) and act on it according to its pre-programmed instructions.

The flight control system uses 3 buses (left, centre and right).

Flight deck control movement is converted into an electrical analogue signal by transducers (XDCRs) fitted to the flying control system under the flight deck floor. This signal is then sent to the ACE unit where it is converted into a digital signal.

The pilot's controls are connected via the ACE unit to the PCU. Other units such as the PFC and the ADIRU (Air Data Inertial Reference Unit) are connected into the system by the ARINC 629 bus and send data to the ACE unit. The drawing below shows a block schematic for the ailerons but the rudder and elevators are similar in principle.

H AFDC (3)

Fig. 39 777 FLIGHT CONTROL SYSTEM

The hand-wheels are connected electronically to the ailerons and flaperons and mechanically by cable to spoilers 4 and 11.

The flight deck control is connected to position and force transducers which signals the pilot's intention by an analogue signal to the ACE. This is in two- way communication via the data bus with the PFC. After digitising and comparing/summing with other signals an analogue command signal is sent to the PCU to move the ailerons in the desired direction. Positional feedback is sent to the ACE which controls the range and speed of movement of the PCU - and hence the control surface.

The flight deck controls have artificial feel to simulate air loads and trimming is achieved by biasing the system neutral by a trim unit actuator.

Should one side jam then the other side can be operated independently by overcoming the force limiters. Should the spoiler control cables become jammed then system operation is assured by the shear-out action of the cable pulley.

Autopilot

When A/P is engaged all three autopilot computers are selected and input data into the system. The back-drive actuator will move the pilot's controls in response to A / P servo commands. Whenever the A/P is engaged the back-drive actuators are active.

When autopilot is selected the PCU is controlled by the ACE, PFC and Autopilot Flight Director Computer (AFDC) via the bus. The AFDC will also send an analogue signal to the back-drive actuator to move the flight deck controls to correspond to control surface?novement.,Thus the system-simulates closely the chara~te6stics of a conventional mechanic$ flying control system.,

' I \ I I

I I I I

1 I

Fault Finding I , ~ 1 I

' / I

\ I !

/ /

Beside3 &hecking the AMM and' ~1hCthkre'i~ an onboard fadlt ' c 0 ~ 3 t e r which is \ accesskd? via the Maintenance dcc!ess ~ e r ~ i h a l (MAT) on th& flight deck.

1 I i I

The MAT comprises: I 1 - / .---

- * A display screen. - --

j: Keyboard. * Trackball - similar to a mouse - controls the cursor. A Selection switch.

A Portable MAT (PMAT) can be used which is plugged into the system at various points on the aircraft.

To use the MAT proceed as follows:

J; Check the logbook for any recorded defects. ~r Check any warning flags and/or displays in the flight deck (elect

power on). These are called Flight Deck Effects (FDEs). * Check for any displays on the EICAS CRTs. (EICAS = Engine

Indicating and Crew Alerting System) (elect power on). * At the MAT in the flight deck you can select:

* Inbound FDEs. * Line maintenance FDEs. ~r Fault history.

* Under each FDE there will be a maintenance message. Select MAINTENANCE MESSAGE DATA and the recommended maintenance action will be displayed.

Some tests can be carried out using the MAT. These can cause the control surfaces to move as well as the flight deck controls, so it is important to ensure that they are free to move with no obstructions and warning notices displayed.

FLY-BY-LIGHT

This system is similar to the fly-by-wire system except that the digital signals are transmitted down a fibre optic cable instead of an electronic signal down a data bus or electrical cable.

The pilot's input is sensed by an analogue transducer near the flight deck. This analogue signal is then converted to a digtal electronic signal and worked on ' - 7

nnm-1 ~+a- n- ~ n m n l l f ~ r ~ The ~ip-d is ~ewerteT1 tn digital light pulses for CI b " I I I p U L U L V* "VLaLr -LVA -. --- transmission down a fibre optic cable. At,the end of the transmission line the signal i s converted back to an electronic digital signal then converted back to an analogue signal to operate a solenoid in the powered flying control unit (PFCU).

I

The advantages of light transmissionof - - data over electronic digital transmission include: ,

J; Not ai'fected by electric or magnetic fields. j; ~ e s s prone to data corruption due to lightning Strikes. JE No risk of cross-talk. * Lighter. * Smaller. ~r Faster transmission rates. * More reliable.

The electronic digital data is converted into light pulses by a light emitting dioac and the pulses are carried in a covered glass fibre to the receiver. The receiver is a light sensitive photo-transistor which converts each light pulse into an electric signal.

ENERGY ABSORBING LOW REFRACTIVE INDEX GLASS COVERING

LIGHT I

\ OPTIC FIBRE OF LOW LOSS HIGH REFRACTIVE INDEX SILICA GLASS

TRANSMITTER

PHOTO TRANSISTOR

RECEIVER

Fig. 40 LIGHT TRANSMISSION

Fibre optic cables usually have several fibre optic strands. Each strand has a silica fibre optic core with a high refractive index covered by material that keeps the light in (low refractive index) and adds strength and bulk.

OUTER JACKET

STRENGTH \A BUFFER JACKET

SILICONE COATING

FIBRE

Fig. 41 FIBRE OPTIC STRAND

Several strands are put together to make up a fibre optic cable. Dummy filler strands might be included to keep the bundle cylindrical. The cable is protected and sirengthened by additionral materid-s,uch as tape a n d m i d yam.

I-- --\ , \ --

Each fibie optic core element is! very sma)l- about 0.0055 in (0.14mm) in diameter, with the fibre optic s{rand being about 0.035 in (0.89mm) in diameter. The fibre optic cable is about 0.2 in (5.1,mrn) in diameter. 1 , , 1

I ARAMID YARN OUTER JACKET ~

DS

SEPARATOR TAPE

OPTICAL FIBRE STRANDS

Fig. 42 FIBRE OPTIC CABLE

Fibre optic cables are easily damaged, so should be handled with care. Damage can be caused by:

* Twisting and pulling - as when pulling through a loom/bulkhead. A Compression - as when standing on it. * Routing through too tight a bend (maximum radius 1 % in).

Note. If damaged the cable may not show any external signs, so a visual inspection must be accompanied by a functional test/BIT.

Fibre Optic Connectors

There are several different types of connectors. The ends of the connectors are fitted with a small protective piece of glass, which should not be touched and only cleaned using the approved materials as laid down in the AMM. Always fit protective caps when connectors are disconnected.

Connectors should only be tightened by hand.

Some connectors are designed to be used in locations that will not warrant their frequent disconnection. These are sometimes called a Butt Type and have good light transmission qualities. Those that require more frequent disconnection are less efficient at light transmission.

FIBRE \

INTERNAL TERMINUS

PROTECTIVE I GLASS COVERS 1

--- - , Fig. 43 DETAIL OJKSINGLE FIBRE OPTIC! BALL

LENS TYPE CONNECTION I

1 / I STRENGTH MEMBER ~ 1 I CONNECTION

\ PROTECTIVE WINDOWS NUT

YELLOW BAND

PROTECTIVE BALL / LENS

CAVITIES (2)

/ ,COUPLING NUT

JGNMENT KEYS

Fig. 44 CONNECTOR - GENERAL ARRANGEMENT

Figure 44 shows the connector assembly containing 5 individual single fibre core element connectors. The connector assemblies are large compared to the cable, but it is necessary to make sure that the two parts mate precisely so there is minimal light loss.

WARNING

When disconnecting connectors it is important to make sure that the system is off, made safe and tagged. This applies in particular when the lenses are to be inspected/cleaned. Although the light from the fibre optic is of low power it can be intense enough to cause damage to the eyes. Never look into the ends of fibre optic cables

MACH TRIM

As the aircraft speed increases towards MCRIT SO the Mach effects start to inf luem the aircraft's stabilityand-performance. These effects vary widely between &craft. On some aircraft the effects can be pronounced and start as low as M0.7, on others the effects can be'mild and start at Lgund M0.9. However, one general symptom that affects most high subsonic aircraft is the rearward movement of the centye jof (C of P) of the wing. This effect (tuck-under) increases as the aircraft-fio~ts --- through the tranponiq s&ed range.

I I --

\ I

'\ ' _/,'

This rearward movement of the/C1 of P produpes a nose down iendency, which requires a downyhrd correction on the tail to overcome the knstability.

,I , I 1

A s this instability is a function bf the ~ach'nurnber a systek 'sensitive to Mach change;which-will autornatic&ly put-an-iLput into longitidin& trim, is required. Figure 45 shows such a system.

As the system is sensitive to Mach number and Mach number is a function of temperature (which is related to altitude), it follows that the system need only be activated at altitude - hence the use of a height switch. As the system is only required at high Mach numbers a Mach switch is also included.

An autopilot system would obviously take care of this kind of instability so when it is operative the Mach trim system is clutched out, but the autopilot may not always be engaged.

The pilot c& carry out manual trim if necessary without upsetting the operation of the system.

The Mach trim input is via the Mach Transducer and the Summing Mechanism.

When the Mach Transducer puts in an input, an output signal is sent (+ or -) to the amplifier to operate one of the relays that will motor the Mach Trim Servo. This will move the cable system and summing gears to select the servo valve.

As the Mach Trim Servo moves so it puts a negative feedback signal into the Summing Mechanism - thus trying to cancel the original input signal. When the negative feedback signal equals the Mach Transducer signal the output to the amplifier ceases, the relay opens and the Mach Trim Servo stops. Meanwhile the servo valve has been selected and the VI tailplane is moving.

The tailplane will continue to move until the feedback link 'catches up' with the position of the Summing Gears (also negative feedback) when the selector rod into the servo valve is returned to neutral.

The two feedback systems work together so that the differences in their timing is not noticeable.

Examples

Boeing 737 Mach trim is connected to the elevators through the autopilot system and on the Boeing 747 it is connected to the tailplane actuator. -

I

I I I

PlTOT STATIC

SERVO VALVE

RETURN

/ SUMMING MECHANISM . ELECT

SUPPLY

Fig. 45 MACH TRIM SYSTEM

blank

STALL WARNINGIANGLE OF ATTACK INDICATION

These systems are fitted to give warning of an impending stall; to provide for stick shaker/stick push warnings/systems; to initiate engine auto-ignition; to provide indications on the flight deck and data to the DADC etc.

Stick Shaker

This may be initiated by a stall-warning device or an angle of attack transducer.

The stall-warning device may connect the stick shakerlstick push directly via a micro switch to a dc supply. The angle of attack transducers send angle of attack data to a computer and signals to the stick shaker/stick push will come from the computer when stalling angle is approached. The computer will also get data on the configuration of the aircraft, eg flap and slat position, which will affect the stalling angle.

-- -- - - - 1 1 1 n 1 1 t t L n>-trnt - ~ ~ n + n r n 1 llL 3 ~ ~ ~ + 3 2 ~ ~ ~ ~ U\3U-Iy bAbbLLLL lLIVLVL uCLUbIIbU LV L L L ~ bVIILL \JY oLbIIL -- \ '7 close to the control column (actually on the column on some arcraft). When switched on it rotates an out-oflbplance whkel, which causes the control column to shake (also operates an aural karning)' - barning the pildt of an iApending stall. 1 , '

, I -' /

-- <\ I \ /

I ' \ / I -

Stick &kh , I ll I I

I ' I

I

/ This sysfem operates a pneumdtic jacklaktuator connected the elevator control s y s t ~ m . When operated, justbefore the aircraft is about, to stall, it gives a push tCf lieeControl column trpitcli the aircraft nose down. ~ l h e pilot can override it by the operation of a switch - t h s released the pressure in the jack. The force provided by the jack is such that the pilot can overcome it manually if necessary. When not supplied with pressure the jack moves freely with the control system.

Pneumatic supply pressure can come from a tapping from the jet engne - typically about 40psi.

The system may have the following inputs:

* Airspeed switch - increases the speed of operation of the system with reduced airspeed.

* Stick shaker relay, which receives the signal from the angle of attack transducer.

* Flap, slat and aileron droop position. When deployed increases the stalling angle.

* Weight switch - system activated only in the air.

Airflow Operated Leading Edge Stall Warning Horn

Fitted to some smaller aircraft and consists of a slot cut in the leading edge of the wing and connected by a plenum chamber and small diameter pipe to an air operated sound reed situated in the cockpit.

A s the angle of attack increases so the stagnation point gets lower on the leading edge, this causes a negative pressure to be felt in the plenum chamber and pipe line, which causes air to be drawn in through the sound reed. This causes the reed to give out a warning sound.

. . - - - -- - -

Fig. 46 PNEUMATIC STALL WARNING

Leading Edge Stall Warning Vane

Fitted close to the stagnation point on the leading edge of the wing on some smaller aircraft so that as the angle of attack is increased the upward airflow of the air at the leading edge will cause the vane to move up (normal angles of attack will keep it pushed down). As it does so a microswitch is operated causing a warning larnp/aural warning to come on in the cockpit.

Rotating Angle of Attack Probe

Consists of a moveable probe within the airflow, with two slots in it, connected to an internal paddle. The two slots in the probe are open to dynamic pressure (airflow pressure), which are connected to two chambers separated by the moveable paddle. The paddle is connected to the probe so that as it moves, so does the probe.

LEADING EDGE VANE

TRANSDUCER

ELECTRICAL CONNECTION

/

TRANSDUCER OPERATING VANE

TRANSDUCER - k c MOUNTING P W E

Fig. 47 LEADIYG EDGE-STALL WARNING VANE-' , ' ' \ I --/

I \ '

! I I I 1 I ~ A s the angle of,attack increases so the dynamic airflow moves towards the

I bottom slot, increasing the dynamic pressure in this slot and under the paddle. Thisrcaui&i it to move u p androtates the probe to move the dots down. This

-- --- action will reduce the dynamic pressure in the bottom ~lo;-and increase it in the top slot and cause the slots to take up a position where the pressures are equal in both. Thus the paddle will take up a position that is related to the angle of attack of the aircraft.

The paddle position is transduced into an analogue voltage signal by being connected to a variable resistor (potentiometer). This signal can be sent to a computer (DADC) where it is converted to a digital signal and stored/sent to all those systems that require alpha (a) (angle of attack) information. These include:

~r Stall warning - warning lights - aural warning - stick shaker - stick push.

* Engine auto-ignition - prevents engine flame-out with turbulent conditions in intakes.

* Instrument sys terns.

Note. Angle of attack is often called Alpha angle in pilot's, training manuals etc.

S LO

AIRFLOW

CHAMBER

POTENTIOMETER

Fig. 48 ROTATING ANGLE OF ATTACK PROBE

Trailing Angle of Attack Vane

This is a trailing wedge type aerofoil that 'trails' in the air flowing passed the aircraft and therefore is always at the same angle as the airflow. Like the PntGlngAnglr of Attack Probe it is mounted on the side o f the fuselage and nA , be fitted to both port and st&6omd~ides of the aircraft to allow for errors due to side slip etc. I ~

I

Its position can be transduced into an electrical voltage sign$ (analogue) by the use of a potentiometer and this signal & be sent to summing units or cornp$ters similar to the Movirig Probe type. I

I

Both Trailing Vane and Moving Probe type units are electrically heated to prevent ice fo5&tion. I

- -

- - -

TRANSDUCER

v

Fig. 49 TRAILING ANGLE OF ATTACK VANE

blank

Figure 50 shows the a circuit for the BAe 146. Note the following inputs/outputs:

* Weight on wheels (squat). * Flap position. * Test. * Power supplies. * Airspeed. * Fail.

DE-ICE

SENSOR

FLAP INPUT

- AOA AIRFLOW SENSOR VANE -

F -

AC POWER I

DC POWER 1 I

RATE INHIBIT - - , AIRSPEED TRANSDUCER I

ONE CHANNEL SHOWN . _ -1

I TWO CHANNELS PER SYSTEM [ --

I I '\ '

AIRSPEED TRANSDUCER AC POWER DC POWER

- -- AOA-AFFLOW SENSOR VANE i

WARN VANE EXCITATION

FAIL

1 T E S T FLAP INPUT

Fig. 50 ANGLE OF ATTACK/WARNING CIRCUIT - BAe 146

CONTENTS

Page

Aircraft fuel systems Fuel tanks

Flexible tanks Rigid tanks Integrd tanks

Refuel/ defuel systems Open orifice refuelling -

_ rressure refuelling I Defuelling

Draining Engine fuel feed systems

Basic system I

Bopst pumps Fukl lines Water scavenging I I 1

Large kircraft system Fuel jettison systems I I

Fuel trim systems Fuel quantii$indication

ManuZl methods ---

Mechanical instruments Electrical instrument systems 42 Electrical/ electronic systems 42

The capacitive system 42 An electronic system 49 An ultrasonic system 51

Exercises 54 Answers to self assessment questions 54

AIRCRAFT FUEL SYSTEMS

Before we start it might be a good idea to define what is meant by 'an aircraft fuel system'.

SAQ 1 Describe what constitutes an aircraft fuel system, where does it start and at what point does it change to the engine fuel system?

A fuel system consists of various sub systems and these can be grouped as follows :

Storage system. There needs to be some form of storing for the fuel prior to it being used by the engines. It is stored in fuel tanks along the aircraft's longitudinal C of G in the wings and centre fuselage. On some aircraft fuel is also stored in the fm/ tail section of the aircraft (trim tanks). The capacity of these tanks is small but the fuel is-also used as a trimming medium - the more fuel that is pumped to the trim taflks-theheavier the tail of the aircraft gets. Useful for ordinary in-flight trimming as well as Mach trirnTihg.

'

I I

Refuel/defuel/drain system. used for fuelling the aircraft and at the same time removing any entrapped air. Sometimes (generally for maintenance) the fuel needs to be removed from the t h k s . - ~ v ~ n ~ a f t e r a complete hefuel, there are still pdddles of unusable fuel left in the tanks and if personpe? need,to enter the tankl(for inspection purposc&s :etc), there heeds to be some means of draining that fuel' away. I

I 1

/

Engin'e fuerfeed system. There-needs to'be a system to feed fuel to the --- -

engine/s burners.. Although engin~~will'suck' fuel using theirdwn pumps, booster pumps are provided in the tanks to provide a positive flow.

At engine shut-down (or during an emergency) there needs to be a valve to cut the fuel to the engine and so various shut-off valves are provided.

Jettison system. For most large aircraft it isn't structurally possible to land an aircraft at its maximum all up weight (on long distance flights aircraft will take- off heavier than they are allowed to land).

During flight fuel is burned and the aircraft mass falls. If a serious malfunction causes a return to base just after take-off the aircrew need to dump some of the fuel overboard to reduce the aircraft mass). This is done using a fuel jettison system.

It is worth noting that due to environmental and cost concerns, aircraft do not jettison fuel often, they will usually opt for an overweight landing - with special checks on airframes, airframe systems and engines to be carried out by the maintenance engineer afterwards.

Indication system. Both the flight-deck crew and engineers require information about the quantity of fuel on board. A variety of systems are available to include mechanical, electrical and electronic systems. Most aircraft also have back-up systems.

STORAGE SYSTEMS

Fuel, for the conversion to heat energy and thence to kinetic energy in the jet engine/piston engine, is stored in tanks normally within the aircraft. On some (usually military) aircraft it may be stored in external tanks and may be transferred to the aircraft in flight via an in-flight refuelling system. Tanks may be either flexible, rigid or integral.

FLEXIBLE TANKS

A l - n r ~ l l ~ r l ' 0011 ' gr %sg tnks'. They %re ms-rle nf fl-IPI r ~ s i s t a n t polymer I UU" b C U A U U ""LA

materialsand are designed to fit snugly inside specially designed compartments of the airframe. They do have several limitations and as such are not seen much today, but you will still need an understanding of them, as they are still used.

I ' ~ ~ s u q l ~ placed within the wing structure, and due to the flexing of the wing the tank is made slightly larger than the volume in which it is placed. This ensures the Aning structure, not the tank, takes the stress loads of the fuel and any in- flight loads.

Bag tanks are - constructed from thin rubberlsynthetic rubber material and usually one of the following is used:

H~catrol. A green coloured Hycar based synthetic rubber available in two thickness', 0.020" (0.5mm) lightweight and 0.040" (lmm) standard.

Flexelite. A red or black Flexsyn material which is synthetic based and similar to Hycar. Also available in two thicknesses, the standard 0.045" (1.14mm) for tanks with capacity greater than 100 imperial gallons (454.61), and lightweight 0.020" for tanks with capacity below 100 gallons.

Marlite. A blue coloured material of a two ply Nylon and Terylene fabric impregnated with a fuel resistant barrier. This is the most common material as its properties include a working temperature range of +lOO°C to -60°C. Also, due to the fact that no 'fuel-extractable' elements are used, these tanks always retain their dimensional stability. This removes the need for stabilising processes that are required in most synthetic based tank fabrication. Usually only available in the one thickness, 0.020".


All three of these materials will have a protective and supportive Iayer around them. This is often made from a fabric impregnated with rubber. This makes them more durable but great care is still needed when handling. As they are flexible they will require support in the wing structure, particularly when they are empty. Figure 1 shows some examples of the most common supporting methods. (a) The tank is supported within the tank bay by a length of cord. (b) The tank has studs bonded to its external surface and these are pushed into corresponding stud fitting holes in the tank bay.

STUDS

TANK BAY STRUCTURE

\

EYE FlTT TAN STR

// BUTTON

REMOVE BY : , 'PEELING'

i

\ A - - - A - ----+- I pF\ SECURE BY SECURING REMOVING PUSHING L -

\ I CAMBERED STUDS I I

, -. \

- -

Fig. 1 FLEXIBI;EI TANKS - FITTING METHODS I 1 I I

A s these tangs are reasonably fragile installation requires car&fvl handling. Figure-2 shows some simple guidelines-ahd the AMM will specify the procedure, but in general:

1. Ensure both the workspace where the tank is folded and the aircraft cavity are clean and there are no sharp edges.

2. Make sure that all protruding rivets, fasteners, brackets etc, in the tank cavity are protected with rubber tape.

3. Fold the tank as laid down in the AMM.

4. Cover the cavity aperture edges, temporarily, with rubber or tape to prevent damage to the tank as it is draggedlpushed through.

5. Remove all personal lose objects from pockets etc before entering the tank area (a standard precaution with all tank work).

SAQ 2 List the advantages and disadvantages of flexible tanks?

- 3 - rnoodull l A-1 009

1. INSPECT TANK FOR DAMAGE. PLACE ON A TABLE PROTECTED WlTH THICK BLANKET OR SIMILAR

2. FOLD TANK AS STATED IN THE AMM. AVOID KINKS OR TWISTS

3. SECURE TANK 4. INSPECT TANK BAY WITH STRAPS T AREA. PASS TANK ALLOW REMOVAL THROUGH ASSESS

HOLE SUITABLY PROTECTED WlTH RUBBER EDGING. REMOVE SECURING STRAPS, UNFOLD & SECURE IN POSITION

1

Fig. 2 FLEXIBLE TANK - PREPARATION FOR FITTING

- -

RIGID -TANKS-

These do not have some of the problems associated with flexible tanks, but they tend to be heavier.

Figure 3 shows an older rigid tank assembly. Tanks are usually constructed from aluminium alloy and sealed during manufacture. Sometimes tanks are fabricated from glass-reinforced plastic or even steel (rare). Larger tanks normally have internal baffles to prevent fuel surge (movement from one part of the tank to another during aircraft manoeuvres). They will have cut-outs to allow for normal fuel movement within the tank.

Each tank is designed such that it fits into a specific space within the airframe with clearance to allow for attachments, pipe connections, removal, re-fitting etc. As with flexible tanks they will have the appropriate connections for refuel, defuel and fuel feed.

Each tank will have its own pumps and hence electrical connections. A s the tank is metal it must be electrically bonded to the aircraft structure via a bonding point.

Some rigid tanks (as well as some flexible tanks*) have an additional external covering that is designed to swell when in contact with fuel. When wetted with fuel it will swell and, with a small leak, seal the tank. This tends to make the tank self-sealing, though it has had limited success.

ATTACHMENT

COVER PLATES FOR INTERNAL F17TINGS

INSPECTIOr I PANEL

I Fig. 3 RIGID FUEL TANK - FUSELAGE FITMENT

I ] I

/ _,

/ i

* When-flexibletanks are covered with a-self sealing membrane-their flexibility is reduced and they are handled more like rigid tanks.

Due to their weight penalty and poor space utilisation rigid tanks are not often used on commercial transport aircraft.

'External' tanks as fitted to military and some light civil aircraft, and are essentially rigid tanks. There are also rigid tanks fitted inside a cradle and loaded into freight bays etc, to increase an aircraft's range (rare).

The B747, for example, had this facility in the early days, but not many operators used it so it has been subsequently removed.

INTEGRAL TANKS

Integral tanks are areas, usually within the wing structure, in which the actual load bearing structure (usually front spar, rear spar and 2 ribs without lightening holes) is sealed to fomn the tank.

The advantages of integral tanks are that they are easily maintained, are cheaper and save weight. They can suffer from leaks, however, and these show up as stains/ seepage on the underside of the wingslfuselage. In some cases the leaks can be enough to wet the hangar floor.

The leaks are caused by the flexing of the wings during flight. This causes minute movement between the skin/rivet interfaces and, even though the interfaces and inside of the structure are sealed, leaks still occur.

Figure 4 shows a typical wing structure of a large aircraft showing two integral tanks in the wings with an additional integral tank in the wing centre section.

During manufacture a 'seal plan' is established during wing construction to ensure the fuel stays in the sealed area. Figure 5 shows the process. Note that the thickness of the various sealants has been exaggerated to show more clearly the details.

Firstly all surfaces are carefully cleaned and joined together with a fuel resistant-(Polysulphide based) sealant as an interfay. This IS a two-part mlx that cures with chemical reaction, not with air. After the joint has been closed (rivetedup), a fillet seal is applied, which, in addition to sealing, helps transfer the structural loads from one &ember to the next.

I 1 I

I

CENTRE CENTRE TANK TANK ACCESS DOORS (3 OFF)

DOORS

Fig. 4 TYPICAL INTEGRAL TANK LOCATION

Following on from this and before the fillet is fully dry, one or two, brush coats are applied, again with the next layer being applied just as the previous one goes tacky. Finally a protective barrier coat is applied; this dries quickly and allows the manufacturing process to continue without contaminating the slower curing sealant underneath. These sealants can take three to four weeks to reach full strength. The same general process is used when-ever repairs are carried on the structure which is part of an integral tank. Of course, always consult the SRM.


CLEAN SURFACES APPLY SEALANT RIVET UP REMOVE EXCESS SEALANT

I I FILLET

APPLY FILLET APPLY FINAL COATS OF SEALANT

Fig. 5 INTEGRAL TANK SEALING

A word of caution about fuel resistant sealants. They all contain harmful substances in their activators; Either manganese oxide orrstrontium chromate. Both y e lethal and specific handling inst&ctions must bb-followed when using these prdducts. j I

I I ~ 1 I 1 I

Integral tanks have a series of access panels often under the wing; these allow access for rework or inspections. They require attention to seding. Also a t various Gositions inside the tank there q e baffle plates a n d fuel dams designed to redhck fuel surge - and to kegp the fuel nearer the boost p u h p inlets.

\ I I l 1

The surge problem is more common on thin wing sections, anti as a solution, the fuel may,be sent from the integral tank,to a collector t h k ' v i a a Fuel/No Air

i _- valve. The-p-umose of this v a l e is to ener&se a solenoid when air is detected going into the collector tank (this might happen at extreme attitudes) to stop the air going in.

The engines are fed from the collector tank so they continue to receive fuel. Not a common system but one that is still in use on some aircraft.

The disadvantages of integral fuel tanks include the initial cost of fabrication, difficulty in repairing and fuel seepage. They are not crash resistant in that any structural failure will lead to significant fuel spillage.

Leak Assessment

Tanks should always be kept fuelled to a certain level, as this helps prevent leaks caused when the sealant (or tank bags themselves) dry out.

During ramp maintenance about the only task the engineer has to perform with tanks is filling them up, and even that job is often contracted out to the fuel supplier - for major operators anyway. However, sometimes tanks do leak and the leak has to be assessed and the appropriate action taken.

Any external leakage on an aircraft fitted with rigid or flexible tanks may indicate that fuel has already collected in the structure of the aircraft itself. This requires immediate attention.

Integral tanks have 'allowable' leaks (see figure 6). The correct procedure for leak assessment is stated in the AMM. In general, ensure the tank is full, clean the area and make sure no external fuel is present. Allow a period of 15 minutes to pass, then carry out the assessment. Where these leaks are and what you do about them depends on the AMM. Where there is a wet area or stain the AMM may designate the area depending on the size of the stain/wet area.

A s a guide, running leaks are not permitted anywhere. Areas around engine hot sections and turbine containment areas are not allowed any leaks at all. Other open areas, that have good ventilation, can have seeps or stains. If a leak is outside the limits it needs repair, and this may be a temporary external repair.

T e m p o r q repairs require regular-inspections and would have to be perhanently fmed as soon as possible. This involves tank draining and entry. It would usually require re-sealing and possibly replacement of: rivets/an airframe repair.

I 1 I 0 0 I ( 0 , -

1 I I ' I

I ' I

0 up to 1.6' 0

1.5 to b"

0

- - STAIN

- - - - SEEP

HEAVY SEEP RUNNING LEAK

Fig. 6 LEAK INDICATIONS

Working Inside Tanks

Besides repair, tank entry, on larger tanks, is also needed during scheduled maintenance and structural inspections. Fuel is inflammable, toxic and a skin irritant so precautions need to be taken before and during tank entry.

SAQ 3 What precautions should be carried out before entering a fuel tank? Answer below.

1. Check the AMM. Ensure correct fire fighting equipment available.

2 . Shut off appropriate stop-cocks. Carry out appropriate electrical bonding. Empty the tank of all fuel and drain. On pressure refuelling systems this may be done using self sealing couplings to a bowser. On gravity refuelled systems this may require a drain- down into a suitable receptacle and the fuel sent for filtering and categorisation at a fuel depot.

3 . Open all access panels and if possible vent the tank (provide a supply of fresh circulating air) for a period of time.

4. Ensure the tank atmosphere is not explosive (ie all fumes are removed) - there is equipment available that can indicate this

--- information by sampling the air. - -

I - \ > -- \ - -

5. Ensure sufficient fre$h air is,available - if there is not then breathing apparatus is required.

' I I

6. Ensure the person enteringdhe tank has a 'permit to work' inside tanks and that he/lshedoes not suffer from claustrophobia. The

I Permit to Work is $ new pi&e of industrial legislation-inthe UK I 1 that,requires that all confined space working be controlled.

~ntroduced by the Fa~tories Act.( I I

7. - - Wear correct clothingto-including a tank suit (fuel resistant plastic -- --

coverall suit). Ensure shoe5, clothing etc havk-no sharp edges to damage the tank lining.

8. There must be a trained safety person present outside the tank who is in communication with the person inside the tank a t all times. He/she also needs to be aware of where to get help, should the need arise.

Inspection lamps, torches and any other equipment taken into the tank must be flameproof and spark proof.

In general, 'normal on-aircraft7 precautions, apply: power off, no food, drink, smoking etc. Wear rubber shoes and correct protective clothing.

REFUEL DEFUEL AND DRAIN SYSTEMS

Refuelling systems can be categorised into two groups:

* Open orifice refuelling (also called over-wing or gravity refuelling).

* Pressure refuelling, under-wing refuelling or closed line refuelling.

Open Orifice Refuelling

Also referred to as 'over-wing' or 'open vent' refuelling, this method of fuelling an aircraft is similar to the way most motor cars are fuelled on the garage forecourt. It is common on small aircraft.

On the top surface of the wing (or fuselage in some cases) there is a filler cap and tank opening. After operating the quick release fastener and opening the refuelling panel the refuelling cap is unscrewed from the tank. In figure 7 the refuelling cap is fitted flush with the wing surface and lifting the tab on the c;,, allows the cap to be removed (by twisting the tab or operating a lever).

The filler port may be connected to pipework delivering the fuel to all tanks or, more often, directly into the individual tank. The cap is usually connected with a lanyard to the structure to prevent loss and there is usually a bonding point.

I -

To rebel, the fuel nozzle from the bowser is bonded with its bonding lead to the tank and the nozzle placed inside the tank orifice. The nozzle lever is operated to allow' fuel to flow into the tank. Care needs to be exercised as the tank starts to get full, as it is a high delivery rate and the fuel will spray back out of the tank when the tank becomes full. Towards the end of the filling operation flow ratesshould be reduced and preparation made to shut-off the control handle quickly .

In addition care is needed when opening the refuelling cap to prevent contamination of the fuel - by rain, if raining - by fine debris - sand etc, if a gale is blowing.

Other potential pitfalls include the danger of walking on the top surfaces of wings and the possibility of damage to the wing from, filler caps, fuelling hoses peoples shoes etc. Also there is a risk of fuel imbalance between the tanks.

After refuelling the bonding lead should be disconnected and the cap closed.

The biggest disadvantage is that of time to refuel. Access to the tanks is difficult and requires step ladders, high rise platforms etc and the actual filling process can be slow.

Most small aircraft only have the one method of refuelling - gravity refuelling, but some larger aircraft have both systems fitted. The B747, for example, has the option of over-wing fuelling, but using this method would require up to 8 hours to fill the aircraft.

Over-wing refuel points on large aircraft are rarely used. Some aircraft are not fitted with them a t all, eg the B777.

FUELLING PORT

-

- - - ,

I \ \

I 1

Fig. '7 OVER WING F E L L I N G POINT/ 1 I ' , I

I I

-- - - / 1 Pressure Refuelling -- - . -- , /

I I

1 -,

This i s a,system ,of refuelling whete fuel uhder pressure [rnax &bout 50psi (345kPa)) is supplied from a bowser, tank&,) or refuelling p i p i n g vehicle. (It is common 'at large !airports to have the fuel' pumped underground. A vehicle connects irito the ground conndctian [gt~r'lifting a steel codef plate] and pumps th6-fuzinto the aircraft); ThFbowserys (fuel tankek) fuel'hose is connected to the refuelling point (figure 8). From this single point there is a pipework system connected to all the tanks in the aircraft. The fuel is controlled into each tank by energising solenoids in the refuel valves.

Advantages of pressure refuelling include:

* Higher pressures and flow rates and shorter refuelling times.

J; Less risk of spillage.

* Ability to fill any tank with any desired quantity of fuel using the aircraft's on-board refuelling control system. Electrically or computer controlled.

~r Reduced risk of fuel contamination.

* Better access. Fuelling points are accessed from the ground.

* Reduced fire risk.


AIRCRAFT REFUELLING

EARTHING OR BONDING STANDARD CROCODILE CLIP HOSE CONNECTOR

SEALING CAP

GROUNDING PIN CONNECTOR SEALING CAP

W $ m w , ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ HOSE

Fig. 8 PRESSURE REFUELLING CONNECTION

i oper4tion

I I I /

I I

The e/e?trical power supply must: be ON to operate the various valves, indicatdrs and computers (if fitted), this can be provided from the aircraft 400Hz supply or even from the aircraft battery. Usudly power to the refuelling systems is removed when the aircraft Eecomes airborne, preventing inadvertent fuel transfer. This power supply also goes to the refuelling panel, which is usually located at the refuelling point. Usually these components are located behind a panel close to the mainplane - check the AMM.

SAQ 4 What ATA chapter of the AMM covers refuelling?

Figure 9 shows a typical refuelling system for a large aircraft, consisting of the pressure connectors and the refuelling panel. Figure 10 shows a typical refuelling valve.

With power available, and the refuel panel door open, the indications on the refuel panel will show (the system may go through a BITE check first).

The dust cap is removed from the refuelling point and from the hose connector. The aircraft bonding clip is connected to the refuelling hose and the hose connected to the aircraft fuelling point via the quick release bayonet type connector.

- 12 - rnoodull l A-I018

The action of connecting the hose to the aircraft opens both the valve on the aircraft refuelling point and the fuelling hose.

With the bowser pumping, the pressure will open the non-return valve and fuel will flow into the system that supplies all the tanks (sometimes called a fuelling gallery) -

FUELLING QUANTITY TEST BUTTON OVERFILL RESET PRESELECT UNITS SWITCH

FUELLING QUANTITY REPEATER INDICATOR

- ' \ pp

L

I

w w1 l la t /:m OIIR) I

/ I

; I i I , Fig. 9 REF~ELLING L - C,ONTROL PANEL '

r ' -

This fuel goes to the inlet of the refuel valves (figure 10). It passes through the upper removal check valve (or non return valve) and, because there is a calibrated leak to the other side of the diaphragm the pressure is felt on both sides.

The spring in the refuel valve body holds the valve shut. If a particular valve is selected open by operating the switch on the refuel panel, the solenoid will energise. This allows the fuel on one side of the diaphragm (side 1) to be ported back into the tank through the lower removal check valve. This ports the fluid faster than the calibrated holes allows side 1 to fill so the fuel pressure moves the diaphragm (to the right in the drawing), and the refuel poppet valve moves to the open position.

Fuel now flows into the specific tank until such time as fuel pressure is removed or the solenoid is de-energised.

Once there are some valves open, and this is indicated on the refuel panel with (blue) lights coming on, the valve position switch is moved and the pressure from the bowser can be increased. (It should be reduced prior to final shutting off).

Note that on this aircraft the diaphragm assembly can be removed without removing the valve body. The diaphragm assembly is the most likely part to fail and so it is useful to be able to change this without de-fuelling the tank.

n ' R::izMOVAL CHECK VALVE

POSITION SWITCH

/

SOLENOID VALVE -- -

SOLENOID

LOWER REMOVAL CHECK VALVE

- FUEL IN

' REFUEL VALVED CLOSED

n

- FUEL IN

'EN

REFUEL VALVE OPEN

Fig. 10 REFUEL VALVE

The gallery may have restrictors installed at certain locations to allow all tanks to be filled at the same time. Care still needs to be exercised to ensure an imbalance doesn't develop (more fuel on one side of the aircraft than the other would up-set the aircraft's C of G laterally).

Sometimes additional refuel points are provided to increase the speed of the fuelling operation - so more than one bowser can be used simultaneously.

Refuel valves can be mechanically operated but they are normally solenoid operated, using the fuel pressure to actually operated them open. This type of refuelling requires overfill protection to prevent tank rupture or fuel spillage.

Also, with some aircraft care has to be taken when refuelling /defuelling so as not to put the longitudinal C of G outside the range between the main gear and the nose gear.

The aircraft's C of G, for a nose wheeled aircraft, is just forward of the main gear and with some aircraft with highly swept wings it might be possible to fuel some tanks in the wings such that the C of G is moved aft passed the main gear - this would cause the aircraft to tip back on its tail.

For tailed wheeled aircraft the C of G is behind the main gear and the tanks are so situated so that it is not normally possible to overbalance the aircraft longitudinally.

Defuelling

Figure 11 shows a typical defuelling system. With the bowser connected, the non-return valves are de-seated-and the defuel handle is turn-ned. On newer

1 m-d . aircraft this may be a swiich sricciiorl on iile rciuci par1el.-tms-connecis iiie engine feed manifold (in the drawing) to the.refue1 points. The engine fuel feed boost pumps are used to provide fuel to the engine feed manifold and through the defuel valve to the bowser. i 1

- /

I I I

-- \ . -\ I - - '

I

JETTISON PUMPS

JETTISON NOZZLE VALVES

VALVES

Fig. 11 DEFUEL SYSTEM

Once the bowser has positive flow the pumps on the bowser can be selected to suck. Again, consideration has to be given as to where fuel is being taken from to prevent imbalances developing.

SAQ 5 How can a combination of the refuel and defuel systems be used to transfer fuel within the aircraft for maintenance?


Refuelling/Defuelling Precautions

Fuel is highly combustible and fuel flow will cause a build up of static electricity - which could cause a spark if the correct bonding has not been carried out.

The following precautions should be observed (also check the book in this series EASA Module 7 entitled Safety):

* Ensure the correct grade of fuel is used (AMM). See the books in this series EASA module 15 for jet fuels and 16 for gasoline and Diesel fuels). Check the bowser driver's log book to ensure that the required dip checks (quality control checks) have been carried out on that particular batch of fuel. This will also indicate the specific gravity of the fuel.

The aircraft needs to be bonded to earth, ideally through a purpose built bonding line but CAAIPs state alternatives for 'field' operations. - ---

The bowser must also be bonded to earth and this is often done through the tyres or a bonding chainllead hanging underneath.

The bowser and aircraft are bonded together. A bonding lead is reeled out from the bowser to the aircraft and connected to the earth point, which is often (but not always) near the refuel point.

The refuelling hose is bonded to the aircraft.

* - The aircraft (and bowser) must be in a designated refuel area that could contain any spillage.

* No smoking or naked lights allowed.

* Aircraft power should not be connected or disconnected during the refuel process. APU's if running should be left running for the duration.

* Ensure fire cover - provided by the airport fire service or through individual hand held CO2 or Dry Power extinguishers. Refuelling should not be performed in a hanger unless additional fire cover is available.

* No radio or radar transmissions allowed during the operation.

* No refuelling during an electrical storm.

* Cleanliness is vital with all refuelling components to prevent fuel contamination.

* Any engine driven ground equipment that is required should be cleared to run in fuelling areas (spark proof exhausts etc).

Draining

Prior to entering a tank for any reason the tank will need to be defueled and drained. For most larger fuel tanks the only way fuel can be removed from the tank during a defuel operation is via the boost pump, the inlet of which is normally not located completely at the bottom of the tank. This will mean that some fuel (typically 20 gallons - 901 for a large tank) is left in the bottom after defuelling. This has to be removed by draining.

Light aircraft are often without boost pumps in the tanks and they have to be defueled and drained completely from the drain ports.

Drain ports are also used for taking fuel samples. - - --

Figure 12c$hows a typical drain Po?. are located at the bottom of the fuel tank. For fuel sampling the primary poppet is pushed up, this de-seats the valve and allows the fuel from the bottomofthe tank to drain \into a container. During the above refuellingldraining operation the primary poppet vhve is removed, this allows the second~arypppet to seat and prev&nts spillage. A tool is insetted to allow the secondary poppet to de-seat and the fuel-to flow - into a cont airler . I \ I I

1 i

, '\ I

LOWER WING SKIN

PRIMARY POPPET

Fig. 12 DRAIN VALVES

Venting

As fuel is pumped into (or out of) the tanks (refuelling, defuelling, engine use) air has to be allowed in and out. Failure to do so could cause the tank to rupture during refuelling operations or a vacuum to occur when the engines are running and thus starve them of fuel.

When allowing venting careful consideration has to be given as a highly explosive fuel/air mix is being moved from the tank to the outside atmosphere. Also when allowing air into the tanks contamination may enter via the venting system.

For small aircraft the vent can be a small pipe connected to the top of the tank which is open to atmosphere. A ball valve is fitted to prevent fuel leakage in the event of inverted flight.

Figure 13 shows a typical venting system for a large aircraft - the purposes of which are to:

* Balance the air pressure within the fuel tanks with ambient air. * Allow for thermal expansion of the fuel/fuel air mixture in the tanks. * Protect the tanks from excessive internal pressures.

. fi *~~ RESET HANDLE FLOAT ACTUATED DRAIN VALVE B &X

SURGE TANK ACCESS DOOR

FLAME ARRESTED

CHECK VALVE

VENT SCOOP

OUTBOARD SURGE TANK ACCESS DOOR

Fig. 13 TANK VENTING SYSTEM

This system uses the top hat stringer sections of the wing to transfer air to the surge tanks. Here there are two valves, one inward and one outward with a flame arrester attached.


During fuelling fuel enters through the refuel valves to displace the air. Holes are provided in the top hat stringer sections (part of the wing structure). These sections are sealed in the same way as the integral tank structure. The air is able to pass along the stringer section and out to the surge tank. The surge tank has an outlet that is protected by a stack pipe. The volatile air/fuel mixture from the surge tank is forced overboard via a flame arrester and air outlet.

TANK MOUNTING

VALVE SEAT ,,,,,Pi"""'

- -- - - ELECTRICAL I

SOCKET MOUNTING FLANGE /

BOTTOM REED TUBE ASSEMBLY

Fig. 15 FLOAT SWITCH

- 19 -

rnoodull I A-1 025

A s the fuel in the tank continues to rise it needs to be stopped from entering the top ha t section via the same hole that the air is going out through. This is done using a float valve (see figure 14). Air is allowed to pass through the open valve but as the fuel rises it lifts the float and closes the outlet.

Overfill Protection

Open line refuelling presents no problem in terms of venting or overfilling but pressure can build u p in the tank/s when pressure refuelling. This could lead to possible spillage or tanklpipe rupture. So protection is built in to provide for this.

One way is to use a float switch (figure 15). As the level of fuel rises to the level of the float switch, the float rises, the magnet moves tolaway from the reed switch and it closes/opens and an electrical signal is sent to the refuel valve, de-energising the solenoid and closing the valve.

Note that the float switch shown in figure 15 has a lower and higher reed 1 mi -- --- L- - - - - 1 L- :---l:--L- L L - C C L swiic~l . + rlt: l u ~ e l u r i c ; rlldy UC; U ~ C U LU L ~ ~ U ~ L ~ L C L L L ~ L LIIE tzilk is getti7i~ k!! SG

reducing the supply prior to final shut-off. I

A rno'e sophisticated (computer) method is where the contents indication system is used to work out when the tank is nearly full and close the valves automatically. A s the fuel level approaches a pre-set 'full' indication the Fuel Quantity System removes the power-to the refuel valve. Some manufacturers call this Volumetric Top Off (VTO).

Some (lkrge) aircraft employ both systems. The first line of defence is the VTO systefn. Should that fail fuel is allowed to travel up the vent lines into the surge tankYA-float switch is fitted here and after a certain amount of fuel has flowed in, the switch will operate break all electrical power to the refuelling system thus shutting the refuel valves.

The fuel in the surge tank should not spill out, as a stack pipe protects the outlet.

ENGINE FUEL FEED SYSTEMS

Whatever type of tank is fitted the principle for delivery of fuel to the engine(s) is the same. Figure 16 shows a system for a light aircraft where the fuel feed is by gravity.

The tanks are higher than the engine and the aircraft is designed for moderate manoeuvres only so the fuel can be gravity feed from the tanks to the engine.

The tank selector valve is used to allow the fuel to be taken from either, or both the tanks. This allows the pilot to control the fuel and balance the aircraft (laterally).

A collector tank is provided below engine level, or a collector compartment is provided within the existing tanks to allow the aircraft to fly inverted. Under normal operation the collector tank/collector compartment has gravity fuel feed.

TANK SELECTOR

Fig. 16 A LIGHT AIRCRAFT SYSTEM -- - - -- -. , r -- -

- - - -

During inverted flight a float valvd clos;~ thk inlet and fuel is &d to the engine from the collector tank/collector dompartment by gravity. ~ h e c a ~ a c i t ~ ~ o f either system is limited and relies on the pilot riihting the aircraftbefore tdo long.

I I - - , I I

On low winged aircraft or aircraft whaq the tanks are lower than the engine/s or where there is a long pipe run f om the, tanks to the engibes-gravib cannot be reliedon to move the fuel. The fuel neeas to be pumped. 1

/ I I

The engine fuql pump does provide some ,suction from the t h k , but a positive pressure is required to keep the fuel-flow rate at an adequate level and to preventcavitation* in fuel lines anckcomponents. So a pump is i'nstalled in the line from the tanks to the engine.

* Cavitation is caused by a negative pressure in fluids and can cause aeration and serious erosion and wear on pumps and other components.

SAQ 6 If the pump in the system described above were to fail, will fuel get to the engine and what sort of redundancy could be build into the system?

Figure 17 shows a fuel tank system for a twin engned aircraft with low wings with some dihedral.

In figure 17 the outer tanks (1 and 4) will gravity feed to the inner tanks (2 and 3), but on some aircraft a pump system is provided. Tanks 2 and 3 are pump fed to the engines. The fuel supply can be controlled by the pilot by the use of the various shut-off valves and non-return valves will be provided to prevent fuel flow in the reverse direction.


Re-fuelling is by gravity feed and de-fuelling is by the use of the tank drains.

ALTERNATIVE SYSTEM ALTERNATIVE NO 1 TO GRAVITY FEED TANK SHUT-OFF TANK FEED PUMP VALVES

ALTERNATIVE NO 4 TANK FEED PUMP

I

Fig. 17 TYPICAL AIYRAME FUEL SYSTEM FOR A TWIN ENGINED AIRCRAFT

I '

I I '

Boost Pumps

Figures'l8, 19 and 20 shows a typical boost pump and its control circuitry. - .-

-- - A

when the pump is selected O N the motor turns the helical impeller at high speed (up to about 10,000rpm). Fuel enters the pump through the inlet and is moved through the impeller and the centrifugal impeller to the outlet.

The fuel inside the pump is also leaked' internally to provide cooling and lubrication - though is kept separate from the electric workings.

Boost pumps provide a low-pressure output, as the impeller style pump is more suited to high volume flow rates rather than high pressures.

Pressures are about 19 to 26psi (13lkPa to 179kPa). This is monitored, using a pressure reed switch and fed through the aircraft indication systems to provide flight deck enunciations (warnings) of any pump failures.

With reference to figure 19 note the 30 11 5v 400Hz supply to power the pump which is controlled by a 28v dc relay. Note the similarity of the pump windings to the windings of a 30 generator.


Boost pumps will normally be fitted on the front and rear spars of integral tanks. Sometimes they are housed in dry bays internally on the wing surface. Bag and rigid type tanks require additional supports for the boost pump fittings.

FUEL TO ENGINE

ELECTRIC MOTOR

FUELIN

FUELTAM

IMPELLER I

BASE

1

I 1

; I I

Fig! ~S-BOO~T - . - PUMP -

I I --

1

\ I I a I

Fig. 19 POWER SUPPLIES TO BOOST PUMP

I I

I I

Redundancy is built in (EASA23 for small aircraft and EASA25 for large aircraft). Normally this is provided by having two boost pumps per tank, each on its own being capable of delivering sufficient fuel to the engines to maintain normal operation. Also the bypass mechanism is available should both pumps fail. Additionally fuel can be used from another tank.

' I

-

NO 2 115V AC BUS /

-

-- - - -- - ---

NO I MAIN AFT BOOST PUMP

I I

NO 2 28V DC BUS NO 1 AFT BOOST PUMP CONTROL

I

NO 1 MAlN AFT BOOST PUMP

- -

0 NO I MAlN AFT BOOST PUMP

NO 1 MAIN AFT RELAY BOOST PUMP SWITCH

SAQ 7

PRESSURE

PUMP

LOW PRESSURE WARNING LIGHT

TO MASTER BOOST PUMP TEST PRESSURE SWITCH

0 U

- - 4 TOMASTER DIM

Fig. 20 PRESSURE SWITCH LOCATION & CIRCUIT DIAGRAM

1 On a multi tanked aircraft which tank(s) should be used first and , how can this be achieved? , I

- -

Crossrfeed System

With &l',systems where there is more than one tank and more than one engine there is ascr ies of pipes connecting-all the boost pumps and engines together. This is called the cross-feed manifold, or sometimes referred to as the 'forward gallery' or 'gallery'.

Fuel from the boost pumps is pumped out of the tanks and into the cross-feer' manifold. It is controlled by a series of valves. To move fuel down to the engine the LP valve or LP cock (sometimes called the spar valve) has to be open. This valve is the last part of the airframe system. From here on the fuel system is considered to be part of the engine.

To move fuel to another engine via the cross-feed manifold, one of the cross- feed valves is opened.

Valves

Can be one of three types: Manually actuated; electrically actuated or remotely actuated via an electric actuator. All must provide the following functions:

* Stop fuel flow in the shut or off position.

- 24 - moodull 1A-1030

* Allow unhindered fuel flow in the 'open' position.

* Provide pressure relief, should the pressure in the LP fuel pipe-work become excessive.

* Indication of valve position, both remotely to the flight deck and physically on the valve itself. (In the case of some light aircraft the fuel control valve is fitted directly to the control knob, which means fuel lines close to, or even in the cockpit. Not allowed in large CAT [Civil Air Transport] aircraft).

Most valves are essentially the same; it is the actuation method that is different. Figure 2 1 shows a typical remotely actuated electrically operated valve. Note the valve body, this is similar on most valves and will be either a butterfly style valve or a 'donut7 type of valve. Both have internal thermal relief, to provide over-pressure relief should (due to thermal expansion) the pipes become over-pressurised. This fuel is ported back into a tank.

I -

\

I

I

,

ADAPTER DETAILS

VALVE BODY

INDICATIONS

Fig. 2 1 REMOTELY ACTUATED ELECTRICALLY OPERATED VALVE

The valve works by a shaft turning in the valve body causing the butterfly to open or close. How that shaft is turned varies from aircraft to aircraft. Common on large aircraft is the one shown; the valve body is buried inside the tank, and it has a n operating link connecting the valve body to a mounting plate. The actuator is connected onto the mount plate via splined couplings and four location screws. (Note the bonding lead).


Indication of valve position is provided in flight-deck.

Figure 22 shows the indication circuitry for a jettison valve which is similar to the one above - with the actuators being identical. It is powered by 28v dc. Moving the switch to open allows the supply to the actuator which starts to run and to the indicator via the closed contacts in the motor and will illuminate the lamp.

The motor runs and the valve shaft and butterfly operate. When it reaches the open position the contacts inside the motor are moved mechanically. This causes the supply to be removed from the open coil of the motor and from the indicator. The motor stops and the indicator goes out.

This type of action is quite common and is called a disagreement indicator. If the indicator stays on, it means the valve has failed. The electrically actuated variant of the valve is the same, except the actuator and the valve body are one unit.

NO 3 28V DC BUS RIGHT FUEL JET,TISON NOZZLE & RIGHT FUEL JETTISON VALVE

TO RIGHT CENTRE RIGHT NOZZLE JETTISON VALVE I VALVE SWITCH

CLOSE B I RlGHT FUEL JETTISON VALVE

W TO MASTER TEST

I I RIGHT NOZZLE VALVE INDICATOR LIGHT

Fig. 22 JETTISON VALVE INDICATION AND POWER CIRCUIT

Manual operated valves (used on small aircraft) use a small diameter [about 1 /8" (3.2mm)l control cable from the fuel cock in the cockpit to the valve body, running through fairleads, pulleys etc.

SAQ 8 What post installation checks are required after a remote fuel system actuator is changed?


Fuel Delivery Pipes (LP)

Fuel pipes are not normally required to handle pressures over 50psi (345kPa) and are usually made from aluminium alloy with the diameter being large enough to cope with the flow rates. This is determined by the amount of fuel the engines need, and typically are about 2%" (64mm) in diameter.

COUPLING NUT COUPLING BODY /

I '

Fig. 23 TYPICAL LP RIGID PIPE COUPLI'NG: s \

, I 1 ' I

I

O-RING SEAL FITTING END i ,

HALVES

RINGS RETAINING RINGS

FLEXIBLE HALF COUPLING FLEXIBLE FULL COUPLING

Fig. 24 TYPICAL LP FLEXIBLE PIPE COUPLINGS

Pipes are thin walled and need care when handling.

Pipes sections are fitted with various connectors. These need to be fuel tight using seals capable of withstanding any flexing that may occur (wings tend to flex considerable during flight and the pipes themselves are not able to take much flexing). Also all pipe-work and couplings must be electrically bonded because of the fluid flow inducing static build-up.

Figure 23 shows a typical example of a rigid coupling and figure 24 shows examples of flexible couplings. There are many types in use and reference should always be made to the AMM for type and fitting instructions. In general:

* Ensure the correct seals are fitted and in the correct way

* Ensure pipes are un-damaged particularly around the seal mating surfaces.

* Torque load correctly.

- -

JC - Ensure correct bonding.

* ' Carry out leak checks after assembly. May need an engine run.

* Some airlines require a duplicate inspection on fuel feed leak checks. -

Pipes in Pressurised and Fire Risk Areas

Where pipes have to pass through pressurised areas (rare) or fire risk areas additional precautions are taken to ensure that any leakage does not get outside the immediate viciniwor to any engine hot sections. These features can include such things as scuppers and channels to direct the spilt fuel overboard. (Engines have a fire-proof bulkhead by regulation). Any couplings near the engine would be enclosed and provided with an overboard drain.

Pipes also run to the back of the aircraft for tail mounted engines, tail mounted fuel tanks (aircraft trimming and fuel transfer) or an APU. Here flexible pipes are used shrouded by a 'normal' aluminium alloy fuel pipe. Any leakage from the flexible pipe is transferred into the shroud. The shroud is ported to a drain mast and any fuel accumulation is drained overboard. A standard 'ramp' check would include checking the mast for any fuel (figure 25). If any is present the leak must be found and rectified.

Water Scavenging

Water can never be prevented from entering the tanks as it is not possible to manufacturer fuel that is completely free of H20. To try to prevent a build-up the water is drained from the bottom of the tanks (water is heavier than fuel and will sink to the bottom) at regular intervals.

- 28 - rnoodull I A-1 034

APU SHROUD -

SHROUD DRAIN LINE

DRAIN LINE APU FUEL SUPPLY

CONNECTION LINE & SHROUD

APU SHROUD DRAIN MAST

CENTRE TANK SHROUD DRAIN LINE CONNECTION

DRAIN MAST OUTLET I 1 APU FUEL SUPPLY LINE CONNECTION , (LOOKING FORWARD)

FUEL +

MOTIVE PORT

n INDUCE PORT

DISCHARGE TO ENGINE

PORT

Fig. 26 EJECTOR PUMP AND SCREEN

With many jet engined aircraft this water is sucked from the bottom of the tank and sent to the engine for burning. The quantity of water is small and makes little difference to engine performance.

Figure 26 shows a typical ejector pump used for drawing off the water. It works on the venturi principal. When fuel is passed through the venturi it speeds u p and the pressure in the venturi decreases (Bernoulli). This allows it to act as a vacuum pump drawing fluid (water) through the induce port. The connection to this port goes to the bottom of the tank, where the water collects and hence it is sucked out and off to the engines.

NO 2 ENGINE

,EJECTOR PUMP MANUAL DEFUEL VALVE :,"t,",","FE

NO 1 IENGINE CENTRE WING

I

- LECTOR PUMP - NO1 MAIN

TANK UMPS NO2MAlN AFT BOOST PUMP

Fig. 27 ENGINE FUEL FEED SYSTEM

The motive force of the fuel is provided by the engine boost pumps via the engine fuel feed system. It is worth noting that the last dregs of centre wing tank fuel often have to be scavenged out and this type of ejector pump might be used here.

Figure 27 shows a Boeing airframe fuel delivery system and figures 28 and 29 show details of the Airbus A 3 2 0 system. Study both systems and work through them to make sure you understand them. There should be no need to commit the details to memory. The CAA, however, will expect you to be able to expIain how a system works; its possible faults and how to rectify them.

FUEL QUANTIW

kgx1OOO- - - I

REFUEL COUPLING

LEFT CTR RIGHT

/MODE SELECT- TEST rELEC POWER-

- , Deluel

TFR

- -- Q, PRESELECTED ACTUAL

VIEW LOOKING FORWARD

Fig. 28 REFUELIDEFUEL POINT - A320

The A 3 2 0 System

The tanks (1 centre tank and 2 wing tanks) are installed in the fuselage centre section and the wings with the refuel panel being situated in the side of the fuselage. The refuel/defuel couplings are situated on the underside of the wings.

The tanks are of the integral type and each tank is made up of an inner and outer cell. Fuel transfers from the outer cell to the inner by gravity via inter cell transfer valves. These are controlled by the low-level sensors in the inner cell.

When the fuel level in either wing inner cell drops to the low-level sensor level, the transfer valves in both the left and right wings will open. There are two low-level sensors in each inner cell. Each sensor will open two valves, one in each wing thus ensuring simultaneous transfer from each outer cell which will maintain trim in the roll axis.

The positions of these valves, open or closed, is signalled to the Electronic Centralised Aircraft Monitoring (ECAM) system fuel page. A n y failure of a valve will be detected by a BITE (Built in Test Equipment) system. A mechanical indicator on the shaft of the valve can be used by the engineer to determine the position of the valve.

,q-- ----- - - - - - - - I

- / / I

R CROSS FEED %

. . m . .,- . '!,AL'!E . .

CZ -I

rri I I I

PUMP, I BOOSTPUMPS I I

I PUMP 1 DEFUEL & TRANSFER 1 7SEQUENCE I I

I I I ' VALVE I VALVES I

I I I I 1 I I I I I I

1 , I I I I I I

REFUEL , I I I I VALVE , I I I

I I

I I , MNGTANK 1 CENTRE TANK WING TANK I - -

-

- - - - -

APU LP Refuel coupling

COUPLING SHUT-OFF

Fig. 29 A320 AIRFRAME FUEL SYSTEM

The amount of unusable fuel in the outer cells is reduced by the fitting of two jet pumps. They use wing booster pump fuel pressure to remove trapped fuel and disperse water from the floor of the outer, and transfer it to the inner cell.

Each wing tank has a vent system, connected by a large diameter duct, to the main vent/surge tank located in the outer ends of the wings, which, in turn vents excess vapour to atmosphere through a NACA duct. The tank vent ducts are capable of allowing fuel to flow in the event of a refuel valve failing to close.


Each outer cell is fitted with an overpressure protector consisting of a carbon disc, which will burst if the tank is subject to extremes of pressure. When this occurs, the protector relieves excess pressure into the inner cell, the pieces of carbon disc being retained by a mesh basket.

A n overpressure protector is also fitted to the inner cell, which relieves overboard.

Any fuel that enters the ventlsurge tank is recovered into the outer cell by a jet pump using wing booster pump pressure. The ventlsurge tank NACA duct is fitted with a flame arrestor, which reduces the risk of a ground fire igniting the tanks. The vent/surge tank is capable of dumping fuel overboard in the event of a refuel valve failing to close.

An overpressure protector is fitted to the ventlsurge tank, which relieves overboard.

, Fuel is supplied to the left engine from the left wing tank by a pair of booster pumps. Fuel is supplied to the right engine . from the right wing @nk by another p&i "6 bos8tei- p-ciTips. These p-ixLps_~pe12&te at ~pFi-o~i i ia te~yy2~gs~ i 1 72kFdj. Centre tank pumps operate at 30bsi (207k~a), therefore the' centre tank will feed first. I

1

I I

I

~f the bodster pumps fail, then thk - - built-in - by-pass valve allbws suction feed by the engine hp pump. One booster pumpiscapable of satisfying-the mBximum

/ demand of one engine. I ; \ \ -

> I 1'; I

I I I A cross-feed valve can be selected, which will)' allow the left tank to supply the right engjne and vice versa. I i

L - L -

l 1 - -1

I -' -

The centrePtG5k-is fitted with two-bomtef pumps, the left pump2supplies the left engine and the right pump the right engine. They do not have a by-pass facility, therefore, in the event of both pumps failing, fuel will not be taken from the centre tank.

By operating the cross-feed valve the left pump can supply the right engine and vice versa.

The centre tank is vented into the left ventlsurge tank through a duct, which will only pass air and vapours, it will not pass fuel. If there is a pressure refuel failure then excess fuel will be passed into the right wing tank inner cell through a pressure relief valve.

An overpressure protector is fitted which relieves into the left wing tank inner cell.

For refuelling/defuelling each wing has a standard refuel coupling positioned beneath the wing leading edge and connected into a common refuel gallery. Only one coupling can be used at a time. Feed pipes are connected to the main gallery and will fill the centre tank and the outer cells of the wing tanks.


Once the outer cell is full, fuel overflows into the inner cell. Refuelling is auto- sequenced:

1st Outer cell wing tank 2nd Inner cell wing tank 3rd Centre tank

Each tank feed pipe has a refuel valve and a diffuser. The refuel valve is controlled at the refuel panel by switches and selected closed by high-level float switches in the respective tank. The diffusers are fitted a t the ends of the pipes inside the tanks and directs the fuel into the tanks with the minimum of turbulence and electro-static build up.

An air inlet valve is fitted in the refuel line near the refuel valve in each wing tank inner cell. Its purpose is to admit air into the refuel gallery to permit the line to drain. Fuel pressure in the line closes the valve. A drain valve allows fuel to drain from the refuel pipe into each wing tank, except when pressure refuelling when the fuel pressure will close the valve.

-- ---

When defuelling, the engine feed pumps are used to move fuei through the engine feed line and then by a transferldefuel valve into the refuel line to discharge a t the refuel coupling.

FUEL; JETTISON SYSTEMS I

I

E A S A ~ ~ states that a fuel jettison system is required if the aircraft does not meet certain climb rate criteria. This has to be done without further endangering the aircraft or it's occupants and needs to be done quickly.

Consider an Zircraft on its take off run. Shortly after V1 (decision speed) it suffers an engine failure. The pilot has to take-off as he/she has insufficient runway left to stop the aircraft in time. Once airborne they will want to land again as quickly as possible. Unfortunately with a full passenger compliment and full cargo holds and full fuel tanks the aircraft is too heavy and it could cause damage to landing gear and structure.

The only option is to dump fuel.

Figure 30 shows a B777 fuel jettison system which is typical of many large jet engined aircraft.

The system utilises the refuel gallery for jettison purposes. The two valves at the ends of the gallery are opened together with the two jettison isolation valves and the fuel is jettisoned at the wing-tips. The valves are motorised open when 'fuel jettison' is selected and the boost pumps in the centre tank, and the jettison pumps pump the fuel overboard.

On some aircraft (the B747 for example) the outer tanks gravity feed into the inner tanks and the inner tanks have jettison pumps.


FUEL JETTISON

OVERRIDE/

FUEL JETTISON PUMP

. . FU'EL JETTISON VALVE JETTISON OUTLET '‘ '

Fig. 30 FUEL J E W I S O N SYSTEM

Control

Jettisoning fuel is fraught with dangexs, and needs to be carefully controlled as iilc airorali is a l c a d y ill i i ~ ~ l g k i - utilr~-wisc it wuu idn ' i LC j c i i i s o n i n g iuei.

I \

' 1 \

I '

SAQ 9 What problems can odcur with the aircraft in the jlettison ' configuration and how can they can be resolved? '

I I I r -. -

, \ , --

\

Selection (and control) of fuel jettison can be! accomplished kither manually or automatically. Both require pilot initiation. Figure 3 1 showy a manual initiated and conti-olledjettison system control pantel,.

, - --

- - -- / - 1

. XmSoN W W I

1

LO t- 10 1 Y" O U I O W - M Y m r l q - q @:@ @:@

maon vrivu U m I I Y * cw- LID*.."

8 8 8 8 <LOW aw CLOY LIDY

nnaon mzru vrsvu

iijj[r.;1gjj aou SO~A~ICU VALM "OY

w ig-- cloy

Fig. 31 FUEL JETTISON CONTROL PANEL

The top row of switches are for the main tank jettison pumps, under that are the selector valves for gravity transfer and finally the bottom row has the switches that actuate the jettison valves.

The aircrew have to select all these switches to start the jettison process. They then have to monitor the contents of the tanks to stop any imbalances and to stop the process at the pre determined level.

In an automatic system the aircrew select a 'fuel remaining' figure on the control panel. They then simply press the arm jettison button and this puts the jettison under the control of a fuel systems control computer, in this case the Fuel Quantity Indicating System (FQIS). It is then automatically controlled.

The pumps run and the valves open and fuel level tanks starts to fall, as this happens the FQIS ensures that the aircraft stays balanced. At the required level the FQIS stops the jettison process.

Figure 32 shows an automated system control panel and the CRT displays. The data comes from the EICAS (Enane Indication and Crew Alerting System) computer on Boeing aircraft (ECAM Electronic Centralised Aircraft Monitor on ~ i r b u k Arcraft).

1 I FUEL JETTISON PANEL

1 FLIGHT DECK FUEL CONTROL PANEL

EICAS DISPLAY

Fig. 32 FUEL JETTISON PANEL AND EICAS DISPLAYS

Jettison Safeguards

Computer systems are usually very reliable, and fuel is rarely jettisoned accidentally. However, errors can happen and the design of a system should allow for safeguards.

To prevent complete jettison of the entire contents, the fuel for jettisoning is taken from a higher point in the tank than the normal fuel for engine use.


This is typically done by having a stack pipe installed on the inlet, meaning the jettison pumps will 'suck air' when the fuel gets down to this level whilst there is still sufficient fuel below the stack pipe available for the engines.

Another, less common, method is to use a float-operated valve that closes and prevents any fuel entering the jettison system below a predetermined level.

Maintenance of the jettison system is not too unlike maintenance of the system as a whole and carried out periodically as per the Maintenance Schedule and will include checks to ensure all the valves and pumps run correctly. In addition a leakage check of the pipe-work and rear gallery would be carried out.

During these checks care needs to be exercised to prevent inadvertent discharge of fuel from the jettison nozzles. To help, the jettison nozzles/system are often protected via air/ground sensing, weight switches, or squat switches which operate when the landing gear shock absorbers compress on landing. This means they cannot be opened with the aircraft on the ground (unless, of course, the, CB's for the weightswitches are pulled).

-

L - I - \ , -- -- -

Jettison of fuel is rare. Even on occasions when you might qeasonably expect the pilot to jettison fuel, hejshei often elects for an overweight landi& instead. Cost and environmental constrdints dictate ,this action. l 1 I

- - l 1 -1

I -- I LONGITUDINAL AERODYNAMIC TRIMMING, FUEL SYSTEM^

I ' I

High speed ~ l i ~ l i t ' I

I 1 I I 1

(Thispart is not applicable for those s tdlyi ig for their Piston ~ n g i n e d Aeroplanes licence (module 11B):Also ~oncorde is used as the example as this is, currently, the only supersonic civil aircraft to have been in service.)

A s the aircraft approaches the speed of sound (usually with a swept wing aircraft) the centre of pressure moves rearwards and as such the lift force opposing weight will put the aircraft equilibrium out of balance. This new moment causes the nose of the aircraft to pitch down (called Mach tuck).

SAQ 10 How is this trim up-set of the aircraft normally countered? C a n you think when this normal method might not be possible? This Q is not applicable for those doing module 1 1B (piston engined aircraft).

An alternative is to move weight (mass) within the aircraft. The most obvious mass to move is fuel (there is plenty of it and it is comparatively easy to move). A s it is moved longitudinally the Centre of Gravity (C of G) and hence the correct longitudinal trim can be obtained (figure 32).

FUEL TRANSFER AFT FUEL TRANSFER FORWARD

Fig. 33 MACH TRIM - CONCORDE

Figure 33'shows an overview of Concorde Mach trim. The system moves significant quantities of fuel from the forward trim tanks to the rear trim tanks as the aircraft goes through the sound barrier and accelerates to Mach 2. (As a matter of interest the Centre 04 Pressure actually shifts about 6 feet (1.8rn), causing a significant pitch moment change). Moving fuel from the forward trim tanks: rearwards counteracts this.

-

I

Concorde holds about 33 tonnes (33.5 tons) of fuel in its trim tanks. As the aircraft decelerates from supersonic to subsonic speeds the fuel is transferred from the rear trim tanks to the forward trim tanks.

Subsonic Flight

Some subsonic aircraft now have fuel tanks in the tail and the fuel is used for normal longitudinal trimming purposes.

INDICATING SYSTEMS

Fuel indication systems can be manual, mechanical or electrical/electronic and are usually associated with fuel quantity (volume and mass) and fuel flow rate indications (if fitted are on the engine side of the fuel system).

Manual Indication Methods

Manual indication methods were the first type of indication, starting with dipsticks similar to the dipstick fitted to the automobile engine for checking the oil contents. Variations of the dipstick include the dropstick and the dripstick.


Dipsticks are fitted to some older aircraft and are used by maintenance engineers to give a reliable indication of the contents of each fuel tank.

To use, unscrew the dipstick and remove complete with sealing cap if fitted, clean with a lint free clean cloth, push back into the tank so that it rests on the top screw cover, pull out and note on the dipstick where the wetted section ends. At this point the graduations on the dipstick will indicate the amount of fuel in the tank. Replace the dipstick, screw into position and lock as per the AMM.

Figure 34 shows a simplified layout for the dipstick, dropstick and dripstick systems.

The dropstick is released from the bottom of the tank, the same as the dripstick, and allowed to fall under its own weight. It will come to rest when the magnet in the top of the stick reaches the iron core of the float. The graduations protruding from the bottom of the tank give an indication of the tank contents.

Fig. 34 MANUAL FUEL CONTENTS MEASUREMENT

Figures 35 and 36 show two methods of manual stick measurement. Both are unscrewed from their mounting and lowered out of the lower wing surface. In the case of the sight glass the prism causes the viewing port to turn black as the prism is immersed in fuel. It is lowered slowly whilst loolang at the prism through the eyepiece - when it turns black note the graduation showing level with the tank fitting - this indicates the tank contents.

Note. Wear protective glasses when doing this.

rnoodul i~ i i?ods

QUARTZ PRISM

SHUT-OFF CAP USED FOR UNIT REMOVAL WITH FUEL IN THE TANK

TANK MOUNTING

AND LOCK

GRADUATIONS

- Fig. 35 SIGHT GLASS TYPE 'DROPSTICK'

The Dripstick is lowered slowly until the open top allows fuel to spill out of the hole at the bottom of the stick.' It is then a simple case of measuring off against the scale and that gives the quantity in the tank.

' I

A development from this is the dripless stick or dropstick (figure 37). In this system there is a ferrous material target attached to the top of the stick. The stick is free to move up and down inside a fuel free tube. The float, which is in the tankyfloats on top of the fuel and has a magnet inside. When required the stick is released and allowed to d r o p u t of the bottom wing skin. The target and the magnet coming together will stop it at the fuel level position.

CALIBRATED HOLLOW ROD\^)

Fig. 36 DRIPSTICK

The reading can then be read off the scale. Usually this reading is given in inches and is converted into a fuel quantity by looking up tables in the AMM and held on the aircraft flight-deck. Care needs to be taken as the readings maybe taken back to front.

FLOAT

MAGNETIC FLOAT

ARMATURE 'i

MECHANICAL INSTRUMENTS

Figure 38 shows a basic mechanical system. It relies on a float (sealed cork or a metal canister) located in the tank moving, directly, or through gears and levers, a pointer within an instrument casing. These are found on some light aircraft. The manufacturers try to locate the dial or sight glass where it can be seen by both the pilot and the person carrying out the refuelling operation.

In some cases, on small older aircraft, a manometer type gauge might be fitted.

GEARING

PUSHlPU LL ROD

Fig. 38 FLOAT TYPE INDICATOR

ELECTRICAL INSTRUMENT SYSTEMS

Fuel level can be measured by any of the following systems: ' - -

' 1. A moving coil instrument. 2. A dc ratiometer. 3. An ac ratiometer. 4. A Desynn system. 5.1 Asynchrosystem.' -

6. A capacitive system. 7., An ultra sonic system.

With syktems 1 to 5 a float will be used to move up and down with the level of the liquid. When the float moves it will move a variable resister for systems 1, 2 and 4;avariable inductor within ac coils for system 3 and a transmitter synchro on system 5. These systems are all described in the book in this series entitled Instrument Systems so will not be dealt with further here.

Systems 1 to 5 are rarely used for fuel level indication.

With systems 6 and 7 there are no moving parts and the capacitive system is the most widely used.

These two systems will be described in this book.

ELECTRICAL/ ELECTRONIC SYSTEMS

THE CAPACITIVE SYSTEM

If two metal plates are placed close together (but not touching) with a dielectric (air or fuel - or any other material for that matter) in between, a capacitor is formed.

The value of a capacitor is given by the formula:

Where A is the surface area of the metal plates and d is the distance between them. These two values are fixed by the manufacturer of the unit. The value of E (lower case Greek letter epsilon) is altered by changing the dielectric constant, and that happens as the level of the fuel is changed. This is because fuel and air have different values of E (permittivity).

DIELECTRIC INNER

PLATE : ; ' I ' I

Fig. 39 SIMPLE CAPACITIVE TANK UNIT - -

I - . I -- 1

As the fuel level rises in the tank, air is displaced by fuel and the dielectric changes to increase the capacitanke of the w i t . Thus the change in the capaciiance is related to fuel quptity. As th& fuel level goes lddwn so the capacitance goes down. , - - - , /

/ , ,-- -

capacitive tank units are normally cdbrated at manufacture and as such require no on-wing calibration.

- a . I . +

CAPACITOR j TANK UNIT i SIGNAL

: j PROCESSOR FUEL GAUGE

I .

i ; I . I . I I ;....; ....

Fig. 40 SIMPLIFIED CAPACITIVE FUEL GAUGING SYSTEM

Figure 39 shows a simple capacitive circuit with one tank unit. There are normally several tank units per tank, wired in parallel and supplied with a small dc voltage. This voltage sets up a potential difference across the plates between which the capacitance is measured.


Figure 41 shows how averaging is achieved. Throughout the tank there will be several capacitive units and their values will be added and averaged. This allows for the possibility of the aircraft being in an attitude which causes the fuel to move to one end of the tank. If, say the tank is half full and this happens, the tank units at one end will indicate a full tank and units at the other will indicate an empty tank - averaged, the values show a half-full tank. The size of the tank and the shape, dihedral angle etc determines how many tank units are fitted.

Figure 42 shows a typical capacitive tank unit (with a compensating unit), sometimes referred to as a stack or tank unit.

TANK UNITS 2 2 4 . A

/ FUEL TANK

' FUEL LEVE

I TANK AT AN ANGLE SO CAPACITANCE OF CAPACITOR FUEL TANK LEVEL -CAPACITORS 1 & 2 EQUAL 1 INCREASES WHILST THAT OF 2 DECREASES BY THE

I SAME AMOUNT

Fig. 41 AVERAGING

- - OUTERTUBE

[GAUGING UNIT]

EFERENCE UNIT UC-

Fig. 42 CAPACITIVE TANK UNITS

The outer plate is called the stillwell and allows the fuel to be moderately static inside regardless of aircraft turbulence, vibration etc.


The cables to and from the tank units are low voltage cables and they are all collected together and passed to a single connection point. This minimises the risk of fuel leakage.

SAQ 11 What effect on the dielectric constant will a change in fuel density (or specific gravity) have?

There is a need to compensate for change in fuel density and there are a number of ways of achieving this. The simplest form of compensator is a small capacitive unit (called a compensator) that sits at the bottom of the tank, often close to the sump drain.

Unless the tank is completely drained it will always be totally immersed in fuel. This means that any change in capacitance of this unit is not due to a change in the fuel level, but due to a change in the density of the fuel. This signal is then used to trim the main tank units to allow for the slight changes in the fuel relative density. This will allowthe fuel contents to be given in terms of mass -- --

\ 7- -

arid rlui just vuiuiile, ir i ~ i kgs aiiiilul~lccess~uiiy irl i i i~cs ui g d i ~ ~ l s . I I

i : \ I

\

,

PRESSURE - --

REFUELLING

CENTRE TANK

FUEL NO 1 TANK QUANTITY GAUGES

TANK UNITS (24)

Fig. 43 EXAMPLE OF TANK GAUGING & COMPONENT LOCATION

Figure 43 shows the location of the various units in an aircraft tank system and figure 44 shows the details of the construction of the unit with a typical range of units available to the aircraft designer.

We shall now consider a 'bridge type' capacitive fuel quantity indicating system. Once again it is important for you to know the system on the aircraft you are currently working on, but it is also necessary for you to understand the principles of this system.

The system is used to measure the mass of usable fuel on the aircraft. The basic principle is shown in figure 45.

SIZE RANGE

Fig. 44 DIFFERENT SIZES OF TANK UNITS

With reference to figure 45 note the following:

(i) Power supply is 1 15v 400Hz. (ii) The many tank units are represented by one capacitor in one leg of

the bridge system and one compensating capacitor in the other. (iii) In the other leg of the bridge is also a fixed capacitor as part of a n

amplifier circuit. (iv) Across the centre of the bridge is a resistor 'R', any voltage

developed across this is fed into the amplifier. (v) The cockpit indicator is driven by a motor in the indicator. It is a

two phase motor with its control phase signal coming from the amplifier. The motor drives the indicator and a balance potentiometer (a variable resistor as part of a negative feed-back circuit).

- 46 - moodull l A-105;!

TANK UNITS I - - - - - - - - - I

POTENTIOMETER

FIXED CAPACITOR

INDICATOR CAPACITOR

2-PHASE MOTOR STAGE

-

- FAIL - --

1 I I - SAFE 1

REFERENCE CIRCUIT PHASE

I ~1 I ,

I \ 1

Systey operation - Balanced condition (fuel quantity steady), I '

I 1. - The current through the-top leg of the bridge IS is the same as the

- -

- -current through the bottom leg of the bridge &; 2. The two currents pass in opposite directions through the resistor

R. As they are equal then no currentflows through R - no voltage is developed across R, and there is no input to the amplifier. The indication therefore remains steady.

Unbalanced Condition (fuel quantity decreases)

1. The capacitance of the tank units will fall.

2. The current in the top leg of the bridge will fall, because as the capacitance has fallen then Xc = 1

{ 2nfc then Xc must have risen, this capacitive reactance (resistance if you like) is the opposition to current flow, therefore current must fall (I = v/ x,) .

3. Nothing at this time has happened to the current in the bottom leg.

4. The opposing currents through R are no longer equal and IB is greater than IS. The current flowing through R is therefore IB - IS, and the voltage drop across R is (IB - 1s)R.

5. This voltage drop across R is fed to the amplifier, where it is amplified and fed onto a discrimination stage.

6. It is important to note that the amplifier must pass the correct sense (up or down) signal to the control phase of the motor, to ensure correct direction of rotation. The signal that it receives could be a 'fuel increase' or 'fuel decrease' signal. The discrimination stage looks at the phase relationship of the signal from the bridge circuit and then it knows whether it is a fuel increase or a fuel decrease signal.

7 . The signal is fed to the control phase of the motor. The reference nLncn ;n farl nff onnther ceor\nrl=rrr nf the rn a i m c ~ ~ ~ ~ l ~ trzqc,fc-mel.

, - ylluvb I" IbU "I* -I" C I I V I "UUVLIUULJ ..A LA-- A a A C Y a a U -.r

The motor drives th&indicator down scale and also the balance potentiometer towards 'empty'.

8. The balance potentiometer controls the voltage across the bottom leg of the bridge. AS the potentiometer wiper arrn is moving towards empty, the voltage across the bottom leg of the bridge is

I I falling, therefore IB must be falling. This is a negative feed-back I signal to cancel the input to the motor.

9. - Eventually IB will fall to the value of IS and when this happens, - .- .

there will be no voltage drop across R, no input to the amplifier, the motor stops and the indicator shows the new lower fuel reading. At this point the bridge circuit is said to be balanced.

Read through this again carefully, and work through what happens if there is fuel increase in the tank. A s far as the equations are concerned there is no need to remember them. There is no need to commit the details of the drawing to memory either but you should be aware of the principles.

In general the principle is that, as fuel quantity changes, the bridge is unbalanced and the error signal is fed to an amplifier and motor which re- balances the bridge whilst at the same time moving the flight deck indicator.

A similar situation occurs when the density of the fuel changes. It will affect all the tank units in both arms of the bridge and change all their reactance's (resistance's) so the bridge remains balanced, but one arm has an extra capacitance unit in the form of the compensator. This means that it will upset the balance and start a similar train of events to when there is a fuel level change.


Both the tank unit and the compensator signals are fed through cabling to processor units. These compensate the signals and amplify them. For servo type instruments the signal is a voltage sent direct to the flight deck instrument. For CRT/flat screen display systems the signal is sent to display drivers or symbol generators to provide for flight deck display.

Note. When testing a capacitance type system a capacitive bridge type test set is used and connected in parallel to the tank units to test points in the system.

An Electronic System

Most modern aircraft have these systems fitted although standby instruments, similar to the type described above, are usually fitted for use if the normal system fails. Electronic components are often housed around the display indicators themselves.

In a typical system (figure 46) the tank unit signals, along with the compensator signals, as well as dens&, temperature and water detection (possibly), are fed to a computer, called,(on Boeing aircraft) .a rue1 Quantity ~rocessilig Unit (FQPU) . \ I

I

I i The signals are sent as either discrete or m!alogue signals. Qnce at the I

/ computer they are converted into digit&d,dta by the analo@e t o digital convertei (AID converter), theniprocessed~through the CPU (dentrd Processing

1 i 1 '\ Unit). i \ '\ 1 ! "

FQPU

Fig. 46 EXAMPLE - FUEL CONTENTS SYSTEM - BOEING

The computer will calculate the fuel level for each tank and the total fuel on board and transmit that data in Binary Coded Decimal (BCD) word format along either ARINC 429 or 629 serial data buses.

The signal is sent to the refuel panel (IRP) for display during refuelling. It is also transmitted to the display drivers (called various things such as Symbol Generators but in figure 45 they are called AIMS - A Boeing term meaning Aircraft Information Management System - and then onwards to the CRT screens in front of the pilots.

It is displayed on the top centre screen as a total quantity. (Most 'glass cockpit' displays have 2 screens in front of each pilot for HSI and AD1 displays, or similar, dealing with aircraft attitude, direction etc, and two centre screens showing systems status [hydraulic, fuel, elect, engines etc].

Boeing call this screen system EICAS and Airbus use the term ECAM.

WIM P I S WE* DCrl tm *mt ss m = a g ~ s , rn n( ss D. -sat

O I I m CIISS 011 w mass

FUEL SYNOPTIC DISPLAY FUEL MANAGEMENT MAINTENANCE PAGE

Fig. 47 FUEL SYSTEMS DISPLAY - BOEING

Fig. 48 A DENSIOMETER

'50 - moodull 1 ~ ~ 1 6 5 6

A button can be selected to bring up more detailed information of the fuel system. See figure 46 and the books in this series entitled On Board Maintenance Systems.

Provided a digital system is being used, there is another way of measuring relative density of the fuel. It uses a densitometer (figure 48).

The densitometer will accurately measure the density of the fuel in the tank. Again located in the bottom of the tank, to ensure that it is always immersed in fuel. It is composed of a nickel alloy, wire wrapped spool, and an excitation coil, a sensing coil and trim resistors all contained in a vented housing. Different sized connectors help 'Murphy' proof the electrical connection.

Fuel density is determined by causing the spool to vibrate at its resonant frequency and measuring any vibration changes. A control unit (the FQPU on Boeing aircraft] sends a signal to the excitation coil and this causes the spool to vibrate. Fuel is allowed to enter the housing through the vent holes.

Dense-fuel will slow the spool-down more and this 'slowing'or change in vibrations frequency is felt by'the-sensing,coil. This is fedback-to-the control unit (FQPU) for processing and used to calculate the mass of the fuel.

l 1 The densitorneter is accurate d d can woik'out the relative dehsity oh the fuel to two decimal places. It also has anoth'er function. If the relative density starts to climb towards 1 .O, it can be assumed ,that water has collecting &,the bottom

I \ of the t d k . I-

' \

Perhaps the ejector pump has b10,cked or i s leaking. A warnfng is sent to the flight deck - and would be displayed to the flight crew as a 'status' message.

/ - - ,'

AN ULTRASONIC SYSTEM

This system senses fuel tank quantity by ultrasonic fuel probes. These and other signals are processed by a Signal Processing Unit which is controlled by a CPU. This unit will send fuel quantity information via a n ARINC data bus to flight deck displays and other systems requiring the information.

The ultrasonic fuel probe consists of two main assemblies; a transducer housing and a still tube. The piezoelectric transducer is mounted inside the shroud and is located against a reference face at the bottom of the still tube. The still tube acts as a guide for the acoustic signal whilst shielding against extraneous acoustic noise and minimising the effects of fuel turbulence.

Fuel Gauging Operation

The piezoelectric transducer in the fuel probe is excited by an electrical pulse generated by the signal conditioning unit. The transducer emits an ultrasound wave upwards into the fuel inside the still tube.

The sound wave travels through the fuel until it meets the fuel surface where it is reflected back down the still tube to the transducer. The return signal is sensed by the transducer and an electrical current generated which is received by the signal conditioning unit.

The 'time of flight' of the ultrasonic wave from the transducer up to the fuel surface and back is a function of the immersed height of the fuel probe and the propagation velocity in fuel. By means of a reference flange in the still tube at a reference height from the transducer a second reflected signal is produced.

Assuming that the propagation velocity is the same in both the calibration wavefront path and the surface wavefront path, then the height of the surface may be calculated directly without reference to the propagation velocity.

Thus the fuel level is determined ratiometrically. Fuel height signals from the numerous probes within the tank are used to calculate the volume of fuel contained within that tank from stored tank shape data. Volumetric quantity is then converted to mass using the signal from the densitometer (working on thp vibration principle). Fuel quantity in terms of mass is then sent to cockpit displays ,&d aircraft systems. - --

I

The datLr detector consists of the basic transducer housing assembly minus a still tAbk and is mounted at the lowest point in the fuel tank where water will tend do collect.

I ' ---

The t{mperature sensor is a platinum resistance thermometer element and is located within the terminal block assembly of a fuel probe. Normally one such sensor is provided per tank.

/ \ ARlNC , , CENTRAL SIGNAL < > 629& < > PROCESSOR - PROCESSING

429 \ / UNIT (CPU) UNIT (SPU) INPUTIOUTPUT TOIFROM AIRCRAFT SYSTEMS

A

FUEL TANK

Fig. 49 ULTRASONIC SYSTEM - GENERAL LAYOUT


The microprocessor-based control unit provides:

(a) Drives for the in-tank sensors. (b) Conditioning of sensor outputs. (c) Computation of fuel quantity. (d) Communication with flight deck displays and other aircraft

systems. (e) Fuel distribution management. (4 Fuelling/ defuelling control. (g) Built-in-test (BIT).

The processor consists of:

(a) Signal conditioning modules. (b) CPU modules. (c) Digital 1 / 0 modules. (d) Power supply module.

Fig. 50 ULTRASONIC FUEL GAUGING - PRINCIPLE

Fuel surface attitude information can be computed from fuel height data of any three probes in any one tank. This allows compensation to be applied to fuel volume calculations based on stored tank geometrical data and compensation for aircraft attitude variations.

In the event of densitometer failure, fuel density can be inferred from the empirical relationship between fuel density and propagation velocity. This calculation uses the time of flight to the reference flange from a number of fuel probes. Fuel stratification effects can also be identified by comparison of the propagation time for a number of fuel probes at differing heights within the tan2;'.

That concludes this book on EASA module 11 fuel systems. To consolidate your knowledge try the following exercises:

Name the three types of storage system. List the precautions when fitting bag type tanks. Where are integral tanks fitted and how are they fabricated? What is the purpose of the 'permit to work'? What are the two types of refuelling systems? What is the purpose of the venting system? Explain how and when you might use a drain valve? List the precautions associated with a refuelldefuel operation. What is the recommended maximum flow rate for a pressure refuel and over wing fuelling systems? What type of pump is a boost pump and why use this particular type? What are the precautions associated with connecting fuel pipes? Where are pipes shrouded and why? How is water scavenging carried out? Explain how a jettison system works and why one is needed. How is the jettison process controlled? What safeguards are built in to prevent total fuel jettison? What is a fuel stability system? What manual methods of indication are used - by themselves, and with more advanced systems? Where are direct reading mechanical gauges usually placed and why? Explain briefly how a capacitive tank unit or stack works? How are the signals fed to the EICASIECAM screens in the fully digital system? What functions do densitometers perform?

S t u d y the AMM on fuel systems for the aircraft type that you are currently working on. Note any similarities/differences to the systems described in this book. If working with ATA configured manuals look in ATA chapter 28-40 and look at page block 1 - 100. Also check out chapters 6 and 12 also.

Answers to all the above (except number 23) are to be found in the text.


SAQ 1. An aircraft fuel system constitutes: Storage, Refuel and defuel systems, engine fuel feed, jettison and indication systems. The start of the fuel system is where the bowser/tanker connects onto the aircraft, and the aircraft fuel system ends at the LP - Low Pressure cock, or LP SOV - Shut-Off-Valve.

SAQ 2. Flexible tanks have the advantage that they can be made to fit any cavity shape. They are fairly easy to replace and as such, leaks in the tanks can be easily found and fixed. Unfortunately they are fragile and prone to deterioration. They add weight to the aircraft. Leaks in the bags may not be detected until there is a build up of fuel in the structure. They are more crash proof than the other type of tanks as they can take some distortion without rupture.

SAQ 3 . Answer in the text.

SAQ 4. ATA chapter 12 gives details of servicing. You may, however, have to look in ATA chapter 6 to find details of the refuel panel location.

SAQ 5. Carry out operations as though you were to defuel the aircraft. Without the bowser connected and hence the NRV7s closed. The fuel will travel from the forward gallery onto the rear gallery. Now we can select a refuel valve open from the refuel panel and then transfer fuel from one tank to another.

SAQ 16.-No-is the answer. Due,to the fact that there is only one-pump in the line if this were to seize then fuel could not get'to the engine. The usual way of allowing for this is to provide a byPass pipe,'with check v q . under normal operatiod the check valve stays closed by fukl pump pressuie. If that pump fails tKe engine will 'suck7 and l-ience open the check valve y d allow 'fuel through. - -

I

I - . \ i I --

SAQ 7.' If; there were a Centre Wing Tank (CWT) this would be ernptred first so that the weight isieduced from itfie fuselage section. Also, kkebing weight

I within the wing &d hence downward force oh the wings, counter balances the upward force,of lift. (This reduces the bpding moment of thje %ng which can be dcsigna_as a lighter strucpre). This is achieved in one of two ways. Boeing uses higher pressure boost pumpsin the CWT, causing it to empty first. Airbus generally uses a single point sensor that allows a float to open when the CWT is empty.

SAQ 8. This follows a similar pattern to the post-fit tests. Check for security of the component; ensure it is correctly orientated and the screws are tight and wire-locked if required. Don't forget the electrical plug, usually a standard Arnphenol type connector and will have witness marks to show correct locking. A bonding check will be required in common with all fuel system component changes. Finally it will need a functional check iaw the AMM.

Carry out the specific AMM checks for replacement that are likely to include open and close checking against the manual indicator. Carry out flow (and no- flow) checks with a number of boost pumps running - these checks ensure the valve opens fully, closes fully and that the thermal relief hasn't failed.

Your company may require a duplicate inspection and all work must be recorded in the aircraft log book/work record cards together with part numbers, any test data etc. All work must be signed for.

SAQ 9. Some of the problems that could occur are the fuel coming into contact with hot exhausts. This is covered by fitting the jettison pipes usually at the extremities of the wings away from any heat sources.

Another problem is fuel imbalance. If one pump were running a t a higher pressure, that tank would empty quicker than the other. Also if the pumps didn't stop for some reason the entire contents could be jettisoned.

SAQ 10. The stabiliser (tailplane) trim mechanism to move the stabiliser nose down (decreased angle of incidence) causing an opposing nose u p pitch. This will counter the natural nose down tendency of Mach tuck. This is the usual way to counter Centre of Pressure changes.

SAQ 11. If the relative density (or specific gravity) of the fuel is lowered the dielectric constant of the fuel is lowered. So applying the capacitive equation we can see that the capacitance will also be lowered.


CONTENTS

Page

Principles Hydraulic fluids Hydraulic systems Power circuits Consumer circuits Hydraulic components Emergency/ standby systems The power pack Hydraulic seals Pipelines Flight deck indications Maintenance

PRINCIPLES

The term 'hydraulics' is used to describe methods of transmitting power through pipes and control devices, using liquid as the operating medium. Hydraulic systems are used in preference to mechanical or electrical systems for a number of reasons, amongst which are ease of application of force, ability to increase the applied force a s necessary, ease of routing of pipelines, and elimination of backlash between components. The most important reason, however, is that hydraulic systems have a good powerlweight ratio.

Liquids are considered to be incompressible - at least up to pressures of 3,000 to 4,000psi (21MPa to 27MPa) - not strictly true but this assumption is generally made. The higher the pressure the more the fluid compresses and at very high pressures (say 20,000psi) then a fluid behaves very much like a gas during compression. Many hydraulic systems have a working pressure up to about 3,000psi so the transmission of fluid power down a pipeline can be achieved with very little power loss.

-- -

However, liquids will expand or contract as a result of temperature changes, and a thermal relief valve is necessary to prevent damage from excessive pressure in any closed circuit which may be subjected to changes of temperature.

Before we start looking a t hydraulic systems you must have a knowledge of some of the rules and laws which govern the behaviour of fluids under pressure.

I r

Pressure. Defined as force per unit area or I

Where P = Pressure F = Force A = Area

Units-Imperial - Pressure = Pounds per square inch (psi) Force = Pounds force Area = Square inches

Metric (SI) - Pressure = The Pascal (Pa). 1 Newton per square meter = 1Pa Force = Newton Area = rn2

Note. The Pa is the Pascal (named after Blaise Pascal 1623-1662). This is a very small unit with nearly 7000Pa to just lpsi.

Pascal's Law states that fluid held under pressure in a container exerts pressure equally, instantly, and a t right angles to all surfaces of the container without loss.

Brarnah's Press. Sometimes called a Hydrostatic Press, it is used to magnify a force. It is a hydraulic press consisting of two cylinders - one larger in diameter than the other. Sealed pistons are allowed to move within each cylinder.

I n general a small force on the small piston will create a large force on the large piston. But in pushing the larger piston up the smaller one will have to move through a larger distance.

During operation the parameters that are common for both cylinders are:

(a) The volume of displaced fluid and, (b) The pressure.

In other words the pressure is the same in both cylinders and the total volume displaced from one cylinder is the same as that received by the other.

-- - - I SMATL- pp

\ i PISTON - - . I

I

I LIQUID LIQUID , i '

I I

I ) I ' I

-4

- Fig. 1 THE BRAMAH PRESS

Example

Take the approximate sizes of an aircraft hydraulic lifting jack. The piston that pumps the jack up is about 0.5 inches in diameter and the jack body piston being about 3 inches in diameter.

QUESTION: Given the above diameters of the pistons can you work out their areas? (5 mins)

ANSWER: Area of ----- small piston -------------- large piston nd2 - - nd2 4 4

'\\

= n x 0.5 x 0.5 = n x 3 x 3 4 4

= 0.2 in2 = 7 in2

- 2 -

rnoodull l A-1 065

LIFTING JACK

SMALL PISTON

LARGE PISTON

VALVES \ PRESSURERELEASE VALVE

Fig. 2 HYDRAULIC LIFTING JACK - SCHEMATIC

If a person can push down on the small piston with a 90 lb force --- [using most of his jher own body weight anci not using a handie (so no mechanic4 advantage is obtained) then what weight will the jack be able to lift?

In general calculate the pressure in the small cylinder then use this pressure in the large cylinder to calculate it's force.

-- -

Pressure equals force per unit area. I

I 1 pressure = force p '=

area A -

- so - Force = Pressure x area F = P x A

SMALL PISTON 0.5" (12.7mm) DIA

LARGE PISTON 3" (76.2rnrn) DIA

Fig. 3 PISTONS - RELATIVE SIZES

1. Pressure in small cylinder = P = F = 90 = 450psi A 0.2

2. This is also the pressure in the large cylinder. So the force on the large piston equals = F = P x A = 450 x 7 = 3 1501bf.

rnoodull ?A-I 066

The jack will lift 3 1501bs - nearly 1 % (imperial) tons (about 1 % tonnes) - the size of a large family saloon car. With a lever on the small piston (as with all lifting jacks - figure 2) for the operator to push on, the jack will lift ten times this load (the lever giving the operator a mechanical advantage).

The same calculation using SI units.

Diameter of small piston = 13mm approx. (0.0 13m) Diameter of large piston = 76mm approx. (0.076m)

Areas: SMALL PISTON LARGE PISTON

&2 & 4 4

= n x 0.013 x 0.013 - - n: x 0.076 x 0.076 4 4

= 0.0001 m2 - - 0.004 m2 r -- -- - -- - J ,

- -. -. The paessure in the system with 4 901b force on the small piston is: (901b prce is approximately equ'al to 40 1 Newtons).

i I , ' I

- - - 401 1 - 4,010 800 N/m I - - 3

1 A ~; = 0.0001 '

i I

\ i

\ or 4,010,000 Pa 1 1 I I I

~ : I I I

I , , oi- ; 4 MPa pressure approx. - 1 \ --- ,' - -

t -

~ h i b ~ r e s s u r e i l l act on the largeFpiston Ad the force produced = pressure x area.

F = P x A = 4,010,000 x 0.004 = 16040 Newtons = 16.04kN

All the above calculations assume that there are no losses in the system and no friction to overcome in the seals.

QUESTION: Is the above statement true?

ANSWER: No - there are losses in all systems. But for general calculations it is assumed that it is true and the answer works out quite well.

QUESTION: Can you work out how far the large piston will move up when the small piston is pushed down?

If your answer is YES then have a go assuming the small piston is moved down 4" (10 mins).

- 4 -

rnoodull l A-1067

If your answer is NO then refer to the answer below.

ANSWER: Remember the basic rule - the volume displacement from one cylinder is the same volume displaced to or from the other.

Volume moved out of small cylinder = & x h 4

where h = height of piston movement.

Sovolume = n x 0 . 5 x 0 . 5 x 4 4

= 0.8in3 approximately

This amount of fluid is displaced to the big cylinder.

Volume = V = Ah -

I

Where V = volume, h = movement height of large & A = piston area of large piston.

So the amount of big piston movement is (h):

h = y I -- , A = 0.8

7 = 0.1" I ~

It doesn't move far does it? - bearing in mind the movement of the small piston is 4in. When you jack an aircraft just-note the differences in the movements of the jack piston for each stroke of the pumping piston.

With reference to figure 2. When the pump piston is moved u p it draws fluid in from the reservoir via the one-way valve. When it is pushed in it pushes the fluid out to the jack via the other one-way valve. To lower the jack the release valve is opened which allows fluid in the jack cylinder to return to the reservoir.

HYDRAULIC FLUIDS

Almost any type of fluid can be used in a hydraulic system, but the special requirements of aircraft systems have resulted in the use of vegetable, mineral and synthetic based oils. They must meet the requirements laid down by the regulatory authorities (EASA 25 for large aircraft, 23 for small aircraft, 27 for small rotorcraft and 29 for large rotorcraft).

- 5 -

rnoodull I A-1 068

An ideal hydraulic fluid would have the following properties:

(a) Be a good lubricant. (b) Have a low viscosity to minimise friction in pipelines and to

provide high-speed operation of motors and pumps. (c) Be anti-corrosive. (d) Have a wide operating temperature range. (e) Be non-inflammable - not true of all aircraft hydraulic fluids. ( Be user friendly (non-toxic etc) - not true of some hydraulic fluids. (g) Be inexpensive - all are costly, some very much so.

QUESTION: What does 'viscosity' mean? (5 mins)

ANSWER: It is the resistance to flow of a fluid. The higher the resistance, the higher the viscosity and the more energy the fluid requires to be pumped around the system. Viscosity may change with temperature

c (called 'viscosity index'))- usually the higher the temperature the , I - .

I 1 lower the viscosity. , I \

I ' I

I I , I

Fluids are coloured which helps recognition, but in general fluids should only be 1 used if they are from an approvrd supplier; in sealed containers; to the correct

specification - as laid down in theAMM (Aircraft Maintenance Manual). I I I

Fluidd tb different specification: must never be mixed. Fluids to the same specification, but produced by different manufacturers, may be mixed when perrnittdd - i n the 'AMM. -

i -_ --- - - -- /

Use of a fluid, which is not approved for a particular system, may result in rapid deterioration of seals, hoses and other non-metallic parts. It may also cause high wear rates and allow sludge to form.

Types of Fluids

There are many types of fluid on the market and it is most important that the correct specification fluid is used as stated in the AMM. Below are listed the more common types:

LOCKHEED 22. (to MIL SPEC H-7644). Vegetable based and almost colourless - but a slight brown/yellow hue. Pungent smell. Used with natural rubber seals and hoses. Used in some braking systems but not often found in modern hydraulic power systems.

rnoodull l A-1 069

DTD 585. (to MIL SPEC H-5606). Mineral based and coloured red. Uses synthetic rubber seals and hoses. Used in hydraulic systems and landing gear shock absorber struts. Excellent lubrication and ant-corrosive qualities, but flammable.

SKYDROL. (to MIL SPEC H-8446). Phosphate ester based (synthetic) and may be green, purple or amber in colour. Used with butyl rubber, ethylene propylene, or Teflon seals and hoses.

Widely used on modern aircraft because of its fire resistance - though high- pressure spray is combustible. Temperature operating range is between - 65°F and + 225°F (- 54°C to +107"C). This fluid requires care in handling as it is an irritant to skin and eyes.

Avoid contact with the body and avoid inhaling the fumes. Always use a barrier cream and protective clothing such as fluid resistant gloves, goggles etc. It will absorb atmospheric moisture and attack most plastics and paints (though not epoxy and polyurethane based paints). May become acidic if overheated.

-- - -

HYDRAULIC SYSTEMS I

Hydraulic systems are used to operate such services as landing gear, wheel brakes, powered flying controls, windscreen wipers, flaps, spoilers etc. Each of these services has its own hydraulicEirCuit within the hydrauliC system. These circuits are usually connected tq supply and return lines running to and from the supply circuit. Thus the complete system is made up of a supply circuit connected to various service circuits. I

O n many *craft more than one hydraulic system is provided and these may be interconnected. On most large aircraft three (or even four) independent systems are used, each with it's own supply pumps, reservoirs, pressure and return lines.

Each system supplies it's own services with the more important services receiving supplies from more than one system.

On most large aircraft the Powered Flying Control Units (PFCUs) for example, receive three independent supplies (the A380 is different in that each PFCU is a self contained electrically powered hydraulic system unit located a t the control surface). Emergency circuits may be provided for use in the event of hydraulic system failure.

Figure 4 shows a basic hydraulic system, which can be made to do work. With the pump running, fluid is drawn from the reservoir and supplied to the selector valve under pressure. When the selector valve is rotated one side of the jack is connected to the supply whilst the other side is connected to return. The jack will move. Returning fluid will go to the reservoir via the selector valve.

rnoodull l A-1070

RESERVOIR DIRECTION OF FLUID FLOW ---* n

(7 SUCTION LINE r\ PRESSURE LINE - & Jack moves in

PUMP

I - I ;;;;::P" RETURN LlNE

Four way selector valve. Valve selected to neutral position to provide a hydraulic lock to stop jack movement at any intermediate position

-.- - -- Valve selected to allowjack to move in the opposite direct~on -- A) Y / --

I '

Fig. 4 BASIC HYDRAULIC , , SYSTEM ,

I l 1

l 1

The selektor valve may be mandally or electrically operated. The drawing shows a simple rotating type of selector valve, buttsome are operated by a slide,type

I I piston. I ' I

I

! I With the valve as'shown the fluid causes the jacklactuator tb retract (move in) - with the - valve -- rotated clockwise by 90° the supply will be connected to the bottom of th2jack and the top will be connected goo.;eturn. The jack will extend. This is called a two-way selector valve- used for systems where only f i l l y in or fully out selections are required, eg landing gear.

Where intermediate selections are required - flaps for example - then a four-way selector valve will be used. In the drawing it is similar to the one shown but it can move 45" from the position shown so not aligning it's ports with any of the connections - supply, return or the jack connections. In this position it will form a hydraulic lock between the valve and jack and the jack will not move.

QUESTION: What is a hydraulic lock? (5 mins)

ANSWER: It is trapped fluid within a pipeline system - normally between a jack (or hydraulic motor or some other actuating device) and some other component such as a selector valve. In the example above the fluid is trapped between the jack (on both lines) and the selector valve. Under these conditions the jack cannot move.

There are at least two disadvantages with hydraulic locks. (1) The lock is not reliable. With all systems off, the fluid pressure, over a period of time, will dissipate - through very small internal leaks. (For example, when jacking an aircraft using hydraulic jacks they are reliably locked in the u p position using a mechanical device - usually a screw thread locking collar).

(2) If the ambient temperature rises the fluid temperature will increase and the pressure will increase. The pressures can get so great that structural failure of the pipes/components will result. Thermal relief valves are fitted to prevent this - more on this later.

The pipeline leading from the reservoir to the pump is called the suction line, with the line running from the pump to the selector valve the pressure line, and the line returning to the reservoir the return line.

Most aircraft have systems which are more complex than the one shown above. I t

( does not have any provision for pressure relief, off-loading the pump, standby suppiies etc. Reai systems have ail these, and for ease of ieGninC%-e'spiit into sub systems, eg power circuits, brake circuits, landing gear circuits etc.

There are two main types of system in use, the open-centre system and the closed system. The former is most usually found on some light aircraft and is not well known. The latter is common and found on'most aircraft - large argl small.

- -

I

The Open Centre System (figure 5)

The main advantage of this system is its simblicity, and the main disadvantage is that only one service can be operated a t a time. When no s e ~ c e s l a r e being operated the pressure in the system is a t a low value, with pump output passing directly to the reservoir round the 'open circuit' - with all valves in the open centre position.

- OPEN CENTRE SELECTOR VALVES

RELIEF VALVE I I SERVICE 1 SERVICE 2

Fig. 5 OPEN CENTRE SYSTEM

When a selection is made the appropriate jack moves. When the jack gets to the end of its travel it makes contact with the selector valve lever and moves it to the open centre position (the position of the left hand selector in the drawing).

With all selectors in this position the fluid is pumped around the system under very little pressure - thus saving energy consumption by the pump. Should there be a delay between the jack getting to the end of its travel and the de-selection of the selector valve, or should something fail, then there is a pressure relief valve to relieve excessive pressures.

The Closed System (figure 6)

With this type of system, operating pressure is maintained in that part of the system which leads to all the selector valves, and some method is used to prevent over-loading the pump. All consumer circuits have the same pressure supply - but s h g e the flow rate so individual flow rates might be different.

- -1 -- - - - -- - -

- -2 I

- - ;'ACCUMULATOR

' +

~ I /

/' IDLING CIRCUIT -

- / - - -- - _, +

I I SERVICE I It SERVICE 2 / - Fig. 6 SIMPLIFIED CLOSED SYSTEM

Closed systems are used on most aircraft.

In systems which employ a fixed volume pump (constant delivery) an Automatic Cut-Out Valve (ACOV) is fitted to divert pump output back to the reservoir via the idling line when pressure has built up to normal operating pressure. A n accumulator is fitted to assist in ACOV operation - more of this later.

In other systems where a variable volume pump (constant pressure pump) is used, delivery is reduced as pressure increases, and an idling line (case drain) allows some fluid flow back to the reservoir to keep the pump cool and lubricated.

- 10-

moodull l A-1073

In some simple light aircraft systems, operation of an electrically driven pump is controlled by a pressure-operated switch, which may be part of a power-pack assembly. The pump comes on when required and switches off when system maximum pressure is reached and no component is selected.

POWER CIRCUITS

The power circuit supplies fluid to, and accommodates the fluid returned from, the other circuits. It may also contain more than one pump. Pumps are usually engine driven, but may be electrically driven. The power circuit may contain one or more of the following components:

Driven pumps. Driven by: the enginels, electric motors, air (the ram air turbine [RAT]), hydraulic (power transfer units). Automatic cut-out valves (if constant volume pumps fitted). Pressure relief valves.

- - H=d pI I1ps. -

Reservoirs. Reservoir pressurisation system. Reservoir pressure refill connection. Oil coolers (in the fuel tanks). ' I

Accumulators. - -

-

Filters. I

Priority valves. \

I

Gauging systems for temperature, pressure, reservoir level, low pressure warning etc. Most modem aircraft this is all via a computer. Self sealing test couplings (ground test connections). Non-return valves (ch$ckvalves) . -

- -

Two mrnp Power Circuit

In multi-engined aircraft it is usual to have each power circuit using two or more pumps. They may both be engine driven (from different engines), or one may be engine driven and the other driven by hydraulic pressure from another system (power transfer unit), or electrically, or ram air operated. On small aircraft there may only be one pump and one supply system.

Figure 7 shows a circuit fitted with two self-idling (constant pressure) pumps, which, should one fail, will still provide fluid flow but at half the normal rate - so all systems will work but at half the normal speed.

The self-idling nature of the pumps means that they will automatically adjust the flow rate to suit the demand, and completely shut-off the supply when demand is zero - although they run continuously. The idling line (case drain line) will allow fluid flow to the reservoir when in idling mode to keep the pump cool and lubricated.

rnoodull l A-1 074

GAS CHARGING CONNECTION

PRESSURE GAUGE

RESERVOIR GROUND TEST CONNECTIONS ACCUMULATOR

TO NON ESSEKTIAL SERVlCES

DRIVEN WMP

1 :v////. 1 ;:En TO ESSEUTM SERVlCES

COMMON RETURN 1 = PIPEUNE; I ----------= ELEcTiaca C o N w c T t m s , I , I /

Fig. 7 TWO PUMP POWER SUPPLY CIRCUIT USING SELF IDLING PUMPS I

---- - I

, --

The pip+ines to the pumps are $led suZtion lines; from the pumps are called pressure lines and idling lineslcase drain. I I

I ' i I

The p ~ r b o s e qf ;he accumulators in this circuit is to give spe6dier operation of cornponknts'&nd provide a source of-hydradic power when the engine-driven pumpare~t -working - f& a-limited tim; period.

As an additional safety factor, some aircraft are fitted with a high-pressure relief valve in the power circuit - in case the pump fails to off-load.

The gauging (from the pump) includes a pressure gauging system (moving coil, dc or ac ratiometer, or synchro systems) a low pressure warning system (Bourdon tube or bellows operated micro-switch) and an oil temperature indicating system (thermister - moving coil or dc ratiometer systems). Gauging at the reservoir includes: oil level; temperature, and pressure - if the reservoir is of the pressurised type.

Indication that the power circuit is functioning correctly is provided by low pressure warning lamps, pressure gauges, and temperature gauges situated in the flightdeck. Filter by-pass warnings may also be available. If for any reason, such as a defective pump, defective valves, low fluid level or leaks, the pressure falls below normal working pressure, the pressure switch in the power circuit will operate a warning lamp - the gauge will also show a low reading - of course.

rnoodul l l A-1075

Over-temperature warning may indicate a failing pump; lack of fluid supply or reduced cooling effect to the returning fluid. In the event of failure of the engine- driven pump, the "second" pump will maintain the supply (but a t half flow rate); the hand pump, if to the flight crew, could be used in an emergency.

The hand pump has its own filter, pressure relief valve and non-return valve.

Both hand and engine driven pumps are provided with non-return valves to prevent the flow of fluid from either pump by-passing back to the reservoir through the pump not in use.

When the aircraft is on the ground, the hand pump can be used for maintenance purposes - operation of cargo doors, bleeding etc.

Some circuits are of vital importance to the safe operation of the aircraft, whilst others are of lesser importance. In general, PFCUs), wheel brakes and other essential services must have priority over non-essential circuits such as powered nose wheei steering and ianciing geeex~ension. Snouici the power clrcuit supply pressure fall below a pre-determined figure, a priority valve shuts-off the flow of fluid to the non-essential services and maintains the fluid prdssure fo; the essential services. I

The priority valve may be situated at the power supply circuit (as shown) with a pipeline going to all essential services and another pipeline going to allmon- essential services. Alternatively, the valve may be fitted a t the supply point to each non-essential service to shut it off in the event of reduced supply pressure.

i ! I

-

QUESTION: It is usual to consiger the lqding gear r e t r a ~ t i ~ n and nose wheel steering circuits as non-essential - why? (5 mins)

ANSWER: The landing gear has it's own emergency down systems - to include 'freefall', gas operated, separate pumps etc, and steering the aircraft on the ground can be achieved by differential braking if the nose wheel steering fails.

During servicing of the aircraft it will be necessary to test the hydraulic system, but a s the engine cannot always be used to run the engine-driven pump (when landing gear retraction testing for example), ground test connections are provided which connect the power circuit to a ground engine driven hydraulic servicing trolley (cart). The ground test connections are of the self-sealing type - either screwthread or bayonet type - usually of different sizes so cross connection is not possible.

Of course, the hand pump may be used for testing purposes, but the rate a t which it can deliver fluid is very slow and it will not reproduce actual operating conditions.

GAS CHARGING CONNECTION

COMMON RETURN = PI~EUNES ......-.----..-------= ELECTRICAL CONNECTIONS I

Big. 8 TWO PUMP CIRCUIT US IN^ CONSTANT VOLUME PUMPS I - I

i I 1

- -- -. I

I i ~ The constant volume pump sho ;~? in figure 8 (which pumps fluid all the time it is running) will need some form 06 '~ressure reliefy when services do not need fluid. If an ~ r d i n ~ , ~ r b s s u r e relief vqvk was fitted:in the system it would do the job,

I I but ,at an energy cost. The pump would have to keep pumping at it's normal working-pressure to keep the pressure relief valve open - using energy (fuel) all the time. So a special relief valve is fitted called an Automatic Cut Out Valve (ACOV) .

When operated it allows the pump to pump fluid at almost zero pressure via the idling line or case drain line back to the reservoir - keeping the pump lubricated and cool. More of this later.

Note. The two circuits shown above have two driven pumps each but may have only one pump on some aircraft, and may have three on others.

A320 System

Figure 9 shows an overview of the A320 aircraft hydraulic supply system.

Take a moment to study the system and note its main features. It has three systems - blue, green and yellow. All have accumulators with priority valves in front of the non-essential services.

RESERVOIR m RESERVOIR

PUMP 140 L/mn 140 L/mn 25 Llmn PUMP

37 Usq/rnn 37 Unq/mn

La RELEVATOR a

LANDING GEAR 1 Wl Fig. 9 THE A320 SYSTEM OVERVIEW

The pumps include a hand pump (for cargo door operation), Ram-Air driven Turbine (RAT) - for emergencies, engine driven pumps and electrically driven pumps.

A power transfer unit is provided between the green and the yellow system. Note that some services have three supplies - PFCUs for example, whilst others have only one - landing gear and steering.

The 13757 System

Figure 10 shows the power supply circuits of the Boeing 757. Take a few moments to study the drawing.

- 15-

rnoodull lA-I078

3- SUPPLY PRESSWE

3 m RETURN

AIR EOP - ENGINE DRlVEN P W P

ACMP - ALTERNATING CURRENT llOTOR PUMP RAT - RAM AIR TURBINE PTU - W Y E R TRANSFER U N I l SOV - SHUT O F F VALVE

With reference to figure 10 note the following:

* Three systems (left, centre & right) using seven pumps. * Pumps: 2 Engine Driven Pumps (EDPs) 3000psi @ nearly 40 galls

per minute; 4 electrically driven pumps (ACMPs). Pressure supply similar to above but low flow rates (about 7 galls per minute; 1 RA - for emergencies - manual or automatic deployment based on both engines low rpm - also possibly connected to airspeed and weight switch inputs. Driven by a V P propeller and supplies about 2000 psi at about ?A of the normal pump flow rate; 1 air driven pump (additional pump on the 767 fitted to the centre system).

* Single point reservoir filling. * Fire Shut-Off Valves (SOVs). * Pressurised reservoirs. * Fuel cooled heat exchanges. * Power transfer unit. Comes on automatically when left engine rpm or

left EDP pressure is lost. Is a bent axis hydraulic motor driving a hydraulic pump. About 2,000 psi at about half the normal pump flow

- --

rate. I I

Note the power supplies to the following consumer circuits: '

* Wheel brakes - 2 * Powered steering - 1 * Landing gear retraction - 1 * Flaps and Slats - 1 * Yaw damper - 2 * Tail plane trim - 2 * ~ u t o pilot servos - 3

-

* Aileron, rudder and elevator PFCU's - 3 -

* Spoilers - 3 * Thrust reversers - 2

CONSUMER CIRCUITS

These are circuits/services that use the pressure supply from the supply circuits. Remember that for most aircraft all services have the same pressure supply (unless there is a pressure reducing valve fitted) but they will share the fluid flow. This means that if more than one circuit is selected a t any one time then the fluid flows to all the circuits will be different.

Consumer circuits can include all those listed for the B757 plus windscreen wipers, cargo door operation, hydraulically operated pumps etc. Each consumer circuit will have a pressure supply line and a return line from/to a supply circuit with, if necessary, switching provision to other supply circuits.

- 17 -

rnoodull l A-1 080

Alighting Gear Circuit

The circuit illustrated shows the layout and components required to control the raising and lowering of one landing gear unit with the other units having a similar arrangement - with the selector valve being common for them all. Fluid is supplied from the power circuit to the control valve via a non-return valve. This valve ensures that the alighting gear circuit is isolated from the rest of the aircraft hydraulic system by providing a hydraulic lock.

NRV = NON RETURN VALVE

PRESSURE SUPPLY TRV = THERMAL RELIEF VALVE

Fig. 12 LANDING GEAR RETRACTION CIRCUIT - SINGLE SEQUENCE

The control valve, which may be manually or electrically selected, directs fluid to the desired end of the jacks/actuators and at the same time connects the other line to the reservoir. The lines to/from the actuators are known as the up and clown lines.

The purpose of the remaining components in the circuit are as follows:

!?herma1 Relief Valve (TRVI. Due to thermal expansion of the fluid (eg, hot climates, operating temperature rise etc) in a closed circuit, there is a risk of burst pipelines and damaged components. To prevent this a thermal relief valve is fitted in both the up and down lines.

The valves will relieve expanding fluid to the circuit return line. They are pressure operated and designed to relieve a certain amount of pressure whilst leaving some pressure behind.

One-way Restrictor Valve. When alighting gear down is selected the free fall (static drop) of the alighting gear could damage the undercarriage unit attachment points on the airframe and cause cavitation in the down line (the fluid being drawn by the falling undercarriage quicker than the pump can supply), therefore, to slow-down the rate of fall, a one-way restrictor valve is provided in the up line.

This valve, which restricts the flow of fluid in one direction, but permits full flow in the opposite direction, offers no restriction to the flow of fluid when alighting gear up is selected. It prevents fluid getting away from the jacks too quickly when down is selected - thus slowing them down.

Shuttle and Fluid Jettison Valves. Separate the Alighting Gear circuit from the Emergency Down gas operated circuit. They allow the landing gear to be "blown downn'"in= emergency such as total-kydraulic failure. More details later.

-\ -

\

~echa!nJcal Sequence Valves (Single Sequence Circuit\. To avbid collision between the ulddrcarriage leg and it's fairing door dqking operation, the' undercarriage compo,nents must move in the correct order,(or sequence). For example the undercdriage leg must fully retract before~it's fairing door start$to close. A typical skquencing operation would be: -, \ ,

1 I

' 1 , U p selection , 1 Pressure fluid to leg jack and leg retracts.

! , 2. Once retracted, 'UP' sequence valve operated I

I / /

LAN mechanically by leg, j ack. I - - 3. Fluid flows through sequence valve to door jack.

4. Door closes.

Down selection 1. Pressure fluid to door jack and door opens. 2. At fully open position 'DOWN' sequence valve operated

mechanically by door jack contact. 3. Fluid flows through sequence valve to leg jack. 4. Leg lowers.

Landing gear without fairing doors do not require this system and some aircraft with fairing doors have the doors mechanically connected to the main undercarriage leg and, therefore, employ only one operating jack.

The sequence valves may be operated mechanically (as above), or hydraulically, or electrically and many of the larger aircraft have double sequence systems:

Down selection Door opens Leg comes down Door closes

rnoodull l A-1083

Up selection Door opens Leg retracts Door closes

Additional sequence valves are fitted to cope with a double sequence system or the system is electrically controlled.

Emergency Operation. Many large aircraft have a free-fall emergency facility and some have a separate dc motor driven pump and hydraulic system to lower the landing gear should all else fail. The one shown has a gas operated system. Some have a manually operated wind-down system.

Tlne gas operated system has nitrogen stored under pressure in the gas bottle. S'hould an emergency arise and the undercarriage fail to lower by the hydraulic system then the pilot can operate a manually operated emergency selector. This allows nitrogen under pressure to the shuttle and jettison valves.

TL'3 " h , ? t t l n r r n l ~ r n ;c ,-.~??cC.A tn -nT,L c,,V-r.&" 0-#-I -0- ,,+-,Aev +-.+-a""7,va z-+o+-- +La 1 J L I ~ O L r u L L r b v - v b ~u b u ~ u v u c u r _ l r u r b u b s u u ~ LUAU ~ U V ULICLUI ~ A L O U U ~ L LIICLL u CIIL

-- -

undercarriage down line. The door jack, sequence valve, and leg jack operate as normal. I

Returning fluid from the circuit has to be,hllowed to be jettisoned (the emergency could be a jammed selector valve --or loss ,df electrical power to the valve, if electrically operated). So the compressed gas operates the jettison valve which allows returning fluid to be jettisoned overboard. I

I I I

Note. When servicing the circ~it~after eme&ency operation it is important to: ,

, -- -

(a) Select emergency selector to OFF. -

(b) Release gas pressure carefully from hydraulic lines. (c) Ensure jettison and shuttle valves are reset. (d) Rectify original fault. (e) Bleed hydraulic system. ( f ) Recharge gas bottle - check for leaks and function test.

MJheel Brake Circuit - Small Aircraft

On some small aircraft the brakes may be operated by a lever and cable system - operating a drum, or brake calliper and disc assembly. On many light aircraft a small hydraulic system is provided. Figure 13 shows such a system.

When the toes are pushed down (on the rudder pedals) on the foot motor, pressure in the hydraulic fluid causes the brake cylinders to operate. When the toes of both feet are pushed down in-line braking is achieved, when one is pushed down, one brake will operate and the aircraft will turn.

- 21 -

rnoodull I A-1 084

The blow-back valve allows the flaps to be blown back if lowered a t an excessive GI- speed (or left down after take-ofq which would otherwise cause them to be darnaged by the air-flow. This provision only applies to flap circuits that operate using jacks - as the aerodynamic loading can be felt back through the jack and hence back through the fluid.

If the flaps are operated using hydraulic motors - as with many large aircraft -the fl,aps will have a separate load relief system. This consists of a Pitot operated pressure capsule connected to a motor connected to the flap selecting linkage. (With flaps down, a s the airspeed increases so the Pitot capsule will make an electrical contact to select the motor to move the flap selector linkage to select the flaps to a higher position) - with pilot warning.

T:he thermal relief valve relieves excessive pressure built up caused by temperature increase in a closed circuit and the throttling valve ensures a constant rate of flow of fluid irrespective of supply rate and is always fitted in the flaa down line. It works in both directions.

QUESTIQN: Why could the flap circuit (or any other circuit for that matter) I ' have a varying flow :ate supply? Surely the pump, particularly a 1 1 constant delivery pump, will supply the fluid at a! cgnstant rate. ! ' (5 mins) I )

I -- ANSWER: I I If a second circuit id selected, thin fluid supply fqom the purnp/s would be shared (on a closed hydraulic system), an? hence speed of

I , opeiation, of the first circuit will be reduced. I I ; I

/ ' 1 I

QUESTION: SO why is it import8nt'that the flaps always move at the same speed each time? (I'm not talking asynchronous here). (5 mins)

A.NSWER: Operation of the flaps causes a trim change of the aircraft. This must be the same on every flight so that the pilot knows how to react. So the throttling valve sees to it that the flaps always move at the same speed each time they are selected.

QUESTION: What other circuits might use a throttling valve? (2 mins)

ANSWER: Possibly PFCU circuits and power steering circuits.

When a selection is made fluid flows to one side of the operating jacks. Returning flluid from the other side of the jacks goes via the selector valve to the reservoir through the return line. Irrespective of the selection the fluid flows through the throttling valve which ensures a constant rate of flow at all times - IN & OUT.

This system operates a simple plain or split flap.

rnoodull l A-1086

PORT STARBOARD OPERATING

I i Fig. 14 FLAP CIRCUIT - PLAIN OR SIMPLE FLAP I I

,

If any pipeline full of fluid is trapped a t any time - in this case, the flap up-line - a thermal irelief valve must be fittedr(EFregulation), This works by pressure which has bee+ created by a n increasd ih fluid temperature.

I I

Flap Synchronisation -- - - - - - - - - -

Port and starboard flaps must go u p and down together, if they do not for any reason, a roll would be induced. If this happened a t low altitude a crash would be almost inevitable. (In the 1960s a BEA Elizabethan twin engined aircraft carrying show horses rolled violently on approach to Heathrow airport. It crashed killing all the crew and the horses, and writing-off other aircraft on the ground. The cause was a failed flap operating linkage on one side of the aircraft. The incident was caught on movie camera by an amateur camera man.)

The most common way to synchronise flaps is to connect them together mechanically using a common drive system - in that way one flap cannot move without the other. For small aircraft an operating rod/torque tube connecting port and starboard flaps may be connected to a pilot-operated handle. For many large aircraft the flaps are moved using hydraulically/electrically operated motors that rotate common drive shafts to operate screw jacks in the wings to port and starboard flaps.

Figure 15 shows a hydraulic method of synchronisation - rare but interesting. It uses 2 constant volume jacks, one connected to the port flap and the other to the starboard flap.

rnoodull lA-10187

SYNCHRONISING JACKS

OPERATING JACKS

Fig. 15 HYDRAULIC

PORT FLAP

! METHOD OF FLAP SYNCHRONISATION

STARBOARD FLAP

Tliey are cross connected in such a way that if one flap is moved the other will be folrced to move in unison because of the tr-ansfer of fluid from one side of the first jack to the other side of the second jack.,Fluid make-up valves and tQermal reliei valves are fitted to the cross feed lines (not shown in the dradirig). \

I I ~ I

1 F'owler IT'ype Flap Circuit (figure ;18)

'

i I i ,-- . Oln most large aircraft the flaps lowered by being pushed backanddown on

I 1 \ track sysfems (though some, th& DC10 aircraft for example, still have a simple I 1 I hinge drrangement) . 1 I

I I / ,

I / he flapsarePushed back on tracksusink ascrew jack and dri+einut assembly. The screw-jack-is rotated by the drive-shafis and the drive-nut (attached to the flap) is caused to move forward or back.

FLUID SHUT-OFF VALVES ASYMMETRIC DETECTOR UNIT /

DRIVE SHAFT RPM PICK-OFFS

MECHANICAL

Fig. 16 FOWLER TYPE FLAP OPERATING CIRCUIT

- 25 -

rnoodull l A-1 088

The screw jacks are rotated using drive shafts and gearboxes. The drive shafts are rotated using a duplicate system of hydraulic motors.

Flap asymmetric operation is prevented, in the first instance, by the drive system being common to both sides, but if it does happen it is detected by detectors at the end of each drive shaft. The detectors pick up the drive shaft rotation and send the signal to the asymmetric detector unit.

This unit compares the signals of both sides and if one shaft rotates quicker or slower than the other then the fluid shut-off valve is caused to operate. This stops flap movement and warns the pilot.

This type of system is also used to operate the leading edge slats/Krueger flaps.

Air speed sensors will operate to select the flaps up (a small electric motor changes the effective link length or a computer operates the normal flap selection system) if for any reason they are left down in an accelerating airflow (after take- off fo< exep le ) . If the flaps ar_e selzted -- -- down at high speed the air speed sensor will pqevent the selection being made to the I hydraulic motors.

1 I 1 : 1

The ~ 3 2 ) O Flap System (figure 17) I /

This is similar in principle to the system-s,hown in figure 16. Each Power Control Unit (PCU) is supplied by two hydraulic systems and these operate drive shafts to operate the flap drive gearboxes.

I

~syn$ne@-y detectors fitted at each drive shaft end monitor their rotation and infoms_the computer if the flaps move asymmetrically. If detected the computer will stop the flaps moving and send a warning to the flight-deck.

When the pilot makes a selection via the computer the hydraulic valve on the hydraulic motor selects and the motor causes the drive shafts to rotate. A s the shafts rotate so a pick-off on the motor power control unit sends positional feedback information to the computer. When the flaps get to the selected position the computer will operate the control valve to the mid-off position to stop the motor - the flaps being held in their new position by a hydraulic/mechanical lock.

QUESTION: What sort of feedback is used in this system? (1 min)

ANSWER: Negative feedback.

QUESTION: Define negative feedback? (2 mins)

ANSWER: It can be defined as a system where the output of the system tries to cancel the input.

Surface support track

Flap (and slat) positions are indicated on the flight deck by ECAM (Electronic Central Aircraft Monitor). The signal will go via a Symbol Generator Unit (SGU) to be displayed on a CRT (Cathode Ray Tube) in colour. The slat operating and asymmetric detection system operate in a similar way to the flap operating and detection systems.

Brake Circuits - Large Aircraft

Figure 18 shows an older type system. Typically two supplies are provided - one normal (or main) and one emergency (or standby). Both are identical from the Non-Return Valve (NRV) to the Pressure Reducing Valve (PRV). The accumulators hold a quantity of fluid under pressure to be used when there is no supply (towing for example). The NRV ensures that this fluid is not used by other circuits when the pumps are off.

Notes: ,--- - - - - -- - - - I -

1. $rC&ure transducers (to f/ighEEk gauges) are fittedto the brake lines Pytween the brake contro! valve andthe anti skid units, and between the emergency brake control valve and'shuttle valves - but these are not shiown for clarity. I I / I

2 . !The brake control valve is;similar to/that shown in figure 40, with each foot hbtor having its own reservoirTOh-some aircraft - the Tristar for example - khk brake control valve is bperated b> cables from the rudde; bar so there i s I I no separate'foot motor hydraulic systerh for each foot on each rudder pedal '(as shorn' in the drawing). I I I I ,

1 '- The pressure operated switch pl.o$dei_a warning on the flight deck should the pressure drop to some low value. The pressure transducer and gauge provides system supply pressure indication on the flight deck. The gauge could be a moving coil, dc or ac ratiometer, or a synchro system . (See the book in this series entitled Airframe Instruments).

The pressure reducing valve reduces pressure from say 3000psi to say 600psi. For supply 1 this pressure goes to the brake control valve.

Lines to the brakes are often fitted with hydraulic fuses so if lines are ruptured (due to runway debris damage etc) the system fluid will not be lost.

The brake control valve is operated by slave units which in turn are operated by master cylinders on the rudder pedals. When a master cylinder is operated this causes the slave unit to move - which in turn causes the brake control valve to allow fluid pressure from supply 1 to the brakes. (For more detailed information on Brakes, Autobrakes and Anti Skid systems refer to the book in this series entitled Wheels, Tyres and Brakes).

rnoodull l A-1091

SUPPLY I SUPPLY 2

NRV

b ACCUMULATOR LOW PRESSURE

PARKING BRAKE SWITCH

PRESSURE LOW PRESSURE WARNING LAMP

PRESSURE GAUGE

SHUULE VALVES

Fig. 18 BRAKE CIRCUIT - LARGE AIRCRAFT

This supply then goes via the anti-skid unit to the brakes. The operation of both rnaster cylinders will achieve in-line braking. The aircraft can be steered using either the left or the right foot motor separately. (For normal operation large aircraft are power steered through the nose wheel).

In an emergency, system 2 can be operated directly using the emergency brake lever.

rnoodull l A-1092

In the system shown the emergency system by-passes the anti-skid valve, but with most aircraft the standby system is an alternate system with a complete set of duplicate components to the main system, but powered from another supply circuit.

When auto brake is selected the brakes are selected ON automatically after wheel spin-up to brake the aircraft quickly or slower depending on selection.

System 2 also stops the wheels rotating during retraction, so they are not spinning in the wheel bay. This is automatic in operation. When 'up' selection of the undercarriage is made fluid pressure causes the spring return jack to operate and the brakes to come on. After a short interval the pressure in that part of the 'up' line is released - the jack spring returns the control valve to the off position and the brakes are released. On modern aircraft this function is controlled by a computer. Fluid pressure from supply 2 goes to the brakes via shuttle valves.

I

-. r- HYDRAULIC COMPONENTS I /

This $art of the book describes vakious aircraft components of a general nature. sever$ manufacturers produce'compone~ts and there are many different types in use. However, all components designed for one task rely upon the same basic princi'ples although their appearance and name may differ.

I ' - -

Reservoirs I

Some, functions of a reservoir are: - - - - - - - - - --

(a) Supply the pumps - with a head of pressure (a positive pressure). (b) Accept return fluid from the system. (c) Hold a reserve of fluid to allow for small leaks. (d) Act as a heat sink. (e) Allow for jack ram displacement (fluid level changes). ( Provide a filling point.

The actual size and shape will vary from aircraft to aircraft depending on capacity, location etc. A simple unpressurised reservoir is shown in figure 19. Study the details and note the level sighting glass, the filter, the vent relief valve, the return and suction connections.

Figure 20 shows a pressurised type with a separator piston - note the contents gauge system, the suction and return lines and the bleed and nitrogen connections (not a common type). Figure 2 1 shows a pressurised type as fitted to Airbus aircraft. Study the pressurisation system and make sure you know how it works. Note the indications - reservoir quantity, oil temperature and pressurisation pressure.

moodull l A-1 093

VENT VALVE

\n

SIGHT GLASS

FILLER CAP

FILLER NECK

FILTER

CONNECTION

AFFLES to prevent eturning fluid jetting

into air space

RETURN CONNECTION the aircraft attitude

I , -- - 1 \ I - - - -

- I

I Fig. 19 UNPRESSU~SED , RESERVOIR , ' 1 I

Fig. 20 PRESSURISED RESERVOIR - WITH SEPARATOR

Note the negative g trap so that fluid is always available to the suction line, even under negative g flight conditions.

Since it is not always possible to mount the reservoir above the pump, and to ensure a positive supply of fluid to the pump, many reservoirs are pressurised. The pressure (relatively low - about 30psi - but check your AMM) in the reservoir also helps to reduce fluid frothing which could affect the operation of the system.

The method of pressurising varies, but may include the use of compressed gas acting against a piston or diaphragm in the reservoir, or air from a compressor stage of the jet engine (pneumatic system). If it is part of the aircraft's pneumatic system the designer must ensure that there is no possibility of hydraulic fumes entering the pneumatic supply to the cabin.

blank

rnoodull l A-1095

ANTI EMULSION

NEGATIVE G T

DRAIN VALVE

I . , - I .. , , '

ENGINE DRIVEN 1 8

OVER HEAT , I

PUMP SENSOR

I ,

RAT PUMP , ' ,, , i /

- ! _ i

GROUND CONNECTION

RESTRICTOR

RELIEF VALVE

Fig. 2 1 THE A300 PRESSURISED RESERVOIR & PRESSURISING SYSTEM

rnoodull l A-1 096

VENT flPE 70 ATMOSPHERE

I

! Fig. 22 RESERVOIR L0,CATION - EXAMPLE

Heat Exchangers, 1 1 I

In some systeys,heat exchange& ,are used to cool the hydraulic fluid from the pumps. This extends the life of the fluid and the pumps. They are positioned in the fuel tanks-so providing a heat sink from the hydraulic fluid to the fuel (also warms the fuel and prevents the freezing of any water at altitude). Not fitted to all aircraft.

Fig. 23 HEAT EXCHANGER

- 34 -

They are normally situated in the idling line (case drain) between the pump and the reservoir - so cooling the fluid returning to the reservoir and where more than one supply circuit is used each will have a separate heat exchanger positioned in a separate fuel tank.

They are positioned low in the tanks so should always be in fuel. If the fuel gets to0 low the pumps can be still be operated provided the hydraulic oil temperature does not get higher than that laid down in the AMM {eg 100°C (2 12°F)) or after the pump fault light comes on.

F'ilters

E;ASA requires that the hydraulic system should be adequately filtered. This means that 'dirty' systems would have more filters fitted that 'clean' systems.

In general fdters may be fitted: - -

% -- - --

L - * After the reservoirin the pump supply line (low $ r e s i r e filter). * , After the pump in tlie bressure line (high pressure fdter). , -k In the return line to the reservoir (low pressure filter). * In front of some circuits that require special prot&cdion - eg where

valves are fitted that rkly-on,&-to-metal contalctifor flulid,'sealing (some PFCUs) . - . I

\ I 1 ' \

The working parts'of hydraulic cbrqponentdI have very small cleqances and working limits A d it is, therefore, most important that the hydqaulic fluid in the system is scgpulously clean otyeryise -- excegsive wear, damage or even blockage could occur. - -,

HEAD

O-RING

-FILTER ELEMENT

O-RING

BOWL

Fig. 24 THE FILTER OF THE A300

moodull lA-10983

Before being put in the system the fluid passes through a fine filter in the dispensing equipment and then through another filter at the filling point (on some reservoirs). The filter in the dispensing equipment is a micronic type, it's filtration level being in microns; one micron = 0.00004in; a five micron filter therefore would allow particles smaller than 0.0002in to pass unhindered.

The type of filter element varies depending on its position in the system and the manufacturer.

The basic construction of all filters is similar with those in the pressure lines being of more robust construction that those in the return lines. Elements are made from:

(a) Felt (b) Paper (c) Wire gauge or cloth (d) Wire wound spool

rp(e) /- Sintered metal -

- -

j(f) i Aluminium foil , (g) Magnetic plug I

I

1 1

I / i

QUESTION: What does 'sintered' mean? (5 mins) r - - - _

ANSYER: It is a process of manufacturing small metal parts using fine I powd'er. The powder is measur,ed out to a precise quantity and placed

into accurately machined dies. Here heat and pressure is applied to I , L - form small parts such as gear wheels, valve bodies and filters.

The item leaves the diewithout any further machining necessary. - -- - -- - -

If the applied pressure is not too great then the fine particles bond together in such a way as to allow small passages between them - thus forming a filter element.

Filter elements are normally made in discs about 2" to 4" in diameter (SO to 100mm) and stacked one upon the other inside the filter body.

Sintered metal filter elements are cleaned using ultra-sonics and a counter flow of fluid in a special cleaning rig.

Many filters are fitted with a clogging indicator. When the filter element becomes blocked (or nearly so) a pressure differential is created (about 30psi - AP) in the filter body to cause an indicator button to be pushed out. This is usually coloured red and gives a visual indication that the filterlfilter element needs changing.

The indicator is kept out by a spring and may be pushed back (after the removal of the transparent cover) once the filterlfilter element has been changed - or as directed by the AMM.

Provision is made that it will not operate until the fluid has attained it's correct working temperature.

1 1 Fig. 25 FI~TER CLQGGED INDICATOR / I

- - ,' - i - ,'

- -- I ----- ' - - --

Pi11 filters must, by regulation, have provision for by-pass should there be any chance that fluid starvation could occur if the elements became blocked. If the filter element becomes blocked then a differential pressure (AP of about 60psi) will cause the by-pass valve to open. This allows fluid through unfiltered - better this than no fluid at all. Of course, the clogged indicator will show.

When changing the filter element always change any seals at the same time, and always make sure that the replacement element is the same as the one removed - or an acceptable alternative as per the IPC (Illustrated Parts Catalogue).

INith some filters, as the bowl is lowered, internal valves close off the inlet and outlet to the filter body thereby reducing the amount of fluid lose. Remember to bleed the system after a filter change - as per the AMM.

Fig. 26

I I -- ' I

COMPONENT LOCATION - EXAMPLE I I

1 AIRBUS

A hand pump is included in some aircraft installations, for emergency use and for ground servicing operations. Figure 27 illustrates a double-acting hand pump (ie a pump which delivers fluid on each stroke).

As the piston moves upwards in the cylinder, fluid is drawn in through an NRV at the inlet connection into the cylinder; at the same time fluid above the piston is discharged through a non-return valve in the outlet connection.

A s the piston moves downwards, the inlet NRV closes and the transfer NRV opens, allowing fluid to flow through the piston. Since the volume below the piston is larger than the volume above the piston, some of the fluid (about half) is discharged through the outlet port. When pressure in the outlet line exceeds the relief valve setting, discharged fluid is by-passed back to the pump inlet.

This a double acting single cylinder pump. Hand pumps are available from manufacturers that are double cylinder and single acting.

OUTLET

PRESSURE RELIEF VALVE

- - -- -- - -

M - 1 -. f INLET

Drawing f r o r n ' , ~ ~ ~ 562 I

I

I ' Fig. 27 DOUBLE ACTING SINGLE CYLINDER HAND PUMP I

I . ~ c c u r n u l ~ t o r s i i \ I -

, I \ '

Fitted to consume? and supply circuits to: '

1 1

I 1

1 * ' I I I !

-- Holds a reseye offluid Lnder pressure for - ~ H e n pumps - - not running. / I

x Allows instantaneous operation of systems. J; Assists the ACOV in operation.

are

Fig. 28 TYPES OF ACCUMULATORS

May also be used to damp out pulsations from the pump and thus cushion the shock loads which the circuit might otherwise have to withstand. However, with modern pumps this function is not always necessary.

Accumulators may also be fitted to essential services, such a s powered flying control circuits and wheel brakes, mainly to provide a reserve of pressure in the event of supply non-availability. They are normally charged with compressed nitrogen, but might, on some small aircraft, be spring-loaded or charged with air.

PRESSURE GAUGE CONNECTION

- -- -

Fig. 29 AIRBUS ACCUMULATOR

May consist of a cylinder and floating piston or a spherical container and flexible diaphragm.

QUESTION: Why charged with nitrogen and not air? After-all nitrogen is expensive, air is more or less free. (5 mins)

ANSWER: Nitrogen does not support combustion - air does. Should a fine spray of hydraulic fluid occur in a pressurised gas container a condition similar to a diesel engine could be set up if the gas is air - and combustion can occur.

Pumps

M:ost aircraft are fitted with multi-piston type hydraulic pumps, driven from the engines. Other types of pumps, such as gear or vane positive displacement pumps miay be found in some installations, but these normally do not provide sufficient piressure - though flow rates are high. Pumps may be powered by:

(a) Another hydraulic circuit (hydraulic motor - power transfer unit). (b) An electric motor - normally ac but may be dc on some emergency

systems. (c) A Ram Air Turbine - RAT. (d) The Engine Driven Pump - widely used (EDP). (e) The air flow in the by-pass section of a turbo fan jet engine. ( The APU.

The pumps may be classified as either:

w'-,-' PRESSURE LlNE

ENGINE DRIVE PUMP t IDLING LlNE

Fig. 30 CONSTANT VOLUME PUMP - GENERAL ARRANGEMENT

PL constant volume pump (figure 30) has 2 hydraulic connections - suction and pressure - and has to have an Automatic Cut Out Valve (ACOV) fitted in the power circuit with an accumulator. A constant pressure pump (figure 31) has 3 c:onnections - suction, pressure and idling line (called case drain line on many aircraft).

- 41 -

rnoodull l A-1 104

ENGINE DRIVEN PUMP

SUCTION LlNE PRESSURE LlNE

I IDLING LlNE OR CASE DRAIN

Fig. 31 CONSTANT PRESSURE PUMP - GENERAL ARRANGEMANT

Constant Volume or Non Self-Idling Pump

These pumps deliver a fured quantity of fluid to the system at a particular rpm, regardless of system requirements, and means must be provided for diverting pump output when it is not required by the system.

- - -- - - - -

n-,-k,- Af nt-.-~+nrn+ + ,A~ , I -P m,llfi-ni~tfnn m11-n i e i l l ~ l ~ t ~ ~ f p ~ in fi0-11t-p '29 The W l l L cyyu "A b u r r u r r u L L ""Aurzr" r r r u r o yAuc_u** y,ss**y -u , A * - u . . - - - u - --- --.3--- --. ---- cylinder(block and drive shaft rot4t.e together, and because of the angle between the cyli<der block and shaft axes,!each piston moves into and out of it's cylinder once ger; revolution.

1

I ' 1

The stationary valve block has t y v ~ ~ @ ~ u ~ r e n t i a l slots leading to the cylinder block,'which are connected to the,fluid>irlet and outlet ports, and are 'arranged so

pistons draw fluid into the cylinders on the outward stroke, and push to the system on the inward stroke.

CYLINDERBLOCK

BLOCK

UNIVERSAL JOINT

Drawing from CAP 562 Fig. 32 CONSTANT VOLUME AXIAL PISTON PUMP

Another type of constant volume pump is illustrated in figure 33. In this pump the cylinders are arranged radially around a crankshaft, so that when the crankshaft is rotated, each piston moves in and out within it's cylinder once per revolution.

rnoodull l A-1 1105

Fluid is drawn into the pump body and enters each cylinder through ducting in tlhe cylinder block, whenever the associated piston moves to the bottom of its stroke. A s it moves outwards into its cylinder, it covers the inlet port, and forces fluid out of the top of the cylinder, past a delivery valve, to the pump outlet connection. The drawing shows one piston for clarity but pumps have many pistons with some having a dozen or more.

/DELIVERY VALVE

- PRESSURE

/ 1 I Drawing f r o r n , ' ~ ~ ~ 562

J?/ig. 33 CONSTANT - XOLUME RADIAL PISTON PUMP -

/

I --

Constant Pressure Self-Idling Pump (figure 34)

This type of pump is similar in construction to the fmed volume pump but the cylinder block and drive shaft are co-axial and piston travel can be varied - so varying the output. The pistons are attached to shoes which rotate against a stationary yoke, and the angle between the yoke and cylinder block is varied to increase or decrease pump stroke to suit system requirements.

When pressure in the system is low as would be the case following selection of a service, spring pressure causes the yoke to move to its maximum angle, and the pistons are a t full stroke, delivering maximum output to the system.

When the selected system has completed its operation, pressure builds up in the supply line and under the control piston. This moves the yoke to the minimum stroke position. In this position a small flow through the pump is maintained (liom inlet to case drain) to lubricate the working parts, overcome internal leakage and dissipate heat.

On some pumps a solenoid-operated depressurising valve is used to block delivery to the system, and to off-load the pump.

Pipeline connections usually vary in size (Murphy proofing), typically:

* Suction line - largest. * Pressure line - medium. * Idling or case drain - smallest.

Fig. 34 CONSTANT PRESSURE PUMP

Pressure Relief Valves (PRV) (figure 36)

These are fitted between pressure lines and return lines. They are designed to relieve pressure should it build up above normal working pressure. They are not designed to operate continuously as this would keep the pump 'on load'. Pressure relief valves are usually a spring-loaded ball or a spring-loaded plate - often adjustable - but by the manufacturer only or in the hydraulic overhaul bay.

- 44 -

rnoodull lA-1107

Fig. 35 ENGINE DRIVEN PUMP LOCATION - EXAMPLE B 7 6 7

ADJUSTER

RESERVOIR

SYSTEM PRESSURE \

-> l~rawing from CAP 562 I I Fig. 36 PRESSURE,' RELIEF VALVE I

I

QUEST1ON:These valves often all look alike but are set by the mahufacturer to operate a t different pressures depending on the system they are designed for. How would you know which is the correct one for a particular system and location? (1 min)

ANSWER: By reference to the description and part number, (on the component), the AMM (aircraft maintenance manual), the IPC (illustrated parts catalogue) and the EASA form 1.

A.utomatic Cut-Out Valve (ACOV)

Fitted to systems employing a constant volume pump, to provide the pump with an idling circuit when no services have been selected. An accumulator is essential when an ACOV is fitted, since any slight leakage internally or externally would result in continuous operation of the ACOV (hammering).

QUESTION: Why wouldn't an ordinary PRV do? It would be cheaper. (2 mins)

ANSWER: If the pump operated a PRV during idling it would have to work all the time against the spring - thus consuming the same energy as when the pump is operating the services normally. Obviously not a good idea. When the pump is off-loaded by an ACOV it does little or no work during it's idling cycle. A s an example one pump made by Dowty used 14HP (10,444 Watts) when on-load and when off-load used only 2HP (1490 Watts) - quite a saving in energy and fuel.

QUESTION: Why would an ACOV 'hammer' if an accumulator was not fitted? (1 0 rnins)

ANSWER: Because the fluid is (more or less) incompressible, pressure build up to normal maximum would be almost instantaneous without any service selected so the ACOV would cut-out almost instantaneously.

There are always some internal (allowable) leaks in a system. The

-- -- quantity (volume) of fluid to be lost to drop the pressure from (say)

I - 3000psi normal working pressure to (say) 2500psi cut-in pressure of

the ACOV would be;vevSmall -just a few drops. This could be lost' I in seconds - causink the ACOV to 'cut-in'.

I

I I When cut-in and on-load, the pump would take a fraction of a second to build up the pressure from 2500psi to 3000psi (as the volume

I '

required is so small). TEEs the ACOV would cut back out again - in a fraction of a second; This cycle, would be repeated very quickly , causing ACOV hammering. The accumulator gives the system some 'resilience' it takes time to build the pressure u p - and time to loose it

1 I ) so the ACOV may cycle, say once every 30 minutes or so instead of

every second or so2_ -

When a service has been selected and the pump is delivering fluid to the system, the NRV is open and equal pressure is applied to the top of the poppet valve and the bottom of the piston. The force of the spring combined with the pressure on the poppet valve is greater than the force on the piston, so the valve is closed and the return line to the reservoir is shut.

When the service selected has completed it's travel and fluid is no longer required for the systems, the pressure applied to the piston is sufficient to lift the poppet valve off its seat. This results in a sudden drop in pressure on the pump side of the valve which snaps the poppet valve open and the NRV closed.

Pressure in the idling line drops to a low value and the load on the pump is removed. Pressure in the system is maintained by the accumulator until further selection is made - when pressure drops, and the pressure on the cut-out piston becomes less than the spring force, the poppet valve closes and pump output is again directed to the system through the NRV.

POPPET VALVE

PUMP SUPPLY - NON-RETURN VALVE

Fig. 37 ACOV

Priority Valve (figure 38)

A priorityvalve is basically a pressure relief valve which is kept-open if system supply pre$sure is normal. Should this drop, to some pre-dedgned-value, then the valve willclose - shutting off the supply toithe secondary servicks. In this way, if the supply is failing, what fluid is available' will be for the (most 1

important) services only. I

1 - - I ! I

The valve, will open to allow fluidlto-the secondary services if noI?nal sudply is I / ', '! / resumed.' 1

PRESSURE INLET PRIMARY

SERVICES

SECONDARY SERVICES

Drawing from CAP 562 Fig. 38 PRIORITY VALVE

A. priority valve is generally used to safeguard operation of important services such as flying controls and wheel brakes. Figure 38 shows the valve in the open position, pressure being sufficient to move the piston against spring pressure and connect the main supply to both the primary services and the secondary services.

A. priority valve may be fitted immediately after the power circuit with one pressure line supplying all the secondary services and another pressure line supplying all the primary services.

rnoodull l A-I 1 10

On some aircraft the priority valve is fitted in front of each secondary circuit. Each secondary circuit will be connected to the secondary services port of the valve and the primary services port will be blanked off.

Pressure Reducing Valve (figure 39)

A pressure-reducing valve is used to reduce main system pressure to a value suitable for operation of services such as wheel brakes. Figure 39 illustrates a pressure-reducing valve, which also acts as a relief valve for the services operating at the reduced pressure.

Fluid enters the inlet port, and flows through the valve to the low pressure sub- system. When the fluid pressure exceeds the spring-loading on the valve, the valve is lifted and gradually covers the inlet port until sub-system pressure reaches the specified value - when the supply is shut-off. If sub-system pressure increases for any reason, the valve is lifted further --- - - and uncovers the return port to relieve excess pressure. - - -

1

/

RETURN

PRESSURE INLET

_, LOW PRESSURE SUBSYSTEM

Drawing from CAP 562 Fig. 39 PRESSURE REDUCING VALVE

Brake Control Valve

This valve is quite 'a box of tricks' so you will need to take it slowly. They vary from aircraft to aircraft so you should not need to commit the details to memory - but you should understand the principles on how it works.

If the brake control on your aircraft is different and you understand it then fine - don't bother with this one. But either way the CAA will expect you to know how the brakes are operated on a large civil aircraft.

- 48 -

rnoodull l A ~ 1 1 1 1

This brake control valve is essentially a variable pressure valve, which controls pressure in the brake system according to the position of the pilot's brake pedals or hand brake lever. The valve usually contains four elements, one pair for the brakes on each side of the aircraft, to provide duplicated control. Figure 40 illustrates a single element, in this case operated by a slave servo from the brake pedal master cylinder.

When either pilot's brake pedal master cylinder pedal on the appropriate side is depressed, or the hand brake is operated, the footbrake servo applies a force to the linkage on the control valve, which, via the lever assembly and plunger, presses down the exhaust valve cap. This action initially closes the gap between the exhaust valve cap and the exhaust valve seat, then moves the cradle down to open the inlet valve and direct fluid to the brakes.

Drawing from CAP 562 Fig. 40 BRAKE CONTROL VALVE AND SLAVE UNIT

Pressure builds up in the brakes and under the valve until it is sufficient, assisted by the spring, to overcome the inlet pressure and the force exerted by the plunger fo'rce. This pushes the whole assembly upwards to close the inlet valve and the increase in pressure to the brakes stops.

An increase in the load applied to the valve plunger will be balanced by increased delivery pressure, and a decrease in the load applied will be balanced by relief of delivery pressure past the exhaust valve cap to exhaust. In this way the pressure applied to the brakes is proportional to brake pedal pressure.

When the brake pedals are released, the cradle moves to close off the inlet port, the exhaust valve cap lifts, and exhausts the pressure from the brakes to the reservoir.

Non-Return Valves

The NRV or check valve or one way valve is a common device used to control the flow of fluid. It permits full flow in one direction, but blocks flow in the opposite direction. Simple ball-type non-return valves are common but designs may vary. When a non-return valve is used as a separate component, the direction of flow is indicated by an arrow moulded on the casing, in order to prevent incorrect installation. The arrow points in the direction of flow.

Hydraulic Fuse

This~vdvEjallows normal fluidflow through, but should it become excessive due, say, to &massive leak down straamihen it will shut and prevent further fluid flow. It maybe known on ~rnerikan aircraft as a Waterman fuse. It operates using the principle of differential pressure across the valve.

I / I

The valve will permit normal flow, but if the flow rate rises akjove a predetermined level the valve will close its outlet linepreventing further flow. Often fitted to wheel brakelines due to the high probability of damage (and leaks) to from flying h n w a y or @re debris. On some airc+aft, fuses are fitted a t locations throughout the,/hydraulic system. 1

1

CHAMBER.

I)

OPEN CLOSE RESET

Fig. 4 1 BRAKE LINE HYDRAULIC FUSE

Figure 41 shows the operation of the valve. The fluid flows into the fuse and enters the upper chamber via a small metered orifice in the spring loaded piston. The fluid also flows through the valve to the brakes.

During normal operation the pressure differential that exists either side of the spring-loaded piston is minimal therefore the piston remains a t the full-flow position and an uninterrupted supply goes to the brakes. If an excessive flow rate oc:curs (due to a large leak), the pressure difference on top of the piston through the metered orifice will be sufficient to push the piston down and block the exit side of the fuse. It will remain closed even if the brake pedals are released - as reservoir pressure will hold it closed. The fuse will need to be bled and reset - with the brakes off.

IIblPORTANT. Always carry out a full brake/system test after fuse operation/ bleeding to check for correct re-setting.

Restrictor valves I --- I -

7 - \

I I I

May be tyio-way or one-way. A two-way restrictor restricts fled ih both directions; a one-way restrictor restricts flow in one direqtion - with full flow in the other.

' I I I

The restriction is usually of fmedi s&e,as Bhown in figure used i$ a number of locations, in order to-limit the speed

I I actuatorjjack. It may, for instanred be used to, slow down landing gear extension. I

i I 1 I

42. 9 qestriStol;'valve is of operation,of a n flap r6traction or 1

Drawing from CAP 562 Fig. 42 ONE-WAY RESTRICTOR VALVE

QUESTION: Why are flaps and landing gear likely to move too fast? (2 mins)

ANSWER: The landing gear will tend to fall under its own weight and the flaps will tend to be blown u p by the airflow (if jack operated).

rnoodull lA_1114

It is important to note that while they restrict the flow of fluid, the actual flow rate through the valve will be related to delivery rate, delivery pressure and the design of the valve.

Selector Valves

These may be manually or electrically operated, and can be of the four-way type (eg, where flap operation would require an intermediate selection), or of the two- way type - where no intermediate positions of a service are required(eg landing gear retraction). The purpose of a selector is to direct fluid to the appropriate side of the actuator/jack/hydraulic motor, and to provide a return path for fluid displaced from the opposite side of that actuator.

A two-way selector valve connects the pressure and return lines to alternate sides of the actuator, without a neutral position. Selectors in open-centre systems will trap fluid (hydraulic lock) in the actuators while providing an idling circuit for the pumpTSFme manually ~ p e r a t e ~ v d v e s are shown in figure 43.

I '

It is sdmetimes necessary to be Able to hold the actuator in an intermediate position (flaps for example). o n 1 some airctaft this is achieved by using a selector which; btocks both lines to the actuator when it is in the neutral position, the selector being manually returned when the desired actuator position is reached. However,, as this could be distraLtingf0r-the pilot at a critical stage of flight, a feed-bacF mechanism is usually, used, which automatically returns the selector to neutral whenever the selected position is reached.

I

I I I

I /

QUESTION: what sort of feed-back --- would - th:ls be? (2 rnins)

ANSWER: Negative feed-back. Where the output tries to cancel the input.

Electrically-Operated Selectors (Figures 44 &, 45)

When the selector valve is located at a position remote from the crew compartment, and to eliminate the need for extensive mechanical linkage the valve is normally operated electrically - or at least initiated electrically - the actual operation is done hydraulically.

The selector shown in figure 44 is a typical electrically initiated (28V dc) two-way valve, which may be used, for example, for operation of the landing gear.

With the solenoid de-energised, the pilot valve is spring loaded against the return seat, and fluid from the system passes to both sides of the slide valve.

rnoodull lA-1115

JACK

JACK

JACK

I P

ROTARY VALVES t JACK

SLIDE VALVE

t i 6 JACK P JACK

*=

Fig.

'For use

SLIDE VALVE with auto return. After selection when the jack gets to the end of its travel the pressure builds up and passes through the NRV on the pressure line. This allows presure to act on the slide piston causing it to slide back to the neutral position.

with open centre systems PRES,SURE RETURN]

SLIDE PISTON

/

I NRVs

I - ,,' ,

I Drawing from CAP 562 ' / /

1 I' I

1 I i I I / I / I ; I ~ ~ Since the1 righfhmd end of the valve is of larger diameter than the left, the valve

moves-to the left and fluid passes to the leftside of the actuator! F;luid from the opposite side of the actuator passes thrmgh the selector to 'the return line.

With the solenoid energised, the pilot valve is held against the pressure seat and supply pressure acts on the left-hand side of the slide valve only, the right-hand side being open to return. The slide valve moves to the right (as shown), and directs fluid to move the actuator in the opposite direction. The other side of the actuator being open to return - via the valve.

Flour-Way Selector (figure 45)

These are used where an intermediate selection of a service is required, eg flap circuits. They have two solenoids (28V dc operated) which control a slide in much the same way as in the previous type.

With one solenoid operated (solenoid 'B' ) the supply to tha.t side of the valve is shut off and the pressure that was there is allowed to return to the reservoir. With pressure to only one side of the slide valve the slide valve will move over to allow fluid to one side of the jack. The other side of the jack is allowed to return via the other port.

rnoodull l A-1 11 6

$ TO SELECTOR SWITCH

SUPPLY

I b3- SOLENOID

I : ~ r a w i n g from CAP562 I

Fig. 44 ELECTRICALLY INITIATED TWO-WAY ! SELECTOR ,VALVE

I I I

,I-- ' I '

With doth solenoids de-energised fluid pressure is allowed to ,both sides of the slide valve. ,

I ~ 1

In figure 45 the effective area oflthe slide valve piston on the left is greater than that1 - of thc - - -- sght.'This p- is due to the fact that the annulus slide is away from it's body stops but resting on the slide stops. Thus the slide valve and annulus slide move to the right until the annulus slide rests on the body stops - the effective area then becomes equal and the slide valve stops - in the middle.

blank

TERMINAL BLOCK EARTH RETURN

I

SLIDE

SOLENOID "B" SOLENOID "A"

I I SUPPLY I t

- 1

(a) Both solenoids de=<nergised I \ \

i \, ' , I SUPPLY I

RETURN

ANNULUS SLIDE

SLIDE VALVE PISTON

e RETURN

J TO JACK

\ FROM JACK

(b) Solenoid 'B" energised

Drawing from CAP 562 Fig. 45 ELECTRICALLY INITIATED FOUR WAY

SELECTOR VALVE - OPERATION

rnoodull lA-1118

Fluid Jettison Valve

Fitted in the emergency gas system this valve permits up line fluid to be dumped overboard on emergency down selection.

On selection of 'emergency down', gas pressure is fed into the base of the valve (figure 46) to act on the piston to unseat the ball valve. UP line fluid can now vent to atmosphere. When gas pressure is released the piston will be moved down by it's spring and the ball valve will automatically reseat.

FLUID FLOW

--

I

GAS -

, PRESSU

I I ,

JETTISON FLUID

I Drawing from CAP 562

Fig. 46 !FLUID JETTISON VALVE I

1 -- -

Shuttle Valves (figure 47)

These are often used in landing gear and brake systems, to enable an emergency or alternate system to operate the same actuators as the normal system. During normal operation, fluid flow is provided from the normal system to the service and the emergency linelalternate supply is blocked.

TO SERVICE

A

ALTERNATIVE I OR

__C EMERGENCY SUPPLY 1

NORMAL SUPPLY

Drawing from CAP 562 Fig. 47 SHUTTLE VALVE

When normal system pressure is lost and the emergency system is selected (or when the alternate system is used), the pressure moves the shuttle valve across, &lowing the emergencylalternate supply to the actuator.

Slome shuttle valves have an electrical connection to them to monitor the normal supply pressure or to indicate piston position. This data being sent to control computers.

RETURN FLUID

Drawing from CAP 562 I

Fig. 48 SEQUENCE VALVES I I ' - - ' - , / " I

Sequence Valves

Sequence valves are often fitted in landing gear circuits to ensure correct sequencing of the landing gear jacks and door jacks. Examples of mechanically operated and hydraulically operated sequence valves are shown in figure 48.

Slequence valves ensure that the landing gear does not extend until the doors are open, and that the landing gear is retracted before the doors close.

Completion of the initial movement of one of the actuators results in part of the rr~echanism contacting the plunger of the mechanical sequence valve, moving the piston and allowing fluid to flow to the next actuator.

T:he two valves shown are operated mechanically/hydraulically. But, of course, valves can be operated electrically using a dc supply; a microswitch operated by the moving component, and a solenoid to operate the valve.

moodull 1A-1120

Modulators

A modulator is used in conjunction with the anti-skid unit in a brake system. It allows full-flow to the brake units on initial brake application, and thereafter a restricted flow. Figure 49 shows a modulator, the swept volume of which would be equal to the operating volume of the brake cylinders.

BRAKE PRESSURE

TO BRAKES

ORIFICE PISTON

Drawing from CAP 562 Fig. 49 MODULATOR VALVE

- - - -

, /-- -- -

~ u r i n k ibitid operation of the biakes, the. piston is forced down the cylinder against gpring pressure, and the 6rakes are applied. Subsequent fluid feed to the brakes, necessitated by anti-skid Gnit operation, is through the restricting orifice and is limited. This limited flow allows the anti-skid unit to completely release the brakes when necessaty, and coLservesqain system pressure. When the brake control valve is released, the pis').on returns to its original position under the influence of the spring and the returning fluid from the brakes.

/

Flow Control V&e

A flow control valve may be fitted in a hydraulic system to maintain a constant flow of fluid to a particular component - similar to a throttling valve. It is frequently found upstream of a hydraulic motor or jack which is required to operate a t a constant speed. A typical flow control valve is shown in figure 50 and consists of a body and a floating valve.

VALVE SEAT

OUT

DAMPER UNIT FLOATING VALVE VALVE HEAD

Drawing from CAP 562 Fig. 50 FLOW CONTROL VALVE

Flow through the valve head is restricted by an orifice, which creates a pressure drop across the valve head. At normal supply pressure and constant demand, the pressure drop is balanced by the spring and the valve is held in an intermediate position - the tapered land on the valve partially restricting flow through the valve seat, and maintaining a constant flow through the outlet.

If inlet pressure rises, or demand increases, the pressure differential across the v,alve head also increases, and moves the valve to the left (in the drawing) to reduce the size of the aperture and maintain a constant flow.

The spring loading is increased by the valve movement, and again balances the pressure drop. Similarly, if inlet pressure drops or demand decreases, the valve takes up a new position, to the right, so as to maintain a constant flow.

Hydraulic Jacks or Actuators ,--

The purpose of a hydraulic jack or actuator\ is to convert hydraulic pressure and --

flow into Linear motion. I I I

/

Drawing from CAP 562 1

Fig. 51 TYPES OF JACK

SCRAPER COMBINED ELASTOMERISLIPPER

Fig. 52 JACKS - EXAMPLE OF SEAL LOCATIONS

Jacks are the main component in a hydraulic system for the conversion of hydraulic power to mechanical power. In one form or another they are found in landing gear circuits, flap circuits, spoiler circuits, powered flying control units, nose wheel steering circuits etc.

moodull l A-1 122

There are two basic types in use depending on the requirements of the system they actuate. The most common is the unequal area type which is adequate for most requirements. However, due to the ram, piston area on one side is less than on the other and thus extension and retraction forces (and speeds) will differ.

Where it is necessary to generate the same force or same speed in both directions an equal area jack is used. In this case the ram on both sides of the piston reduces both areas by the same amount. The type of jack will also dictate the speed of operation. Given the same fluid supply in each case an unequal area jack will move in faster than it will move out. With an equal area jack it's speed of operation is the same in both directions.

QUESTION: Which one of the jacks shown in figure 5 1 would be fitted in a power steering circuit? (5 rnins)

ANSWER: One double acting equal area jack - similar to the A320 or two

!, - 1 double acting un@al Zea jacks - similar to the B747. I

I

QUES~ION: Which way should the unequal area jack move to retract the I landing gear ideally? (3 mins) I

ANSW~R: It should move out or to the right in the previous drawing. This

1 1 will mean that it uspsits largest area and provide its greatest force.

, SPRING

RESTRICTOR VARIABLE RESTRICTOR -

--\-

Drawing from CAP 562 Fig. 53 THERMAL RELIEF VALVE

Thermal Relief Valve

To relieve slow pressure rises due to thermal expansion of the fluid a thermal relief valve is fitted between the pressure line it is to protect and a return line. Similar in principle to a pressure relief valve but it incorporates a restrictor which ensures that when it opens it will relieve only some of the pressure. Works on pressure only.

Throttling Valve

A specialised restrictor valve similar to the flow control valve which automatically regulates the flow-rate in inverse proportion to supply pressure to maintain a constant speed of component operation. These may be one-way or two-way.

Figure 54 shows a two-way valve in operation. When too much fluid tries to flow through the valve the piston is moved so that the metering needle enters the outlet orifice, thus restricting the fluid flow. The valve is so designed that the o-utlet fluid flow rates are within the normal operating parameters.

NORMAL FLUID FLOW RATE

4

SPRING

, I t ; ' HIGH FLUID '

, _F_L_OW RATE - /

~ r a w i n ~ f f r o 5 ~ ~ ~ 562 Fig. 54 THROTTLING VALVE

Pressure Relay Valve

A pressure relay valve is a component which transmits fluid pressure to a direct reading pressure gauge (Bourdon tube type) or to a pressure transmitter which el.ectrically indicates pressure on an instrument in the flight deck.

In some cases both types of indication are provided, the direct reading gauge being fitted in the hydraulic equipment bay, adjacent to the relay. It transits fluid pressure but not fluid flow.

During normal operation the piston acts as a separator, moving within a limited amount to transmit fluid pressure from the supply side to the gauge side. If a leak develops on the gauge side, the piston moves to the gauge end of the cylinder, and the valve seats in the cylinder head, thus preventing leakage from the system (there will be a small leak from the gauge line, but only a few drops).

moodull 1 A-1 124

The valve also permits bleeding when a new gauge, or gauge line, i s fitted. This is done by clamping (with a special clamp around the body) the piston to prevent its movement. Pressure build u p from the system will move the valve to allow fluid to pass through and bleeding a t the gauge to take place. The clamp must be removed afterwards (it has a long warning streamer attached).

TO GAUGE

P IS~ON VALVE

Drawing from CAP 562 Fig. 55 PRESSURE RELAY VALVE

-- ---

7 I- -

-

Quick: Disconnect Couplings , I I I I

In positions where it is n e c e s s h to frequently disconnect a coupling for servicing purposes, a self-sealing, quick disconnect coupling is fitted. The coupling enables the line to be disconnected without-loss of fluid (without the need for subsequent bleeding) and they be bayonet or screw type.

' I 1 ,

I 1

Pressure Release Valves I I ,

\ - - ~ i t t e d t o en-able pressure to be released from the system for servicing purposes. The valves are manually operated, and consist of a valve body with a n inlet and outlet port, the passage between the two being blocked by a spring-loaded valve. Operation of an external lever opens the valve against spring pressure, and allows fluid to flow from the pressure line back to the reservoir.

Drain Cocks/ Sampling Valves

Drain cocks are simple manually operated spherical valves, and are located in the hydraulics bay at the lowest point in the system. They are marked to indicate direction of flow, and are used to drain the system in order to replace the fluid, or in some systems to change certain components. A sampling valve allows fluid to be taken for analysis a t periods specified in the maintenance schedule or at times when hydraulic fluid contamination is suspect.

- 62 -

rnoodull l A-1 125

Fire Shut-Off Valve (SOV)

E;lectrically operated and mounted between each reservoir m d each of the engine- driven pumps (within a fire zone). Closed when the flight deck fire handle is operated.

SAMPLING

AMPLING VALVE

II

Fig. 56 FLUID S A M ~ L ! ~ G VALVE --, \

--- - \*- FLUID SAMPLING

' I

I I \

1 I 1

1 These make it poidible to cut-offltqe supply:offluid quickly seconds) in the event of a fire in the enginej1The shut-off valve's are normally controlled~fr~m/the hydraulic panel\by th,e/PUMPS ON-DUMP- HUT VALVES switch&. The shut-off valves c&e when-the associated swiche$-are placed in the SHUT VALVES position. When the switches are a t either of the other two positions, the valves are open.

Fig. 57 FIRE SHUT-OFF VALVES - POSITION

- 63 -

rnoodull lA-1126

In the event of an engine fire, the fire control handle is pulled and the fire shut-off valve will close irrespective of the PUMPS control selector position.

Hydraulic Motors

This converts hydraulic power into rotary motion, usually by the swashplate principle. These motors which are reversible, may be operated continuously, intermittently, or stalled without damage. Used for powering generators (rare) and in systems such as spoilers, flaps, etc. The speed of the motor depends on the fluid supply rate or the angle of the swash-plate and the torque it can produce depends on the supply pressure.

S W BEARINGS

' I I I VALVE PLAT; INLET SLOT

~ /

- - ~ i g . 58 BENT AXIS PISTON TYPE HYDRAULIC MOTOR -- -

Although the basic principle of operation is common in all types of hydraulic motor, each aircraft system will employ its own specific design of motor.

Pressure fluid enters the motor through a valve block (figure 58). It is directed to the individual cylinder bores through a valve plate. The non-rotating valve plate provides the timing for the motor.

Pressure builds up in the cylinder bores until the resulting force on the drive shaft overcomes the resisting torque. The pressure of the fluid is transmitted from the piston rods to the drive shaft, overcoming the torque resistance of the load connected to the drive shaft. Fluid under pressure enters via the inlet slot to those pistons on their outward movement. The pressure exerted on the piston will push it out - but it can only go out if it moves up (in the drawing). This it does, and at the same time moves round. This causes the cylinder block to rotate in the direction shown (figure 58).

Tlhe universal link keeps the drive shaft and cylinder block in alignment since the piston rods have ball joints on both ends. The drive between the cylinder block and the drive shaft is maintained by the universal link. The link however, does not transmit load torque. Each piston rod has pressure-lubricated ball joints at both the piston and drive shaft ends.

EMERGENCY / STANDBY SYSTEMS

One or more of the following systems/components may be fitted to an aircraft to pirovide standby/emergency hydraulic supplies:

1. Accumulators. 2. Electrically driven pumps - ac and dc. 3. Duplicate/ triplicate systems.

I 4. Duplicate/triplicate components - PFCUs, yaw dampers, A/P servos. ' 5. Power transfer unitSfrorone system to another;---- 6. # Manual reversion (PFCUS). \ , \ I -

7. Gravity freefall (UJC down). ! 8. Gas operation. I I

~1 I I 9~. Ram air turbines (RAT). I _ - I

1'0. Air driven pumps. I / I

Ill. Separate system with Kd~-driven pump (U/C down on some aircraft). \ I I

1 I I I ' I

A RAT System I

I 1 /

L- - - I j

The RAT-^&<^^ stowed in the Kottom - P- of thefuselage or in th& underside of the wing - usually inboard. The following is an example of a hydraulic RAT system based on the Airbus A300.

R A M AIR TURBINE

Fig. 59 RAT LOCATION

rnoodull l A-1 128

1 \ 1 1 G R E E N H Y O R A U L

R A H A A S

1

Fig. 60 HYDWULIc RAM AIR TURBINE (HYRAT) -,

! -

I

This vkriable pitch ram a i r -d r i~dn~~ro~e l l e r is housed inside the wing root and the housiqgis closedlby two mechahi&ally operated doors. Two control handles at the first and second officers positions allow RAT extension.

The RAT leg is secured in the retracted position by an up-lock which is mechanically operated by a control cable. This runs from the flight-deck to the inboard wing using turnbuckles, fairleads, cable quadrants etc.

When the RAT is released the ejection jack extends under it's spring pressure to thrust the RAT into the airstream. The latter stage of the jack extension is retarded by the hydraulic damping device to avoid high impact at full extension of the unit. A down-locking pin secures the RAT in the extended position.

The doors are mechanically opened and locked by the RAT movement.

With the RAT in the stowed position the test selector is set to provide a permanent supply of warming fluid. This maintains the power pump and the gerotor pump (an Airbus term) at the correct operational temperature. The flow is controlled within the control module and by the flow-sensitive valve in the leg.

The return line is provided with a filter monitored by a clogging indicator.

- 66 -

rnoodull l A-1 129

A:j soon as the RAT is down and locked the turbine is simultaneously unlocked by withdrawal of the locking rod and immediately commences to rotate due to the airflow. After the momentary off-loading effect provided by the flow sensitive valve, th~e power pump operates to supply full pressure to the yellow system. Hydraulic pressure supplied by the gerotor pump for operation of the speed governing mechanism passes via the turbine shaft sleeve assembly into the piston valve.

As the turbine rotates, centrifugal force causes the governor weights to pivot outwards and move the piston valve against it's spring pressure. This allows high pressure fluid to pass through ports in the sleeve and shaft and flow to the annular chamber.

Here the pressure acts on the cylinder which is moved forward against it's spring, turning the blades toward the fine pitch position. Low pressure fluid, displaced from the space forward of the cylinder, passes via the cylinder guide and central ducts of the piston valve and drive shaft to return to the gerotor pump.

rnoodull l A-I 130

Any increase in turbine speed will occur until the pre-determined governing speed is reached. Any tendency of the turbine to change speed is corrected by variation of the position of the governor weights and piston valve. Which thus maintains a balance of pressures to keep the speed constant.

In the event of an excessive speed occurring, due to any malfunction of the Mechanism a spring loaded dump valve will open under centrifugal force and high pressure fluid will be bled from the cylinder chamber into the low pressure cavity to balance the pressure. The cylinder will thus move to coarsen the blade pitch and reduce the speed.

Figure 6 1 shows the RAT system schematic. There is no need to remember the details, but you should be able to follow the general principles.

Rat Test

It is possible to test the RAT pump -- performance on the ground. A special test selector allows the pump to be ulsed as a motor to drive the propeuer.

I I

INERTIAL EFFECT

MI\\\.\\\\\\\\\\

TO YELLOW SYSTEM

Fig. 62 RAT TEST

To carry out the test the Yellow hydraulic ground supply must be connected and aircraft electrical power on. The M T leg must be extended and locked down. The area must be cleared with position warnings posted and the area roped off.

The general procedure is:

1. Consult the AMM.

2. Connect and start test rig and select switch to ON. Check that red light comes on.

3 . Set the test switch to 'SPIN'. This energises the test selector solenoid. The pressure and return lines are crossed-over to supply the pump with fluid pressure and make it work as a motor.

4. Check the speed of the propeller on the rpm indicator. 5. When normal speed is reached (5300rpm) release test switch. The test

selector solenoid is de-energised and the motor is returned to pump mode with the pressure/return lines configured for pump operation.

6 . Power is now supplied to the pump by the inertia of the propeller and pressure output is monitored by a pressure switch which activates the green test rig light if the delivered pressure is correct.

7. Reset the RAT, reconfigure the aircraft and record all work done and sign paperwork.

Reset Procedure

1. I Reset the control hrmdle. - - -- -

I - 2. Check that the filter cJogging igdicator is 1N'.

I ' \

3. i Set the blades perpendicular to allow the blade locking pin to engage the spline.

4. Pull back and lock the door control pins. I

I 5. Release the down locking-pin' ar5d push RAT leg dpward (two man

! job), lock u p mechanicallyusiG the up-lock unit; ,

6. Release the door contrbl pins &pd\release the door lbck. 7;. ~1ose''Ihe doors - enbure they fit flush.

8. ~ecokd all work dond and sign ihe appropriate dohments. /

I ' i I ,

i - -' ' -

I / - - 1

THE POWER PACK

Light aircraft may be fitted with a power-pack which contains all the hydraulic components including pumps, valves etc to operate the services. The pump is usually powered by an electric motor which is caused to switch off when normal maximum system pressure is reached. It comes on automatically when system pressure is required (figure 63).

When the selector lever is operated the pressure from the accumulator starts to move the actuator. A s the accumulator ram moves up, the switch operating collar moves away from the motor switch and the pump commences it's pumping cycle. The actuator will extend until full travel is reached. Further pump action will recharge the accumulator until the fully charged position is reached and the switch-operating collar will operate the motor switch and stop the pump.

rnoodull l A-1 132

AIR CHARGING VALVE

\

, Drawing from CAP 562 Fig. 63 POWER PACK

I I

To prdv&nt leakage of fluid past pibtons, &ston rods, etc, special seals are fitted. The matbrial used depends on the type of fluid used in the system, while the shape,of a seql is governed by the pressure of the fluid and the purpose of the seal.

- - - - - - - - -

Hydraulic seals are made of natural rubber, synthetic rubber, or other synthetic material (check the earlier pages in this book on hydraulic fluids). Modern seals are made to meet the fluid specification (MIL SPEC or local o r national specifications) to cope with the temperature range, operating conditions and fluid types.

When replacing a seal it must be replaced by a seal of identical shape, material and part number. This may be indicated on the seal and/or on the package. Seals to MIL SPEC H5606 will usually have a red or blue band. For Skydrol seals the band is green. Check the AMM, components manual and/or the IPC. Some seals may have coloured dots or strips painted on them for identification purposes.

Care must be taken when replacing seals as they can easily be damaged. Special tools are provided for there removal/fitment.

Square Section Seal. This provides sealing in both directions and can be used as a piston ring seal. It is either located in a square groove in the piston head or supported by Tufnol rings on either side.

rnoodull lA-1133

Fig. 64 SQUARE SECTION SEAL

'O' Section Seal. This seal is used to provide a fluid tight joint for contact surfaces where no movement takes place (ie between the mating faces of two halves of a component where a fluid tight joint is required), especially where the shape of the joint is irregular. If used on moving surfaces with pressures above 1500psi they should be backed with backing rings to prevent distortion.

-

- : Fig. 65 ROUNDO~~O' SECTION SEAL -- \ I - - -

! I I

\

I '

~ o n d e d Seal. Consists of a met4 washer with n rubber seal bonded to its inner - surface!, the rubber being slightly thickerhah the metal. It is u ~ u d l y fitted with a hollow bdlt that is torque loaded and iS used for sealing end c k p , banjo/unions, e1.c where no movement takes p l a c e 7 ' \

-

a

! ', . I

I

I

- i

- - I RUBBER

METAL

Fig. 66 BONDED SEAL

Wiper Ring/Scrapper Ring. This is not a seal as such, but fitted to prevent - dirtldebris from entering the seal assembly on a piston ram, undercarriage oleo leg etc.

/ WIPER EDGE

Fig. 67 WIPER RING

- 71 -

rnoodull l A-I 134

Duplex Seal. Consists of a hard rubber square section ring with a soft rubber square section ring bonded to its inner face. Resists both high and low pressures and is suitable for fluid/gas components.

HARD RUBBER @ SOFT RUBBER

Fig. 68 DUPLEX SEAL

Chevron Seal. If used singly will withstand pressure in one direction only. If used backed as a pair will withstand pressures in both directions.

The drawing shows a single chevron seal designed to withstand pressure in the direction-shown. It is normally-supported by back-up spacers and spreader packingipieces between it and the component.

1 @'""'" 0" 'A'""

I CHEVRON SEAL,

! I

i t I PRESSURE DIRECTION I

Fig. 69 CHEVRON SEAL -. - - -

'V' Ring/'UY Ring Seals. Similar to the chevron seal, though the W' ring seal has a shape similar to a U' rather than a V'. Normally used without backing pieces and are fitted singly.

PIPELINES

All pipeline information including materials, methods of connection, and identification are part of the module 7 syllabus. The following, however, is included as revision and is based on the Airbus aircraft.

The pipelines are placarded with Skydrol-resistant identification bands or metal rings, on which are indicated the system colour code; a brief description of the line function, and the direction of fluid flow. Pipeline materials are stainless steel for high pressure rigid pipelines, aluminium alloy for low pressure pipelines (except where they pass through fire zones) and for flexible hoses - braided, Teflon lined.

rnoodull l A-1 1:35

Hydraulic components are identified by means of self-adhesive placards showing the component number a s given on the hydraulic system general drawing. A s a rule, numbers beginning with 1, 2 or 3 are used for Green, Blue and Yellow system components respectively.

1 1063 SELF ADHESIVE

IDENTIFICATION LABEL

HYORAULIC SYSTEM NUMBER ANDCOLOR CODE

SYSTEM

GREEN OREEN SYSTEM

BLUE SYSTEM

FUNCTION AND FLOW ARROW

I -

1 I

- ~ --

Fig. 70 COMPONENT & PIPELINE IDENTIFICATION 1 I '

FLIGHT-DECK,INDICATIONS I I 1 \

\ \ i 1 These are covered in more detailini the book on instruments in this module but below i's a generdl outline of instrumentation as specifically applied to hydraulic

I ' systemk. , I 1 I 1

-- - I - -

FlightdEB indications will inclbde instru~ents/indicators that wbrk on the fbllowing principles:

* Warning lamps with signals from pressure/ temperature/contact operated switches. Discrete signals.

* Warning horn/chimes with signals coming from similar sources. * Direct reading Bourdon tube type gauges. Rare on large aircraft,

though are fitted in servicing areas within the aircraft. Comrnon on older small aircraft.

* Moving coil type instruments. k dc & ac ratiometer type instruments. k Synchro-resolvers. x CRT screens/flat screen displays.

]For details of the principle of operation of these instruments systems you should ]refer to the book in this series entitled Airframe Instruments.

rnoodull l A-1 136

A Typical Indicator System

The indication systems which follow are based on the Airbus family of aircraft. These can be considered as typical for a large modern aircraft.

BLUE I GREEN I YELLOW SYSTEM SYSTEM S Y S E M

I SELECTORS

I

1 I I

INDICATION I I

I 1 I 1

I

1 I ! ' ~ i & 71 HYDRAULIC1 SYSTEM CONTROL PANEL - A300

1 I

-

- - - - --

The hydraulic systems are controlled and monitored from the flight decl hydraulics panel. The panel arrangement represents the fluid flow from to the pumps and high-pressure delivery to the systems.

C

the ti mks

The control panel shown (figure 71) is vertically divided into three sections, one for each of the three independent systems. The panel shows pressures for all three systems for reservoir pressurisation and system supply pressure.

Warnings are given for:

* Low reservoir pressurisation. * Low system supply pressure. * High fluid temperature. A Low reservoir fluid level.

The panel allows for operation of: dumplshut-off valves for the EDPs (shuts the fire-valve and de-activates the EDP); electrically operated pumps and the power transfer unit.

rnoodull l A-1 137

YELLOW

RAT

LO A I R > L O AIR LO A I R PRESS PRESS PRESS

OVHT L OVHT TAT t I9'C C.W.60300 KG SAT t 1B'C C.G.ZB.1 %

@ Reservoir LO AIR PRESS indicatian :

@ Rcscruor OVHl indication :

@ FLRE VALVE indicrtion :

@ Elcc pumps O W 1 indication :

@ RAT control :

@ ELEC indication :

@YELLOW ELEC PUMP cunrrol :

@ PTU control :

@ ENG PUMPS control and low prcrue i n d c a h .

\ / - - -- -

1- r-

Fig. 72 HYD&LICS CRT PAGE OF THE A320 I

1 ' I ~ I I

I 1 I '

Figure ;72 shows an example of 2 f l & h t d e c k , ~ ~ ~ display of thd hydraulic system of the A320. It is a colour display and gives,details such as: pressure indications; low pressure warnings; fire-valvdl pbsitions;,oyerheat etc. It is belectedusing the page selekor panel in the flight deck and i ipar t of the ECAM1(~lectronic Centraliskd Aircraft Monitor) system of the aircraft.

I 1 1 1 i 1 1

Note that in fi&re 73 the OR gate in-the ldgiL circuit means that a warning will be given if any one of the following~occurs: A

1. Low hydraulic pressure from the pump pressure gauge. 2. Low hydraulic pressure from the low pressure transducer. 3. Low reservoir pressurisation air. 4. High temperature hydraulic fluid. 5. Low fluid level in the reservoir.

Fleservoir Level Indicating System

This is a float and arm mechanism which drives a synchro-transmitter shaft via a magnetic damper which rninimises the transmission of the float oscillations to the synchro-transmitter.

PL direct reading gauge pointer is mounted on the opposite end of the synchro- transmitter shaft which allows the fluid level to be read at the reservoir location without power being on. The electrical output of the synchro is sent to the flight dleck gauges/ computer.

Fig. 73- WARNING LOGIC \ / -

-

I 1 \ ' The LEVEL LO warning light is activated by a switch which is closed by a cam mounted on the synchro when fluid level gets low - about 5 1 (1.3 US gallons).

I

I ! I / Figure 714 shows the wiring circuit forthe level indicator sys ted - note the lines to I I the repeater gauge (for the aircrpi servicing panel). The low level switch grounds

the LEVEL LO light circuit and thk HYDRAU inputs the MWS. 1 I I I I 1 I

The pde$s-to-test caption light slvitch, grounds the light circuit to test the lamp and [to activatk the master warning.

- pp -

i n d l TO REpE,iiAuGE

26V GUANTITY AC OUANT.lNDICATOR TPANSMITTER

Manifolds & Tanks LOW LEVEL

SWITCH

MWS

28V DC

Fig. 74 RESERVOIR LEVEL AND WARNING INDICATION CIRCUIT

- 76 -

A high temperature warning circuit is provided for each of the three systems

T:he overheat sensor consists of one arm of a resistor bridge. The bridge output signals are amplified and when the temperature gets too high, (85°C [185"F]) it triggers a transistor which grounds the TEMP HI, switching it on together with a warning chime.

Pressing the TEMP HI light caption unbalances the resistor bridge to simulate a high temperature.

The reservoir low-pressure switch grounds the AIR PRESS LO light circuit to put the warning light on in the flight deck.

The amber LO AIR PRES light illuminates when the pressure drops to about 1.5 bar (22psi).

1 C:leanliness is of tQe utmost importjance. The filters fitted in the 'Arcraft system I will normally protect the components from the effects of partidle contamination,

I / blut it is important that any ground equip,ment used for servicing purposes is kept scrupulously --- clean, and that thefluidis - filtered to a similarl-stddard. --

- - MAINTENANCE -- , I -

All maintenance is carried out 4 accordance with the AMM q d , the maintenance schedule! If it is routine maintenlance then thk specified tasks! T e carded out a t tlie times specified in the approvedl maintenance schedule. If it is non-planned maintenance such as fault rectification-tlien , it is carried out followinglp~ocedures

C:ontamination from other fluids must also be avoided, and provision is usually made for taking fluid samples.

- - published in the AMM. i i

',

Whenever a connection is broken or a component is removed, precautions must be taken to prevent the ingress of foreign matter or moisture. If it is necessary to top-up the system, fluid should be poured directly from a new fluid container into the reservoir, or a sealed dispensing rig should it be used.

1 - - /

When the system is topped-up from a can, any unused fluid in the can should be discarded or poured into a dispensing rig - unsealed cans with fluid in them should never be left (the fluid could become contaminated and be used in an stircraft system at a later date).

- 77 -

rnoodull l A-1 140

Sampling

Samples of the system fluid should be taken at the periods specified in the approved Maintenance Schedule, and whenever contamination is suspected.

If a fluid sampling kit is available it should be used strictly in accordance with the manufacturer's instructions, but if such a kit is not available, the sample should be sent to a laboratory for examination.

The bottle into which the fluid is drained must be scrupulously clean, to avoid any contamination that may be already present in the sample. The bottle should be washed with soap and water to give a clean, bright finish, rinsed in clean water, then in filtered alcohol, and dried with clean dry air.

It is usually recommended that plastic sheet is interposed between the bottle and the cap, to prevent the formation of loose particles when the cap is screwed on.

When-taking a sample, a suitable sefice should be operated to circulate the fluid, and aisrnall quantity should be'\,dl'alii~d from the sampling point before filling the sample bottle.

I I I

Precautions should be taken to prevent contamination of the sample, and instructions in the AMM, or in the test kit should be carefully followed.

The aition which is necessary to d&e following the testing of a sample of fluid, will ddpend on the degree of contahination found. The parameters to be tested are - acidity, spkcific gravity, viscosity, water 'content, and particle contamination, and acceptable - values are specified - in - the appropriate AMM.

- --

If slight contamination is present, the fluid should be circulated by operation of the services, and a further sample taken. If heavy contamination is found, the affected system should be flushed or drained, and re-filled with clean fluid.

Flushing

Flushing is normally required after extensive replacement of pipelines or components, and is carried out by operating the particular service a number of times, so that any particle contamination may be trapped by the filters.

When it is necessary to flush the main system, the filters should be changed and the fluid should be circulated by operating the largest hydraulic jack a number of times. Either an auxiliary pump, or an external hydraulic test rig may be used for flushing, but, if an auxiliary pump is used, it is recommended that it is subsequently removed and inspected for possible damage.

During flushing always ensure the reservoir is kept topped-up.

Draining

The hydraulic system should be drained (or partially drained) whenever components which are not provided with self-sealing couplings have to be removed; and also when overheating occurs (above a certain value), or when there i:s extensive mechanical failure of a pump or similar complex component; or there i:s introduction of extraneous fluids or foreign matter.

Make sure that all driven pumps are configured so that they cannot be operated before commencing draining. If any of them were to be run dry (ie without fluid) then serious damage would occur.

The hydraulic system should be made electrically safe (by tripping off circuit- breakers or the removal of fuses, as appropriate). The hydraulic pressure should be released by operating one of the services or operating the pressure release valve, and the air pressure should be released from the accumulators and

\ reservoir. The reservoir filler cap -- should be removed, and fluid should be drained into a cleari container of suitable capas@; by means of the Lystem drain cock. , I- -

Drained fluid should be returned, in apprdpriately identified dontainers, for reclamation by an approved process. I I

I 1 I ' ' ,/ I ' I

If fluid cdntarnination is the r eadonfo~drah in~ , it will also be: &cessary to remove the filters, and to clean or replicithk fdter elements As appropriate. Cleaning is usually by an ultrasdnic cleanirtg brocess, but wabhing in trichloroethylene,,'r&ay also be petmissible as 2 temporaq rneas4re.

I 1 I I I

If the coritarnir;/ation is found to belfluid t&t/has an adverse dffect on the seals of the s$stem-components, then t'hese-components may requirk replacement - check the AMM or contact the manufacturer. If particle contamination is severe the source must be found and the component changed (if it is component break-up).

Internal Leak Testing

Some circuits/components will leak internally - from the pressure line to return. This applies to those components where valving employs sliding members that rely on metal-to-metal contact for sealing purposes such as PFCUs.

The leak-rate can be checked and if it is found to be higher than that laid down in the manual then the source of the leak must be found and the component replaced; and if an external leak, it must be found and rectified - and the leak- rate re-checked.

To check for leaks connect a hydraulic ground power unit equipped with a sensitive flowmeter and pressurise the circuit to the correct pressure as per the fiMM. Set the PFCU servo control selector to OFF to isolate the flight controls.

moodull l A-1 142

Close all shut-of valves - note the fluid flow.

Open the centre shut-off valve to pressurise the tail section of the aircraft. This will pressurise the following:

* The rudder and elevator PFCUs. -k The tailplane trim (horizontal stabiliser trim). * The yaw damper, etc.

Note the leak-rate flow and compare it to the nominal rate stated in the AMM.

To test the wing areas successively open the right-hand and left-hand shut-off valves and compare the two leak-rate flows to the laid down in the AMM.

If a leak above the nominal is located, conventional trouble-shooting procedures are carried out to isolate the fault. The component is changed if internal, if the leak is external it is rectified. Re-test system.

-- - - -

\ -- - - i

Fig. 75 LEAK-RATE TESTING OF THE TAIL SECTION OF THE A 3 0 0

Internal Leak Management

Some aircraft are equipped with an internal leak management system which is used to detect abnormal internal leaks in the PFCUs.

rnoodull l A-1 143

System Filling

Filling may be carried out through an open reservoir filler neck, or through a priming connection in the ground servicing bay using an external priming rig, or through a common self sealing coupling with a selector valve to fill anyone of several reservoirs.

The system is pressurised for priming purposes by using either an aircraft e:lectrically operated pump; an external hydraulic test rig; or hand pump.

Bleeding

To ensure correct operation of the system all air must be removed from the pipelines and components. Some components are bled by slackening the pipe connections allowing fluid to escape then retightening. Some components are fitted with bleed valves/nipples, and others are purged by operating - the service and {orcing'any trapped air to return toth'e,reservoir. Sometimes -- called self- bleeding systems, they often have to beoperated several times to purge all the air out of the system to the reservoir +d to athosphere. 1 l~

I

The airkraft should be jacked in accordance with the AMM an;d the accbmulators s houldbe charged with nitrogen to-the c6rf6ct pressure. Ground- electric& power shouldlbe connected and the apprdprixe fluid level and pump overheat warning lzunps should be tested. 1 I I

I I ' I I ' I n general the fohowing points sdould be observed, but always refer to the AMM:

I L -/

/ _- / i

- 1. -- Use-the aircraft e16ctrically-driven pump or hand pump. 2. Keep the reservoir topped up and pressurised if of this type. 3 . Start bleeding from the lowest and most distant point in the system -

if all the system has to be bled. 4. Always bleed from the small volume end of the jack - ie the end that

the piston is at. 5. Use a clear plastic hose placed over the bleed nipple with the free end

immersed in a glass jar half filled with the same fluid. 6. Using the hand pump slacken the bleed nipple until an air-free-flow

of fluid is obtained from the end of the hose. 7. Tighten the nipple and allow the piston to be pumped to the other end

of the jack. (The jack may have to be disconnected from the component it moves to allow access).

8. Bleed this end of the jack as in 5 to 7 above, and reconnect jack if necessary.

9. Carry out this procedure on all jacks/ actuators/ bleed connections in the system or circuit affected.

10. Place components - landing gear, flaps etc in correct position and top-up reservoir to correct level.

rnoodull l A-1 144

1 1. Carry out functional checks. 12. Check for leaks, wire lock all bleed nipples/connections used during

bleeding, clear log-book/work cards and sign CRS.

NOTE. The fluid obtained from the bleeding process should not be re-used.

Checking Fluid Level

In general the following points should be noted:

1. Check hydraulic system pressure. With many aircraft this has to be exhausted - the operation of a service such as wheel-brakes or the depressurisation valve will relieve the pressure.

2. Check that the gas pressure of the accumulators is correct. 3. Check that all the services are in the correct position - usually

---- landing gear down and all other services in or up (ie flaps up, spoilers '

- - -

in etc). I- ,

. Check reservoir level v d t o p up if necessary. This may involve I I topping up through /& open fill& neck, or using a pressurised filling I I

1 rig via a self sealinglconnection. 5. / Check flight-deck level indicators show correct reading - as well as

, 1 direct reading indicitors (withla pressurised reservoir check if the I pressure should be pn).-- , I

1 i \

Note. If excessive amount of flulid is required to top-up to the correct level then investigate the reasons why - add rectify.

I

- -- -

Component Replacement

Components are replaced when they become unserviceable or when they are life expired, or when required in the maintenance schedule. When carrying out the task check the following points:

1. Check the AMM for the correct procedures. Check IPC for correct replacement identification.

2. Depressurise the system/component. 3 . For an electrically operated component isolate that electric circuit. 4. Check that a correct replacement is available - check description

part numbers, mod state, and life remaining - if applicable. 5. Check 'new' item against stores release certificate (EASA form 1). 6. Drain/ shut-off fluid. 7. Jack aircraft if necessary. 8. Remove 'old' component.

rnoodull l A-1 145

9. Remove blanks from 'new' component; inspect for general condition and fit. Secure pipeline connections and connect any electrical connections; inputs form control systems etc.

10. Return old component suitably labelled, blanked etc to stores. 1 1. Prime/ bleed system. 12. Carry out a functional test, record and sign relevant paperwork.

System Testing

E;!efer to the AMM. One circuit only may require testing but the points listed below assume that the whole system is to be tested.

Jack the aircraft. Check system is complete - reservoir levels are correct and accumulators changed to correct gas pressure. Connect external electrical power.

-- -

Insure flightdeck selectors - - - are in the same position as their respective systems. Carry out any BITE checks. Use aircraft electrically driven pumps (with external electrical power) or connect external hydraulic test rigs. Quick release self seal external connections are as follows: /

* Constant volume pumps. All 'aircraft - 2 - one suction, one pressure. I

* Constant pressure pumps. Some aircraft - 3 - one suction, one pressure, one idlinglcase drain. Other aircraft1- 2 - one suction, one pressure.

Where applicable check each circuit for: (a) Smooth operation. (b) Range of movement. (c) Leaks. (d) Flush fitting doors - landing gear, cargo etc. (e) Correct clearances. (f) Correct sequence - eg doors and legs on landing gear. (g) Correct inter-relationship of one circuit to another (eg power

steering centres on undercarriage up-selection, differential spoilers work with aileron movement).

(h) Correct flightdeck indications - position - temperature - pressure etc.

(i) Timing of each operation. (j) Correct warnings, config etc given on flight deck. (k) Correct sense (eg flap lever UP, flaps UP). (1) Synchronous operation (of flaps). (m) Correct operation of stand-by systems, cross-feed supplies etc. Reconfigure aircraft to original configuration. Record all work carried out on the aircraft and sign for the work done.

rnoodull lA-1146

Generally pumps can be tested:

x Engine driven pumps - during an engine run. * Air driven pumps - during an engine run. * Electrically driven pumps - with external electrical power. * RAT pump - with external hydraulic test rig connected. * Power transfer units - with power source system running. * Hand pump - when driven pumps are off.

rnoodull l A-1 147

CONTENTS

Page

Ice formation on aircraft 1 Ice detection 4

Spot lights 5 Black rod ice detector 5

i Vibrating rod ice detector 6 Pressure operated ice detector - 9 Rotary ice detector

- 11 Inferential method 12 Beta particle ice detector 13

Methods of ice protection 15 Electrical airframe ice protection --- systems 17

Windscreen ice protection 20 ~ indsc rken wash & rain repellent systems 26

Windscreen wipers I 28 Fluid airframe de-icing systems 31 F'ropellers - - 34 F'neumatic airframe de-icing systemms 37 Hot air de-icing systems 41 Other anti-iced areas 47

rnoodull l A-1 148

ICE FORMATION ON AIRCRAFT

Icing on aircraft is caused by a combination of freezing conditions and moisture in the atmosphere. If moisture impinges on the forward facing surfaces of a n aircraft and those surfaces are a t or below O°C, then a build up of ice will be caused which may seriously alter the aerodynamic qualities and increase the weight. This applies particularly to small objects, which have a higher catch rate efficiency than large ones, as small amounts of ice will produce relatively bigger changes in shape.

The actual amount and shape of the ice build-up depends on surface temperature, which results from a n energy balance arising from heat input from viscous or kinetic air heating, hnetic heating by water droplets and the latent heat of fusion, and losses from evaporation or sublimation, convection and by warming the impinging droplets.

When the temperature is less than 0°C all the impinging water droplets are frozen, and when it is above 0°C none are frozen. However, for a particular set of atmospheric conditions and altitude it is found that there isquite a wide aircraft speed range over which the energy balance gives a skin temperature of 0°C and ,this energy balance occurs a t one end of the speed! range by all the droplets freezing and at the other by none freezing. The potential "catch rate" or "impingement rate" and the actual_ icing rate are thus not simply relatcd in this region. I

I -

The "no icing hazard" speed depends, therefbre, on the free wker content of the atmosph,ere as well as the temperature and altitude. For severe conditions it is about the-mmaximum speed of subsonic aircraft. The final influencing factor of note is that icing does not occuiabove about 12,000m (40,000 ft) since the droplets are all frozen and in the form of ice crystals and will not adhere to the aircraft's surface. Relative humidity is also very low.

It is interesting to note that icing can occur within the temperature ambient range +25"C to -15°C depending on the energy balances stated above. Outside this range ice will not form.

Hoar Frost

Hoar frost occurs on a surface which is a t a temperature below the freezing point of the adjacent air and, of course, below freezing point itself. It is formed in clear air when water vapour is converted directly to ice and builds up into a white semi-crystalline coating. Hoar frost is feathery and is the white crystalline deposits that can be seen on trees and hedgerows after a frost. I t is easy to scrape off car windscreens.

When hoar frost occurs on aircraft on the ground, the weight of the deposit is unlikely to be serious, but if not removed may interfere with the airflow and attainment of flying speed during take-off, the windscreen may be obscured, and the free working of moving parts such as flying control surfaces may be affected.

Rime Ice

This ice formation, which is less dense than glaze ice, is an opaque, rough deposit. At ground level it forms in freezing fog and consists of a deposit of ice on the windward side of exposed objects. Rime is light and porous and results from the small water drops freezing as individual particles, with little or no spreading.

Aircraft in flight may experience rime icing when flying through clouds with the air temperature and the temperature of the airframe below freezing point; thc icingbuilds u p on the leading edge, but does not extend back along the chora. Ice of this type usually has no great weight, but the danger of rime is that it will interfere with the airflow over wings etc, and may choke the orifices of the carburettor, air intake and Pitot static probes

I t is difficult to scrape off car windscreens.

I

Glaze Ice I '

Glaze ice is the glassy deposit that forms over the village pond after a frosty - - - -

night. On aircraft in flight, glaze ice forms when the aircraft encounters large water drops in clouds or in freezing rain with the air temperature and the temperature of the airframe below freezing point.

I t consists of a transparent or opaque coating of ice with a glassy surface and results from the liquid water flowing over the airframe before freezing; glaze ice may be mixed with sleet or snow. It will form in greatest thickness of the leading edges of aerofoils and in reduced thickness as far as one half of the chord length.

Ice formed in this way is dense, tough and sticks closely to the surface. It cannot easily be shaken off and, if it breaks off at all, it comes away in lumps of a n appreciable and sometimes dangerous size. The main danger of glaze ice is still aerodynamic, but to this must be added the dangers of increased weight, unequal wing loading, propeller blade vibration and ice shedding debris damage to tailplanes, rear engines etc. Glaze ice is the most severe and the most dangerous form of ice formation on aircraft.

Pack Snow

Normally, snow falling on an aircraft in flight does not settle, but if the temperature of the airframe is below freezing point, glaze ice may form from the moisture in the snow. The icing of the aircraft, however, is primarily due to water drops, though snow may subsequently be embedded in the ice so formed.

Conclusions

I t if ice continues to be deposited on the aircraft in flight, one or more of the following effects may occur:

(a) Decrease in lift may occur due to change in aerofoil section resulting in loss of streamlined flow around the leading edge and top surfaces.

(b) Drag will increase due to the rough surface, especially if the formation is rime ice. This condition results in a greatly increased

I

s h n friction and thicker boundary layer. I

(c) Decrease in propeller efficiency. With turbo-prop and piston engines, the efficiencyofthe propeller will decrease due to alterations of the blade profile and increased blade thickness.

(d) LOSS of control. LOSS of contrdl may occur due to ice preventing movement of control surfaces. I

-

- - - -

(e) Increased load and wing loading. The weight of the ice may prevent the aircraft from maintaining height, or even taking off.

(f) Loss of the inherent passive stability may occur due to C of G movement caused by the weight of the ice.

(g) Loss of vision - if the windscreen becomes iced up .

(h) Malfunction of flight and engine instrumentation and air data computer. This would occur if Pitotlstatic vents, PI probes and air data probes became blocked.

(i) Debris damage. Caused by sections of ice breaking away from the airframe and hitting tail-planes, rear mounted engines etc. This has caused serious accidents in the past.

Ij) Drain systems (grey water) becoming blocked.

(k) Icing of intakes, carburettors, cooling ducts etc causing reduction in efficiency in engines, engine failure, overheating of systems.

(1) Flutter. Ice can change the pattern of the airflow around the control surfaces, mainplanes etc, a s well as change the centre of gravity of the component - thus allowing flutter to occur.

It is evident from the above that the airframe and systems prone to icing be kept ice free both before and during take-off (ground de-icing) and during flight (aircraft de-icing). Ground de-icing is carried out by maintenance personnel prior to aircraft take-off and aircraft de-icing is carried out by on-board systems.

For ground de-icing refer to module 7 in the LPB series.

ICE DETECTION

EASA25 states that if an aircraft is to fly in icing conditions then it must be proved that it has all the necessary equipment on-board to be capable of flight in a safe manner.

I

Most detectors detect the actual ,presence of ice on the airframe, some (rarely) measure the two parameters required for icing to occur (moisture and temperature) and infer that icing conditions exist.

I

Many different methods are used to detect the presence of ice. The actual method used is dependent on the aircraft type and manufacturer. With some aircraft there is a main detector system (often electrical/electronic) with a back-up system (often low-tech) .

Ice detection systems use one of the following methods of detecting and assessing the formation of ice:

(a) Ice Accretion Method. Ice is actually allowed to accumulate on a probe which projects into the airstream and in doing so operates a warning system.

(b) Inferential Method. This method uses two detectors, one to detect temperature and the other to detect the presence of water droplets or moisture. So ice is not actually accreted but that it is inferred that it might from the two measured parameters. Used in wind tunnel testing but rarely used on aircraft.

We will concentrate on ice accretion detection methods in this book.

rnoodull lA-1152

Ice Formation Spot Lights

Ice formation spot lights are mounted on either side of the fuselage so as to light up the leading edges of the mainplanes in the dark. When required they can be switched on to allow for visual examination. They can, of course, be used at any time by the flight crew to visually check that part of the aircraft for any signs of damage, malfunction etc, not necessarily connected with ice.

Fig. 1 ICEFORIMATION SPOT LIGHTS

Black Rod or Hot Rod Ice Detector I

This consists of an aluminium alloxbase on which is mounked a black steel mast detector of aerofoil section, angled,back from the verti'cal and mounted on the outside of the fuselage so that it can be seen from the flight-deck. The detector houses a heating element, and in the base there is a built-in lamp.

/ ICE DETECTOR

TERMINAL BLOCK

Fig. 2 HOT ROD ICE DETECTOR

The heating element is normally off and when icing conditions are met ice accretes on the leading edge of the mast which can be seen by the crew as it turns white. During night operations the built-in lamp is switched on to illuminate the mast. Once the mast turns white it is de-iced by switching on the heater element. This clears the ice, the mast turns black and the heater element is switched off.

QUESTION How would you check this unit for correct operation? (5 mins)

ANSWER Carefully. A little more seriously though - check for damage, security of attachment and corrosion. Check lamp for correct operation. Check operation of heater (carefully) with the fingers and check that the mast is coloured black.

Vibrating Rod Ice Detector -

-

When ice is detected it provides an indication in the flight deck. The system consists of a solid-state ice detector and warning system. The ice detector is attached to the fuselage with it's~probe protruding into the airflow. The ice detector probe is caused to vibhte a t it's resonant frequency of 40kHz. [Resonant frequency is affected by (a) the Forcing Function and (b) the Mass of the object being vibrated - the greater the mass the lower the frequency].

ICE DETECTOR 2 Fig. 3 VIBRATING ROD ICE DETECTOR

When ice forms on the probe, its mass increases and the frequency drops. The ice detector circuit detects the change in probe frequency by comparing it with a reference oscillator. At a predetermined frequency the ice detector circuit is activated. Once activated the probe de-icing heater is triggered to de-ice the probe and return it's frequency to normal. The heater then turns off, to allow the probe to ice up again (if icing conditions still exist).

rnoodull l A-1 154

This cycle is counted by the system and after a certain number of cycles an ice warning signal is sent to the flight deck, and certain de-icing systems are switched on and/or a signal is sent to a computer - depending on the aircraft.

This cycle will be repeated for as long as the icing conditions exist and when icing conditions cease the system re-sets itself, turning the ice attenuation lamp off in the flight deck, and deactivating any de-icing systems that it have switched on.

QUESTION How would you check this unit for correct operation? (5 mins)

ANSWER In general the following points apply:

(a) Check the AMM. (b) Operate any BIT systems. (c) Check heater (carefully) by switching ON and touching

- - ---

with fingers. I

(d) Holding the probe using "oven gloves" while the unit is switched on will cause it's frequency to drop and the system to operate.

(e) Letting go of the probe (as the heater comes on - -

automatically) will allow the warning to cancel (after a delay period). -

1 I

The following describes the system as fitted to the B747-400.

Two independent ice detection systems provide flight deck indication of icing conditions and automatic activation of engine intake thermal anti-ice systems. Components used by the ice detection systems are; 2 ice detector probes; the nacelle anti-ice control switches on the anti-iceIrain removal module, and the Engine Indication and Crew Alerting System (EICAS) display. (A CRT - Cathode Ray Tube - type display on the flight deck).

The left hand ice detection system is powered by 115V ac from the 115V ac bus 1 and the right system from the 115V ac bus 3. The ice detector probes are located on the lower front part of the fuselage just aft and below of the Pitot static probes.

The ice detector probe contains a sensing element, a heater and control cards. The sensing element is part of an electromechanical resonant circuit which oscillates a t approximately 40kHz. Ice build up on the end of the element lowers the resonant frequency by increasing the effective mass of the element. After a specific amount of ice has accumulated, the heater is activated to melt the ice and return the resonant frequency to its initial value.

rnoodull l A-1 155

If the heater cannot melt the ice after 25 seconds, or if the internal temperature of the ice detector exceeds 300°F (150°C), the heater and the detector outputs are automatically de-activated.

Each ice detector provides an output to activate the corresponding left or right side engine anti-ice systems and a separate output for indication of wing icing conditions. Computations for identifying icing conditions and for activating anti-ice systems are performed in the ice detector.

Control of the anti-ice system is via the Anti-Ice/Rain Removal Module on the flight deck. The module includes 4 nacelle anti-ice switches. Each switch has 3 positions to place the corresponding anti-ice system ON, OFF or in AUTOMATIC mode under control of the ice detection system.

A s ice builds u p on the ice detector sensing element during flight, the resonant frequency of the sensing element decreases. When a 0.02in (0.5mm) ice layer accumulates, the heater turns on. When the resonant frequency returns to if ini6d value, the heater turns off, about - 8 seconds after it has turned on. This process is repeated as long as icing conditions exist. The ice detector keeps count of the number of icing/de-icing cycles of the sensing element.

I I

When 2 consecutive cycles are counted, the detector outputs a signal to activate the engine anti-ice system for about 3 minutes. IE an additional icing/de-icing cycle occurs during this time interval, the 3 minute engine anti- ice activation period is begun again. If it takes more than 15 seconds to melt the ice ,on the first icing/de-icing cycle, the ice detector will activate the engine anti-ice system for 3 minutes without waiting for a second cycle.

When 10 consecutive probe icinglde-icing cycles are counted, the detector outputs a wing icing signal to the EICASIEFIS interface units (EIU's) for 3 minutes. If an additional icing/de-icing cycle occurs during this time interval, the 3 minute wing icing cycle is begun again.

If the wing or nacelle anti-ice control switch is placed in the OFF position then a message is sent to the EICAS display, the same happens if a control valve fails to open. If the wing or nacelle anti-ice control switch is placed in the O N position and the total air temperature exceeds 12°C a message is also sent to the EICAS display. If a n ice detector probe fails a message is sent to the EICAS display.

When the aircraft is on the ground, the icing/de-icing cycle is held a t zero so that there is no automatic activation of anti-ice systems. However, the ice detector heater is still enabled on the ground.

moodull l A-1 156

The ice detection system can be tested in 2 ways: on the ground using the Central Maintenance Computer (CMC) and doing an operational test.

CMC Test. For this test electrical power is applied and the CMC does a test of the ice detection system. If the system fails then it will produce a list of message pages which include CMC messages, CMC message numbers and relevant ATA chapter numbers. The CMC Message Index in the fault isolation manual shows the corrective action for each CMC message.

Operational Test. It is important for this test that the engines must not be running. Electrical power must be on after confirming that it is safe to do so and the aircraft must be put into the Air Mode in accordance with the AMM.

Hold your hand near each of the ice detector probes to make sure that they are not hot. Set the nacelle anti-icing 1, 2, 3 and 4 switches on the anti-iceIrain removal module to AUTO.

To check the engine anti-ice systems for the left ice detector, put on a heat resistant glove and apply preisur6 to the probe tip with the th-rnb and forefinger for one or two seconds., Release, the probe tip for 15 seconds. Apply pressure again to the probe tip ,with thumb and forefinger for one or two seconds. Release the probe tip.

I

-

Make sure that the EICAS caution &&sage. >ICING NAC shows for about 3 1 1 minutes: It is advisable to use a wet cloth or a piece of ice to cool the probe

strut (ndt the probe tip) while you do this test, this helps to increase the life of the probe. Repeat the same test for the right hand detector. If1 necessary run through a "CMC Test" afterwards in accordance with the AMM.

- - - - - - --

Return the aircraft to the Ground Mode in accordance with the AMM and remove electrical power. Record all work done and sign the appropriate documentation.

Pressure or Pneumatic Operated Ice Detector

This consists of a short tube, which is closed at its outer end and mounted at right angles to the airflow and connected to a case. Small holes are drilled in the leading edge and in the trailing edge of the tube. Dynamic air enters the holes at the leading edge and passes into the case returning to the airstream via the holes at the rear of the tube. A heater element is fitted in the tube to allow it to be cleared of ice (once formed).

Also there is a thermostatically controlled anti-icing heater in the case that is on all the time to prevent it icing up.

WARNING SIGNAL ELECTRICAL SUPPLY

ELECTRICAL SUPPLY

\ I \I [ 1 1 CASE HEATER

I I I I I I I I I

I

OVER-HEAT SWITCH SWITCH

DIAPHRAGM

. CASE

CALIBRATED ORIFICE

FLOW OF AIR

PlTOT HOLES

I

I

' I , Fig. 4 PRESSURE OPERATED ICE DETECTOR I 1

I I

I

In flight, Pitot pressure will build up inside the case and push the diaphr down. This will keep the switch in the position as shown in figure 4.

- - - - -

When ice accretion builds u p on the leading edge of the probe the Pitot pressure is cut off. The air pressure now drains away through the calibrated orifice and the rear static vents. The natural springiness of the diaphragm will cause it to move up and the switch to close - thus giving a warning signal to the flight deck.

The switch operation will also cause the heater element to come on in the probe - thus de-icing the probe and resetting the system. This cycling will continue until icing conditions no longer exist.

A delay may be built into the warning system to keep it activated until after one complete clear cycle of the system - this prevents the flight deck warning lamp from flashing on and off during the cycling process.

The system is connected to the weight switch or an airspeed capsule.

It is not fitted to many aircraft.

QUESTION Why is the unit connected to either a weight switch or airspeed switch? (5 mins)

ANSWER If it were not de-activated as the aircraft slows down it would give an ice warning as the Pitot pressure drops. So it will be de-activated either by a weight switch or an airspeed capsule some time during the landing approach/ landing run.

QUESTION How would you test the unit? (5 mins)

ANSWER Refer to the AMM. The probe heater can be checked with the fingers - carefully. The internal heater may be felt from the outside of the case or be checked using electrical test meters. Of course, damage, corrosion etc to the unit and probe can be checked visually. The functional test would require the use of a special adapter and the supply of Pitot pressure. The system would also have to be configured so the unit thinks the aircraft is in the air (operate weight switch or airspeed switch). -

Rotary Ice Detector (or Napier Ice Detector) I

The Napier Servomechanisms ice detector consists of an ele'ctrically driven, externally-splined rotor revolving close to a fmed knife edge cutter. Clearance between the rotor and cutter is 0.'002in (0.05mm). The rotor is mounted with it's axis at right angles to the direction of thd airflow.

AIRCRAFT SKIN

KNIFE EDGE CUTTER

/ ROTOR

Fig. 5 ROTARY ICE DETECTOR

- 11 -

rnoodull l A-1 159

Any ice accreting on the front of the rotor is scraped off as it rotates to the rear where the cutter is. Ice scrapping starts a s soon as the thickness exceeds the clearance between the rotor and the cutter.

The scraping action on the ice produces a torque reaction on the motor. This is mounted on a spring loaded gimbal and the torque reaction causes it to rotate a few degrees in the opposite direction to that of the rotor. This action closes a micro switch and sends a warning signal to the flight deck.

While ice continues to accrete the torque reaction will remain and the warning signal stays on. When ice accretion ceases half a revolution later (of the rotor) the scraping ceases - torque reaction ceases and the motor (within its housing) gimbals back to it's original position, (under the influence of a spring). This opens the micro switch and the flight deck warning is cancelled.

QUESTION How would you check this detector for serviceability? (5 mins) - -

I -

ANSWER Naturally refer to the AMM, but in general, apart from security, i ' corrosion, attachment and lay of cables etc, the unit is checked for:

(a) Operation - switch it on and see that it rotates. (b) As it rotates place a screw driver in the slot at the end if the

rotor and gently hold the handle in your hand. The torque reaction on ,the rotor should cause the warning on the flight deck to operate.

Moisture and Temperature Detector (Inferential Method)

Ice can only be formed when two conditions exist: (1) moisture is present and (2) the temperature is a t or below freezing. With the inferential method these two parameters are monitored separately.

The system is not used widely on aircraft but may be found in wind tunnel testing equipment.

The temperature and moisture detectors are wired in series so that they both have to show positive for a warning to be given. But note - no actual ice has been detected.

The temperature detector can be a contact thermometer or any other temperature detector. The moisture detector consists of two heated elements wired as a bridge. One element is open to the airflow, the other one is protected.

When moisture occurs the exposed element cools due to the latent heat of evaporation being drawn off. This changes its resistance and the bridge becomes unbalanced which causes a relay to close a switch. When moisture ceases to occur the temperature of the exposed element rises to that of the protected element - the bridge balances and the relay opens the switch.

So when the bridge is unbalanced (due to the presence of moisture) and the temperature switch is made (due to temperatures at or below freezing) then a warning is given.

Beta Particle Ice Detection Probe

This particular system was to be fitted to the A320 but so far there has been no experience of this. I t may be fitted to some aircraft.

Principle -

Beta particles are absorbed by ice-so if one probe emits beta particles the other probe will detect them. If the probes get iced up then the beta particles will get absorbed and the detector probe will detect fewer beta particles. At a certain beta particle count rate a relay in the detector probe will operate causing a warning to come on in the flight deck. The system is reset by a heater element. The system is tested by using a strip of Teflon wrapped around one of the probes - th i s will-absorb the beta particles and cause a warning to come on and the de-icing system to work. In general:

(1) -The probe protrudes perpendicularly from the fuselage wall. It-is installed in the forward fuselage in a zone where the local liquid water content in icing conditions is always greater than that in the upstream ambient air.

(2) The beta particles are emitted from a Strontium 90 Sulphate capsule (maximum activity = 50 microcuries).

The radiation counter is a Geiger tube installed in the probe base. The beta particle source and the count rate are located on either side of the sensing surface so that the thickness of ice on the leading edge can be monitored through two opposite facing windows.

The radioactivity of the probe is, at a distance of 41cm (16 in) roughly equivalent to that emanating from a luminous wristwatch. The fuselage wall forms a screen which provides efficient protection for occupants. However, for ground handling of the probe during maintenance, some safety precautions should be observed.

rnoodull lA-1161

The ice thickness sensed by the probe sensing surface increases with the combination of the following parameters:

* Cloud liquid water content. * Aircraft speed in icing conditions. * Ice detection duration.

When the ice thickness increases, the pulse rate of the Geiger tube measuring the attenuation of the beta particle stream decreases. The triggering threshold corresponds to 0.4mm (0.0 15 in) of ice on the sensing surface of the probe. When the triggering threshold is reached, the internal relay is energised and causes simultaneously:

* Activation of the ECAM system and illumination of annunciator lights on the overhead panel.

* Probe de-icing.

(3) I, The probe also incorporates a de-icing electrical circuit and an overheat protection device'which trips if probe de-icing operates for :excessively long periods. De-icing power is 5A with 115V ac.

(4) 'The particle count signal is fed from the probe to the controller via a ' c6axial cable.

I I

I

Indications are as follows:

Repetition warning: ICE legend of ICE DET push button switch. Central warnings (ECAM system): ANTI ICE warning light on the warning light display panel, single chime, display on the left ECAM display unit.

The response action consists in supplying power to the engine air intake and airframe ice protection systems. The ICE repetition legend on the ICE DET pushbutton switch goes off 60 seconds after the last ice detection signal.

Maintenance Panel

(1) The system can be tested from the flight compartment by pressing briefly the ICE/TEST pushbutton switch. The test is positive if momentary action on the ICE DET/TEST pushbutton switch triggers the ice warning (ANTI ICE warning light on the warning light display panel, single chime, display on the left ECAM display unit). When the pushbutton switch is released (out), this cancels the ice warning: The ANTI ICE warning light on the warning light display panel and the ICE legend of the ICE DET pushbutton switch go off.

moodull l A-1 162

(2) Test of the system can be performed on the aircraft on the ground by simulating ice thickness. A Teflon strip is temporarily wrapped around the sensing surface of the probe. A 0.13mm (0.005 in) thick strip does not cause ice warning activation whereas a 0.52mm (0.020 in) thick strip does.

METHODS OF ICE PROTECTION

Ice protection for aircraft can be by a de-icing system or an anti-icing system and is necessary for:

Aerofoils. Tailplane (stabiliser) and fin. Intakes - engine and APU. Propellers. Spinners. Control surfaces. -

Windscreens. Detector probes - TAT, Pitot, engine EPR, angle of attack etc. Slats. Carburettors. Drain masts (grey water and potable water masts). Waste water connections, escape slide attachments etc.

WATER SERVICE - - - - PANEL HEAT

LOCKING MECHANISM

AOA SENSOR (ON ALL DOORS)

WINDSHIELD

(ON UNDERSIDE) ENGINE ANTI-ICE

ANTI-ICE

TAT PROBE HEAT

PITOT PROBE HEAT

Fig. 6 AREAS TO BE DE-ICED/ANTI-ICED - A 3 3 0

- - 15 - -

rnoodull l A-1 163

Not all these systems will be fitted to all aircraft but you have to be aware that they do exist and how they work.

De-icing implies the removal of ice after it has formed and anti-icing refers to the prevention of ice formation. Some equipments can be de-iced - mainplanes for example - where any removed ice (by the airflow) cannot damage any other part of the aircraft. Engine intakes are usually anti-iced as de-icing might cause lumps of broken-off ice to damage the compressor.

Anti-icing/ de-icing Methods

Method Application Principle

FLUID

_ I

Wings, tail units, propellers, windscreens.

Wings, tail units.

A chemical which breaks down the bond between ice and water and can be either sprayed over the surface, eg a windscreen, or pumped through porous panels along the leading edge of a surface eg a wing, or allowed to flow by centrifugal force along the leading edge of a propeller.

I

Sections of rubber boot along the leading edges are inflated and deflated causing ice to break u p and with the aid of the air-stream break away.

THERMAL

Hot air bleed Wings, tail units, Hot air from turbojet engine engine intakes, compressor, (rarely exhaust, via a windscreens, slats. heat exchanger) passed along inside

of leading edge structure or directed on outside of windscreen.

Combustion A s above. heating.

Hot air from a separate combustion heater or from a heat exchanger associated with a turbo prop engine exhaust gas system.

continued

- 16-

rnoodull l A-I 164

Electrical heating.

Wings, tail units, Heating effect of electric current intakes, (dc or ac) passing through wire, propellers, rotor flat strip or film type elements blades, windscreens, attached to the outside of the drain masts, component, or heater elements Pitot heads, embedded within the structure probes, drain masts, (composite materials, windscreens). control surfaces (Dash 8 elevators), carburettors.

Hot oil. J e t engine nose cone. Engine oil heating - which helps cool the hot oil.

De-icing Materials

The various fluids used for frost and ice protection and for de-icing aircraft on the ground are inflammable a n i s m e t i m e s poisonous. Care must be taken when handling. Use only the correct type of fluid for the s~stem/application as laid down in the AMM chapter 20 Maintenance Practices (typically for airframe de-icing systems TKS R328 or any fluid to specification DTD 4068).

Some airframe de-icing fluids will have an adverse effect on laminated windscreens. When topping u p thk windscreen de-icer bottle ensure that it is the correct fluid. -

ELECTRICAL AIRFRAME ICE PROTECTION SYSTEMS -- --- -- -

Heater elements can be:

* Sprayrnat system - leading edges, intakes. * Heater mats - Intakes, propellers and spinners.

Each mat is designed for a specific application, the heat output being obtained from whatever electrical source is available. Mats are available both for anti- icing or de-icing and are rated in watts per m2 (W/m2) or kW/m2.

Anti-icing mats for intakes, are supplied continuously with electricity, while the de-icing mat is intermittently heated. The total area to be heated is often divided into several smaller areas with independent mats for each. The electrical power is then arranged to be switched to each small area in turn.

Thus, on any particular area, there is no heating for a given period during which the ice builds u p and then, when power is switched to that area, adhesion is broken by heat and the ice removed by the airflow. (Non intake applications).

- 17 -

moodull lA-1165

LEADING EDGES

Fig. 7 HEATER ELEMENTS

Spraymat

Spraymat is so called because the heater mats are sprayed directly onto the surfaces to be protected. The surfaces are pre-treated and each spray application is preceded by a masking process. It comprises of:

- r

(a) The base insulator,which is sprayed on to the pre-prepared surface 1 to be protected and is made up of either a rubber or polymer

insulating material.

The conductor element which is sprayed on to the base insulator and is made of either aluminium or Kuminol (copper magnesium alloy).

The outer insulation which is of the same material as the base insulator.

(d) The protective coating ("Stoneguard") which is used when the heater requires extra protection from mechanical damage, eg on leading edges.

Heater Mats

These are used on air intakes on the engine and other components. Heater mats differ in design and construction according to their purpose and environment. The latest mats have elements which are made from a range of alloys woven in continuous filament glass yarn. Other elements are made from nickel chrome foil. The insulating material is usually PTFE which gives a higher limiting temperature than synthetic rubber.

The supplies to heater mats are controlled by units similar to those used in spray mat systems.

moodull l A-1 166

The supply for the heaters can be either ac or dc.

The anti-icing supply is continuous, controlled only by master switching and overheat protection devices. The de-icing supply is cyclic, pulses being fed through a cyclic switch or similar device. The timing of the pulses is determined by the manual setting of a flight deck selector.

1. 200 V 3 0 ac bus bars. 2. Current transformer. 3. Power indicator. 4. Control relay and current balance relays - and power relay. 5. Intakeslaerofoil heaters. 6. Continuously heated element. 7. Cyclicly heated element. 8. Cyclic timer.

9. Propeller elements. 10. Cycling light. 11. Onloff sw~tch. 12. FasVslow switch. 13. Weight switch. 14. To voltage control unit of generator

system. 15. CBs. 16. 28 V dc bus bar.

Fig. 8 ELECTRICAL DE-ICING SYSTEM

System Operation

Figure 8 shows a typical electrically operated anti-icing/de-icing system. It is simplified with the thick line representing the 3 cables of the 3 phase (30) ac supply and u p to 3 control boxes (current balance relay, control relay and overload protection) being replaced by one control box (4).

QUESTION Study it carefully and then explain why the control of the system is via a 28V dc supply using relays, and why the system is not switched directly by the pilot. (5 mins)

rnoodull lA-1167

ANSWER The heavy current carrying cables (30 200V ac) are taken the shortest route from the bus bar to the heater elements (the cables are heavy). This will then require the use of remote switching equipment in the form of relays (electrically operated switches). To route the heavy cables via the flight deck would involve a weight penalty with large switches for the pilot to operate.

When the system is switched on, direct current energises the power relay (4) via closed contacts in the overload sensing device (part of 4 in the drawing) thus allowing the 200V ac to flow directly through to the continuously heated elements and to the timer switch (8). The cyclic timer normally controls all aircraft cyclically heated elements with slow cycle normally selected a t temperatures below about -5°C.

In the event of an overload, the heater elements are protected by the sensing device (within 4) which when actuated interrupts the supply to the power rel;. this= turn interrupts the supply of heating current. The current balance relay (part of 4 or a separate unit) is actuated whenever there is an imbalance between phases.

I

(6) shows a continuously heated element whilst (7) and (9) are cyclically heated. /

For operation on the ground, the applied voltage is reduced to prevent overheating. This is effected by the automatic closing of the weight switch (13) fitted to/ the landing gear.

I

WINDSCREEN ICE PROTECTION

Misting/fogging (on the inside) and frosting (on the outside) can occur on the surfaces of aircraft windscreens. Misting/fogging occurs because of the warm air in the cockpit contacting the cold windscreen and condensing. Moisture and freezing temperatures on the outside will cause ice to form on the outside of the windscreen.

The methods of preventing the formation of mist or ice on windscreens are as follows:

(a) Dry air sandwich (anti-misting). (b) External fluid spray (anti-icing) . (c) Wiper (rain clearance). (d) Gold film - inside outer surface layer of windscreen (anti-icing). (e) Gold film - inside inner surface layer of windscreen (anti-misting). (f) Hot air - outside (anti-icing and rain clearance). (g) Hot air - inside (anti-misting).

rnoodulll A-1 16,3

Dry Air Sandwich

Basically the dry-air sandwich windscreen consists of two transparent panels spaced apart to allow an air gap and sealed by a frame of Vinyl (polyvinyl butyral). I t is a double-glazed windscreen. This method is also used for transparent plastic windows on the fuselage as well as windscreens.

The interspace insulates the inner panel from the outer panel, thus reducing the coldness of the inner panel and helping to prevent condensation. To guard against misting in the interspace, complete dryness is necessary and achieved by electrically-heated elements or more generally by the use of a desiccant such as silica gel.

EXTERNAL INTERNAL PANEL PANEL ASSEMBLY ASSEMBLY

: 1 OUTER GLASS , 2 CONDUCTIVE COATING 3 VINYL , 4 INNER,GLASS -

-- -

Fig. 9 THE DRY-AIR SANDWICH WINDSCREEN OF-THE DC8

The dry-air sandwich windscreen may be:

* Completely sealed. * Fitted with a valve and expansion bag. * Fitted with a tube connected to a silica gel container. * Unsealed.

Expansion bag type. To ensure complete dryness, the interspace is completely sealed except for a valve which seals the space while the windscreen is in storage or being serviced. When the windscreen is fitted to the aircraft, the valve is depressed by the expansion bag connecting pipe, so allowing the interspace to "breathe" into the rubber expansion bag. This 'breathing' is necessary in order to equalise pressure changes due to variation in aircraft altitude. The expansion bag must be deflated before take-off.

moodull lA-1169

Dw air type. The windscreen is fitted via a tube from the interspace valve to a container filled with a desiccant. The desiccant, (silica gel), is in the form of crystals and has the ability to absorb moisture. The crystals, when fresh, are blue in colour, but turn to pink/no colour when saturated with moisture and need to be changed. A Perspex window in the container enables the colour of the crystals to be inspected. They can be re-activated by gentle heating in a oven when they will turn blue.

Unsealed type. Some windows/windscreens that are "double glazed have a small hole drilled through the inner panel to the air gap which allows the air pressure to equalise between the air gap and the inside of the aircraft. Common on fuselage windows.

MANUFACTURER PREFIX

DATE OF MANUFACTURER

ONFIGURATION CODE

PART NUMBER

Fig. 10 TYPICAL WINDCSREEN IDENTIFICATION (B747) I

I

~xterna l Spray System I

Figure 11 shows a system as itted to a fighter aircraft, though all external spray systehs will work using the same principles. The system comprises a de- icing fluid tank connected by a pipe line to an electrically operated pump, the delivery side of which is coupled to a spray device arranged in front of the windscreen.

It is fitted with inlet, vent and supply pipes, together with a small drain pipe. The inlet and vent pipes connect with the fluid tank, the supply pipe to the spray device and the drain to the atmosphere. The pump is controlled by a switch in the flightdeck.

Servicing will include keeping the holes in the spray device clear, cleaning the tank filler-neck filter and refilling or topping-up the tank with fluid of the correct specification. For full details of servicing see the aircraft maintenance schedule and the AMM.

Functional checks include switching the pump on and checking for fluid spray coverage on the windscreen.

PUMP

DOOR

Fig. 11 WINDSCREEN FLUID DE-ICING SYSTEM (HAWKER HUNTER)

Electrical Heat , -

Electrically heated windscreens are used on the majority of modern aircraft, both for ice and mist preventioq and to increase the resistance of the panel against bird strike at very low temperatures.

I

A transparent'film element (Gold 'Film) is fitted between the laminations of the windscreen-and connected to t h e electrical system. The gold-film is thin enough to see through but will take an electric current and act as a heater.

A temperature sensitive element is incorporated in the panel to automatically regulate the film temperature.

QUESTION Where is the gold film heater element positioned in the windscreen? (2 mins)

AI\TS'wER It is fitted inside the front glass layer (anti ice) and one may be fitted between the inside glass layer and the vinyl layer (anti mist). See figure 12.

Electrical Control

The heater film element has to be protected from overheat as this could cause windscreen delamination.

Figure 13 shows the heater element controlled by a temperature control unit.

Within the unit is a bridge circuit with one element of the bridge connected to the windscreen film heater. When the switch in the flight deck is switched on it operates a relay within the control unit. This causes the heater element to start to heat up. Initially the bridge circuit within the temperature control unit is unbalanced, but as the screen temperature begins to increase so the sensing element gets hotter and it's resistance increases. At a pre-determined temperature (say 40°C) it's resistance is high enough to balance the bridge - this de-energises the relay and switches off the supply.

CONDUCTIVE \ COATING (ANTI ICE)

CONDUCTIVE / COATING

(DEFOG)

I Fig. 12 CROSS-SECTION-OF. DClO/MDl l SOLID WINDSCREEN

I I REI!AY GANGED SWITCH I

I-- ------- 115V ac SUPPLY

- - - --

28V dc SUPPLY

- TEMPERATURE CONTROL UNIT

SENSING ELEMENT

WINDSCREEN

Fig. 13 WINDSCREEN ANTI-ICE TEMPERATURE CONTROL

moodull l A-1 172

With the power supply interrupted the heater element and temperature sensing elements start to cool. When cool enough the bridge will become unbalanced enough to cause the relay to come on. When it does so it will start to heat up and start to unbalance the bridge again.

The system continues in this mode until switched off. A separate overheat standby sensing element is also fitted (but not shown).

Modern windscreens, such as the A320, are computer controlled (see figure 14).

Hot Air Blowing

To prevent icing and misting, hot air is sometimes blown over the outside (anti- ice) and inside (anti-mist) of the windscreen and other transparent panels. The principle of the outside air blast system is also used for rain clearance on some aircraft such as the DC8 and the King Air. -

-

Great care must be taken with ground running aircraft engines with this system to ensure the system is switched off to prevent overheating the windscreen.

The hdt air is usually taken from a tapping on the pneumaticducting upstream of the conditioning packs. This ducting is usually from the compressor stage of the jet engine leading to the packs.

The A32OPWindscreen Heating-System (Figure 14) - -

Electrical heating is provided for anti icing of each windshield and demisting of side windows of the cockpit.

Two independent window heat computers (WHCs), one on each side, automatically regulate the system and provide overheat protection and fault indications.

The window heater comes on automatically when the engines are started or, switched on manually prior to engine start or if the aircraft is in the air (signals via the OR gates). The heating operates at low power on the ground and at higher power in flight. The change-over is automatic via a weight-switch signal to the computer.

System status data is sent via ECAM (Electronic Central Aircraft Monitor) and symbol generators to flight deck displays.

- 2.5 -

rnoodull l A-1 173

ECAM 4-:

:w_~ _C - A

ENG I OR 2 RUNNING -

FLIGHT

L!??F- ., . . -. J

w PROBUWINDOW RH SIDE

IDENTICAL

Fig. 14 320 COMPUTER CONTROLLED WINDSCREEN ELECTRICAL I ANTI ICING SYSTEM

-

I

I WINDSCREEN WASH AND RAIN REPELLENT SYSTEMS

Most ;windscreens are kept clear by windscreen wipers. These are not too dissimilar to those used on automobiles and are operated hydraulically or electr-ically. Used to maintain clear-vision through the windscreen when raining, sleeting or snowing and will clear snow and ice in conjunction with the de-icing system.

They work well on the ground at normal taxiing speeds and take-off and landing but in-flight in heavy rain their effectiveness is limited. In some cases to make rain clearance easier a rain repellent solution can be sprayed onto the screen. Figure 15 shows a system as fitted to the B747 and figure 16 shows the flight deck control panel as fitted to the B777.

In figure 15 each front windscreen is fitted with a two speed electric wiper. The wiper is controlled by a rotary switch. Emergency park selection is used if the wiper fails to park when switched OFF. In moderate to heavy rain, a rain repellent may be sprayed onto the windscreen to improve visibility. This is spread by the wiper to form an even film which allows the water droplets to be blown away by the airflow. Pushing the same switch down will operate the wash bottle - which helps clean the windscreen of debris (insects, dirt etc).

rnoodull l A-1 174

I WASH I

@YQ SWEEP

WHEN EMERG EMERG WIPER PARK OFF FAILS SHOULD SLOW TO ONLY PARK BE AT OFF USED

PARK

U FOR 0.6 SEC ON I

NOZZLE i

OPE1 EACH APPLICATION

1 WINDSHIELD WASH BOTTLE

Fig. 15 TYPICAL WINDSCREEN WASH AND RAIN REPELLENT SYSTEM

I

--

WINDOW HEAT - WINDOW HEAT SWITCH

L- R- (ALTERNATE ACTION)

ON -WINDOW HEA; IS APPLIED TO THE SELECTED WINDOWS

INOP INOP INOP ILLUMINATED AMBER - * THE SWITCH IS OFF * AN OVERHEAT IS DEECTED, OR

ow: FRG WIPERS

-.-..-. . -..-

I I * A SYSTEM FAULT HAS OCCURED

OVERHEAD PANEL

m u 1 ILC

Fig. 16 WINDSCREEN HEAT CONTROL PANEL - B777

h -

Rain repellent is removed from the windscreen on the ground by the use of rain repellent solution as listed in the AMM.

On some aircraft such as the Boeing 777 the front windscreens are treated with a hydrophobic coating (hydrophobic meaning water-repellent). This means that there is no need for a liquid rain repellent system as the water droplets disperse by the airflow without additional help.

WINDSCREEN WIPERS

The wiper action is similar to that of an automobile in that a wiper blade is caused to move back and forth over the windscreen - thus clearing water droplets, snow, sleet etc to provide clear vision. The wiper blade is made of rubber with a knife-edge in contact with the windscreen. It is supported in metal support channels that ensures even blade pressure contact with the windscreen.

Figures 17 and 18 shows the electrically operated blade system for the B757. It is typical in that it is electrically powered and sweeps the blade over a part of the forward facing windscreen.

Operating parameters for the blade can include:

* Parking the blade off the windscreen when switched off. ~r The number of blade sweeps per minute (may be variable). * __ The angle of sweep. * The contact pressure of the blade on the screen - to ensure good

i I rain clearance and prevent displacement of the blade away from the windscreen by the airflow.

* Speed. The speed a t which the blade completes each sweep. * Smooth operation.

The maintenance of the blade system involves:

x Checking power supplies to motor and flight deck controls. JC Checking for general serviceability of all components including the - -

condition of the rubber blade. * Changing any defective components with the most usual change

being the rubber blade. * Operational checks on the system (see below).

Operational checks are specified in the AMM and are carried out when specified in the maintenance schedule or after a component change or when a fault is reported. Remember when operating the wiper that the windscreen must be kept wet a t all times - otherwise damage to the screen may result.

Before commencing tests ensure windscreen is clean and the blade is in contact with the screen and not any part of the blade supported structure.

Ensure the sweep area is clean and smooth from the park position to the screen.

Ensure that windscreen electrical anti-ice supply is disconnected. On some aircraft there is a possibility of an electric shock from the screen if this is not done.

WIPER 5'

..

-- WIPER BLADE METAL SUPPORT SYSTEM

Y// ' WIPER ARM

WIPER BLADE (MIDSTROKE POSITION)

LIFT 'BLOCK WIPER BLADE - (PARK POSITION)

W I P ~ R ARM HUB / (SEE DETAIL A)

ADJUSTMENT SQEEVE PRESSURE

(FITTED TO MOTOR ADJUSTING BOLT

I DRIVE SHAFT) , ;--

\ I

-

] DETAIL A WIPER ARM HUB --

Fig. 17 WINDSCREEN WIPER SYSTEM

The checks that can be carried out include:

Controls. Check that the controls on the flight deck operate the left and right wipers correctly and as directed. Ensure all flight deck indications are correct.

Sweep. Check that sweep length is correct (as per AMM) and that the blade does not go over any windscreen sealsledges except when selected to park.

Speed. Check the 'sweeps per minute rate' is within the limits laid down in the AMM and correspond to the setting control in the flight deck (LOW, MEDIUM, HIGH). Rates can vary u p to about 300 sweeps per minute. For high sweep speeds it i s recommended to use a commercially available Sweep Counter.

rnoodull l A-1 177

WINDSCREEN WINDSCREEN

w

Fig. 18 WIPER ARM DETAIL

check for smooth operation on each sweep and that the blade clean .s the winddcreen correctly.

Blade Cbntact Pressure. usingla spring balance (scale 0-20 lbs) or similar device attach at the wiper armposition where the blade assembly attaches to arm and pull the blade away from the windscreen. A minimum force should be required to lift it off (eg 15 lbs).

Park. When selected to PARK ensure blade parks correctly.

Rectification

Should it fail any of the above tests then rectification can include, as appropriate :

* Checking electrical supplies. * Changing control panel in flight deck. -k If sweep or park position is not correct this can be adjusted for at

the adjustment sleeve on the wiper arm pivot bolt.

- 30 -

rnoodull lA-1178

* If speed cannot be maintained within limits (sweeps per minute) then the motor will have to be changed.

* If blade contact pressure is incorrect this can be adjusted for by adjusting the pressure adjusting bolt.

* If the blade does not give a clear smear free wipe of the screen then change the blade.

Remember, when working on the blade or the system to isolate the electrical supply and tag. After any adjustments/replacements, retest, re-configure the aircraft and clear all paperwork.

FLUID AIRFRAME DE-ICING SYSTEMS

In systems of this type, a de-icing fluid is drawn from a storage tank by an electrically driven pump and fed through filters to a number of porous metal distribution panels. The panels are formed to the profiles of the leading edge of the s t ~ c t g r e onto which they-refitted. At each panel the fluid passes into a cavity, and then through a porous plastic sheet to a porous metal outer skin. A s the fluid escapes it breaks the bond between the ice and the skin of the aircraft, and the fluid and ice together are blown away by tkie airflow.

-- -

- - -

I

I 1 PRESSURE GAUGE , LOW PRESSURE

WARNING SWITCH

TANK TO DISTRIBUTION PANELS

--- -- --

F '*-*-I PROPORTIONING UNIT

FILTER

TO PROPELLERS + + - Hi

FLUID CONTENTS POWER SUPPLY

Fig. 19 FLUID SYSTEM

The porous metal element may be made of sintered stainless steel or laser drilled titanium (expensive).

(Sintering is a process whereby a fine metal powder is pressed into shape with heat applied to produce a component. If the pressure is not too great then the component is porous - filters and de-icing fluid distribution panels. Laser drilling produces hundreds of minute holes per square inch which under normal conditions are only just visible).

rnoodull lA-1179

POROUS ISINTEREL. -.. . . .-...-...

METERING TUBE

POROUS STAINLESS STEEL (SINTERED) OR TITANIUM (LASER DRILLED)

--SUPPLY

Fig. 20 DISTRIBUTOR PANEL

QUESTION The distribution panel is fitted on the leading edge of the wing in such a way as to cover the full range of the stagnation point. What is the stagnation point? This is theoG-of-flight revision. (2 mins)

ANSWER The stagnation point is where (in theory) a small molecule of air I stays in front of the aerofoil throughout the flight. The actual point

on the leading edge will change with the change in angle of attack - moving down relative to the leading edge with a high angle of

I

attack and up with a reduced angle of attack.

The system-shown in figure 19 is a simplified system but is typical of those found on aircraft. The fluid is drawn from the tank and pumped to the various systems via a filter. The pump is operated either via a switch in the flight deck or automatically via the ice detection system.

There may be a tank contents indication system fitted, either a capacitive type or a float type. Other system indications may include a low level warning, a pressure transducer and a low-pressure warning switch. Before the fluid reaches the distribution panels it goes via a proportioning unit. This meters the quantity of fluid to each panel depending on the panel's size. On a propeller driven aircraft de-icing fluid may also be fed to the propellers via slinger rings.

It is normal not to have the same system supply fluid to the windscreen as the fluid can adversely affect the structure of the windscreen.

rnoodull l A-1 180

Installation and Maintenance

Distributor panels. Before installation of a distributor panel the part number should be checked to ensure that it is the correct type (IPCIAMM).

Where aerofoils are not symmetrical odd part numbers usually indicate a left- hand (port) component and even numbers right-hand (starboard) component. A check should also be made for signs of corrosion, damage and deformation of the panel profile.

Panels which are to be sealed by an epoxy resin compound should also be checked to ensure that the PTFE coating release agent is undamaged. Check panel for security of attachment.

Before fitting check the aircraft structure to which the panel is to be fitted for corrosion and damage - repairlreplace as per SRM if found damaged.

Tanks. Before installation, the tank should be inspected for signs of damage and interior cleanliness, and the tZnk supporting structurd checked to ensure that it is in a suitable condition to receive the tank.

I

With the tank accurately positioned it should be firmly secured by the mounting straps or mounting bolts as appropriate to the type of tank. Blanks should be removed from all pipe unions, new sealing rings fitted and the pipelines connected. I

I ' I

Pumps. Before installing a pump kt should be inspected for signs of damage, and checks made that the part number of the pump is correct (EASA form 1). Connect electrical power and carry out a functional check.

Note. The inhibiting fluid should be retained in the pump for the functional check, the duration of which should not exceed that specified. Do not dry run the pump.

Pipelines and Couplings. Pipelines should be supported by clips spaced not more than 18in apart (457mm). Where there is more than one pipe in a run the pipes should be cleated together every 6in (1 52mm).

Filters. After installation, the system should be operated until fluid flows from the filter outlet connection or, if provided, from a bleed screw hole. The pump should then be switched off and the filter outlet connection or bleed screw refitted.

proportion in^?; Units. Before installing a check should be made that they are of the correct type. Check part number, description, IPC and EASA form 1.

rnoodull lA-1181

Priming and Testing

After the installation of a component and at periods specified in the Maintenance Schedules, priming and functional testing of the system should be carried out as specified in the AMM.

In general, check that the tank is full and that, when the pump is operated, fluid flows out from each of the distributor panels. They will appear to sweat and the drops will run down the underside of the wing to fall on the floor. Pipes to propeller slinger rings will run fluid freely. Check for correct fluid distribution and leaks and wire lock all unions.

PROPELLERS

Propellers may be anti-iced by using:

* /- Fluid * Electrical power via contact brushes at the back of the spinner.

x Induced ac electrical power via coils behind the spinner and coils on the back of the spinner itself.

I I

The fluid system provides a film pf de-icing fluid to the propeller blade surfaces during flight which mixes with1 the water or ice and reduces the freezing point of the mixture. Where ice has already formed on the blades, the fluid penetrates under the ice and lbosens it sufficiently for it to be thrown off by centrifugal action.

- --

Fluid is distributed to each propeller blade from a ~linger ring which is mounted on the back of the propeller hub or spinner. The fluid is pumped into this ring through a (stationary) delivery pipe.

Some propellers have rubber overshoes fitted to the leading edges of the blade, to assist in the distribution of the fluid. On this type of installation fluid is fed from the slinger ring to a small cup, which is part of the overshoe, and is then forced by centrifugal force along spanwise leading edge grooves (about 1 to 2mm deep and 5mm wide) in the overshoes. The groves finish about 1 /3'd to 1 / 2 of the blade length.

The action of the centrifugal force and the airflow will distribute the fluid to both sides of each blade.

On propellers which are not fitted with overshoes, fluid is fed from the slinger ring through a pipe to the root of the blade, as before, and is then distributed by centrifugal force and the airflow over the blade.

rnoodull lA-1182

The fluid may be pumped to the slinger ring from the supply tank by an electrically driven pump but pneumatic pressure is sometimes used. The electric pump is controlled by a switch and, in some installations, the pump speed may be varied by means of a rheostat. Non-return valves are sometimes provided to prevent loss of fluid when the pump is not operating.

The Slinger Ring

The slinger ring rotates with the propeller. The pipe and nozzle which deliver fluid to the slinger ring should be positioned so that there is clearance between the pipe and the side of the ring to prevent interference when the propeller is rotating. This clearance is important as the tolerance is small and an error will cause excessive wear or the system will spill fluid.

- PROPELLER FLUID DISTRIBUTION SHOE

SUPPLY PIPE I

SLINGER RING (ROTATING)

PROPELLER BLADE I

- Fig. 21 PROPELLER FLUID DE-ICING

VP Propeller

For VP (variable pitch) propellers, the propeller feed pipe leading from the slinger ring to the front of each blade is positioned to end just inside an oval rubber fluid cup bonded to the front of the blade. There is a small clearance between the feed pipe and the cup and the arrangement allows the blade to change pitch and still receive de-icing fluid.

Function Test

If there is any doubt as to whether the propeller de-icing system is functioning properly i t should be checked during an engine ground run. The propellers should be painted with commercial whitewash (or methylated spirits and whiting) and allowed to dry.

rnoodull l A-1 183

A suitable dye should be added'to the fluid so that when the de-icing system is operated the dyed fluid will stain the white propeller and indicate the distribution pattern over the blades.

Ideally, carry out the test when the engine has to be run for other reasons, and towards the end of the engine run switch the anti-icing system on. The pump will pump the fluid to the slinger ring where centrifugal force takes over and the fluid is forced out through the hole in the slinger ring (one at each blade), through a short pipe to the beginning of the fluid distribution channels on the front of each blade.

Centrifugal force and the airflow help to distribute the fluid evenly over both the front and rear surfaces of the blade.

Before stopping the engine switch off the anti-icing system. Shut down the engine and inspect the blades.

~ a c ~ blade should show an even distribution over both surfaces. Uneven distribution may be caused by the slinger ring being fitted incorrectly, by the feed pipes from the ring being incorrectly located or by obstructions in the pipelines.

The blades should be cleaned to remove'all traces of fluid and whitewash.

Often the system is just switched on during a ground run and the blades checked by feel after engine shut-down. This saves a lot of time painting and cleaning.

Cleaning the Fluid System

When the system is to be out of use for a long period it is advisable to remove all traces of de-icing fluid. This may be done by draining the tank and re-fillinb with a mixture of 95% methylated spirits and 5% distilled water.

The system should be operated until the tank is empty. During this operation the engines should be run if the system involves propeller fluid de-icing.

Inhibiting the Fluid System

The fluid used in de-icing systems is stable and non-corrosive but leaves a gummy deposit after drying out. Inhibiting the system is at the discretion of the aircraft operator, but if it is not inhibited it is advisable that a certain level of de-icing fluid is maintained in the tank and the system operated at regular intervals.

rnoodull l A-1 184

To inhibit the system after draining and/or cleaning, fill the tank with inhibiting fluid as per the AMM or use fluid DTD 5540. Run the system to ensure all internal surfaces are coated. Inhibiting is carried out for aircraft storage and if the aircraft will not be flying in icing conditions for some time.

To de-inhibit the system flush through with a cleaning fluid as per the AMM. Refill with the de-icing fluid and test.

Electrical De-icing of Propellers

This is done using dc or ac supplies with heater mats bonded to the leading edges of each propeller, and sometimes to the spinner. Each propeller heater mat starts a t the root and has a length of about l/z the blade length. On large aircraft the supply is 200V 30 ac, on smaller aircraft the supply is dc.

Heater elements may also be bonded to the inside of the propeller spinner with \

connecting wires from the spinner to the junction box fitted-to the front of the propeller bulkhead - which r ~ t a t e ~ w i t h the propeller.

Wires from each of the blade heater mats also lead to the junction box. The junction box is connected to slip rings on the rear of the propeller bulkhead. The slip rings are in contact with stationary brushes fitted in a brush housing on the engine support structure. The electrical supply goes ithrough the positive brush gear to the positive (rotating) ring gear through the heater mats and back via the negative rotating ring gear and negative brush gear.

I

On some modern aircraft electrical power is transferred to the rotating propeller via induction coils on the stationary airframe maprotating coils inside the rotating propeller hub - with no physical contact.

For more information on propeller de-icing refer to the book on propellers in this series.

PNEUMATIC AIRFRAME DE-ICING SYSTEMS (DUNLOP OVERSHOES)

This is a mechanical de-icing system and would not be used for anti-icing.

Air pressure and vacuum is supplied alternatively to inflatable rubber overshoes fitted to the leading edges of mainplanes, tailplane and fin and when alternately inflated and deflated cause the ice to break away.

Pressure and vacuum is supplied to the distribution valves which cycle the supplies alternately to the supply pipes to the de-icer boots.

rnoodull l A-1 185

It is inexpensive to fit but disturbs the airflow (at a critical point on the wing - the stagnation point) when in operation - reducing lift and increasing drag.

The Overshoes

The overshoes or de-icer boots are made of a rubber/rubberised fabric or a polymer material and are manufactured in such a way as to produce inflatable tubes that run spanwise on some aircraft and chordwise on others. The span wise boots may run for nearly the whole length of the wing. The chordwise boots start at just under the leading edge and run over the leading edge to finish a few inches after the leading edge on the top of the wing.

The tubes within the boots are connected to pipelines that supply alternating pressure and vacuum. The external surface of the boots are covered by a conducting film to allow the discharge of static electricity.

- ----

T h e ~ are, fitted to the leading-edges and can run for the full wing span length. The width of the boot i s related to thestagnation point range of the aerofoil.

TO PORT WlNG BOOTS

ELECTRONIC CONTROL

ION VALVES

TO STARBOARD WlNG BOOTS - Fig. 22 PNEUMATIC DE-ICER SYSTEM

Air Supplies

The tubes are inflated by air from the pressure side of an engine driven vacuum pump or from a high pressure reservoir or in the case of some types of turbo-propeller aircraft, from a tapping from an engine compressor stage.

This provides the inflation stage of the operating sequence.

The deflation cycle of the sequence (and whenever the system is switched off) is provided by vacuum air from a vacuum pump, or, in some systems an engine compressor tapping using the venturi section of an ejector nozzle.

Controls and Indicators

The system can be switched on by the pilot but may be switched on automatically from the ice detection system.

The timing of the cyclic distribution valves is carried out by an electronic control, and flight deck indications include indicating lights and electrically operated pressure and vacuum gauges. For a specific installation reference should be made to the appropriate AMM.

Operation

When the system is switched on, pressure is admitted to the boot sections to inflate each alternate tube. The inflation weakens the bond, between the ice and the boot surface, causing the ice topbreak away. At the end of the inflation stage the tubes are fully deflated 'by the vacuum supply.

I , 1

The inflation aqd deflation cycle is repeated continuously while the system is on. When the system is switched off vacuum is supplied continuously to all tubes to hold them flat against the aerofoil leading edges thus minimising drag.

- - p-p

Installation

Full details of the checks to be carried out on the components prior to installation, and installation methods, are given in the AMM. De-icer boots may be attached by rivets or special fasteners or may be bonded into position.

Inspection and Maintenance

The majority of inspection and maintenance associated with these systems is related to the de-icer boots, since their location on an aircraft makes them vulnerable to damage from airborne debris (ice, rain etc) and ground handling (step ladders, refuelling hoses etc) .

Precautions when considering the boots include:

(1) Re-fuelling hoses and other equipment must not be dragged over the surfaces.

(2) Ladders or service platforms which are placed near the boots during servicing operations must have soft padding fitted to prevent damage.

(3) Oil or grease on the surface of boots must be removed as soon as possible with soap and water or with a clean rag moistened with lead-free petrol. Petrol should not be allowed to dry on the surface; it should be wiped off immediately with a clean dry cloth.

Surface Deterioration

The conductive surface of cemented de-icer boots deteriorates slowly in service through - general abrasion and this will show as a roughened surface.

-

-

Repairs I I

For specific repair procedures refer to the AMM/SRM or the boot manufacturer's repair manual. In general small holes and minor damage may be repaired using a repair kit sirhilar to a bicycle puncture repair kit (follow the repair kit instructions). These cold patch repair kit covers cuts, holes and cracks up to 2"' (51mm) long. For larger damage emergency repairs only may be carridd out. The best way to repair extensive damage to a boot is by replacement. - - - - - - - Removed boots may be sent back to the manufacturer for repair by vulcanising or it may be scrapped.

Functional Tests

These must be carried out at the periods specified in the maintenance schedules; when a malfunction occurs; or a component has been replaced, and also after repairs. The method of testing depends on the type of aircraft and details must therefore be obtained from the AMM, but in general:

(1) Tests may be carried out using either the aircraft engines or air supplies from a ground test trolley. If a system is to be tested by using a test trolley, the air supply must be clean, moisture-free and at the correct pressure.

(2) Pressure and vacuum indicators should be checked to ensure that supplies are maintained at the specified values.

rnoodull l A-1 188

(3) With a system selected 'ON' boots should be checked to ensure that they inflate and deflate in the correct sequence and for the correct periods of time.

(4) When checking pipelines etc for leaks it is important to remember that any leaks in the vacuum line will suck in any detection fluid used (Snoop etc). So do not use detection fluid on the vacuum lines to the distribution valves or the lines from the distribution valves which cycle from pressure to vacuum.

Storage of De-Icer Boots

Before storing, boots should be inspected for condition and defects. They should be cleaned and repaired if necessary and dusted with French chalk.

Connectors should be blanked off and the boots rolled up. Rolling should be ' ,

comm6iced at the end remote-from the valve which should be on the outside of the finished roll. Where connectorsZe located near the centr63f the boot, a pad of corrugated paper should be placed oper the connectors to protect the contacting surface.

, The rolled boot should be carefullywrapped up in a heavy paper to exclude all light, a& then stored in a cool,: dry, d&k place, away from ky electrical running equipment where it will not be crushed or wrinkled. In cases where boots are bonded to detached leading edge sections, the seqtions should be wrapped up and supported on their trailing edges.

- -

- - --- ---- -

QUESTION When storing anything made of rubber it must be away from electrical running equipment. Why? (2 mins)

ANSWER Electric motors and particularly generators generate ozone (0 3)

when running - this is the principal ageing element of rubber.

HOT AIR DE-ICING/ANTI-ICING SYSTEMS

In general the leading edge sections of wings, tail units, intakes and slats are usually provided with a second inner skin to form a small gap between it and the inside of the leading edge section. Heated air is ducted to these sections and passes into the gap, providing sufficient heat in the outer skin of the leading edge to melt ice already formed and prevent further ice formation.

The air is exhausted to atmosphere through outlets in the skin surfaces and also, in some cases, a t the tips of wings and tail units.

rnoodull l A-1 189

The temperature of the air within the ducting and leading edge sections is controlled by a valve system which depends on the type of aircraft.

Hot air can be obtained from:

(a) A tapping from the compressor stage of a jet engine (common for large jet transports).

(b) A combustion heater - using a heat exchanger and ram air (used on some piston engined aircraft).

(c) A heat exchanger system using the heat from the exhaust gases. Not common but used on some aircraft. The air being used for cabin heating, anti-icing etc.

QUESTION What sort of engines would use (c) above? (5 mins)

ANSWER Small piston engined aircraft might use a heater muff (heat exchanger) around the eihaust pipe. Ram air is passed through

I which picks u p heat from the exhaust pipe and is used for cabin I heating,anti-icingletc.

I Some turbo-prop engines use the jet efflux gases via a heat exchanger for anti/delicing purposes. The gases being passed back

I I I into the jet pipe. -

QUESTION Why might the effl'p gases be used in a turbo prop engine and never on a pure jet engine? (10 mins)

L- ---

ANSWER Approximately 90% of a turbo-prop engine's power is used to drive the propeller with the residual thrust from the engine accounting for about 10%. So interfering with the jet efflux will have little effect on the overall engine performance - though it will have some of course. 'Mucking about' with the efflux of a pure jet would seriously affect it's performance as all it's thrust is obtained from the exhaust. The speed of the efflux would also be high and airflow control would be difficult.

In many cases the tapping from the compressor side of the jet engine is integrated with the cabin air supply system (refer to figure 23) with the anti- icing air being tapped off to the anti-icing system before the (hot) cabin air enters the cold air units/conditioning packs to be cooled.

The tapping will be from about the 7th stage, and on some engines there are two per engine (say 7th and 14th with the 14th modulated a t high rpm) - but check the AMM.

With combustion heaters and exhaust heat exchanger systems there is a risk of carbon monoxide poisoning of passengers and crew should any part of the ducting down-stream of the combustion area leak. Carbon monoxide detectors are fitted where there is a risk of such poisoning.

Duc ting

In general the ducting may be made of stainless steel, light alloy, or composite material. The material used must be of the correct specification to resist the air temperatures encountered at that particular part of the ducting where normally metal like stainless steel is used early on in the system where the temperatures are highest.

Temperature Control

The conEol of the air temperature within ducting and leading edge sections is important as overheating and burnsg of the metal and ducting may occur.

In a typical compressor bleed system, control is effected by temperature sensing units which are located , at _ various _ points in the ducting and by valves in the air supply ducting. The sensing units and valves are electrically interconnected and the valves &re automatically positioned to,-regulate the flow of hot air. \

When heat exchangers are employed, temperature control may be obtained by the use of adjustable flaps and valves-to decrease or increase the supply of

--

heating &d/%cooling air across the exchangers. The mefhod of controlling the flaps and valves varies, but a typical system incorporates an electric actuator, which operates automatically by a device controlled by a temperature sensing element fitted in the duct on the warm outlet side of the heat exchanger.

In systems incorporating combustion heaters, the temperature may be controlled by thermal cyclic switches located in the heater outlet ducts, so that when the temperature reaches maximum the fuel supply to the heaters is automatically switched off. When the temperature drops to a pre-determined value the heater is started back up.

Installation

Refer to the AMM. Always check for security of attachments, corrosion and damage of components and in general the following applies:

rnoodull l A-1 19'1

Heat Exchangers. Before installation, heat exchangers should be inspected to ensure that no foreign matter has entered the various connections, and they are free from obstruction.

Combustion Heaters. Before installation, combustion heaters should be inspected and when necessary, pressure tested in accordance with the manufacturer's maintenance manual to ensure that no fuel or combustion products can leak into the air supply of the system (see GR 11 of CAP747). Pressure tests are usually to 2psi (or as stated in the manual) with the unit submersed in water, special bungs fitted and in a suitable safety cage.

There should be no leaks.

Combustion heaters and their hot air outlet ducting should be dismantled, inspected, reassembled and pressure tested at least once every 500 heater operating hours. These hours are assumed to be the same as aircraft flying hours unless agreed by the CAA to be a percentage of the flying hours. \

System inspection and Maintenance

~ h e c g the AMM and the maintenance schedule. The following general points apply./ Check all components for:

(a) Damage, cracks and evidence of leakage. ,(b) Security of attachment, loclang and corrosion. (d) Security of fittings, ducts, electrical connections, controls. (e) Damage and security of cables and moisture ingress.

Testing

Carried out in accordance with the AMM but in general:

Carry out any BIT checks - check on-board fault computer. Ensure all intakes/exhausts are free from debris, birds nests etc. If ram air is required - as for combustion heaters - use a ram air supply servicing trolley. Ensure electrical power is on. Run engines if engine hot air is used. Select temperature required and monitor temperature outlet/ s. Check flight deck indications and warning captions. Check temperature of airframe/cowling surface by hand or use surface temperature thermometer. Check for leaks in the ducting.

rnoodull l A-1 192

CONDITIONING PACKS

DUCTS AFT OF SPAR

ANTI-ICING (OUTBOARD)

- -

SEPARATE FLIGHT DE-ICING ENGINE PRECOOLER DECK SUPPLY (INBOARD)-- BLEED

Fig. 23 BAe 146 PNEUMATIC SUPPLY SYSTEM

SPLITTIER LIP AND 1st STAGE INLET VANES ANTI-ICING /

SPINNER ANTI-ICING \ /

INTAKE ANTI-ICING

\ 7th STAGE BLEED

Fig. 24 BAe 146 ENGINE ANTI-ICING DETAIL

Example System - BAe 146

Figure 23 shows the BAe 146 anti-icing hot air system. The air supply comes from a tapping on each engine and also from the APU. The air is fed to the mainplanes and tailplane (stabiliser). Anti-icing air comes from a separate (7th

stage) tapping on each engine for the intake, splitter lip (between bypass air and core air) and LP compressor inlet guide vanes.

rnoodull l A-1 193

ENGINE ANTI-ICE SELECTOR

I I I I (THREE POSITION

ROTARY) ANTI-ICE

OFF - THE ENGINE

L ENGINE / I ANTI-ICE VALVE IS

I AUTO COMMANDED CLOSED

AUTO

OFF@ON

APPROACH IDLE IS SELECTED BY THE EEC AND CONTINUOUS

I

L WING ANTI-ICE SELECTOR (THREE POSITION IGNITION IS

ROTARY) AUTOMATICALLY Z U ~ D L ~ C ~ nu

OFF@ N OFF@ N

I

OFF - BOTH WlNG ANTI-ICE VALVES ARE COMMANDED CLOSED

AUTO ENGINE - ,N ANTI-ICE FLIGHT, VALVE THE IS OPENED OR CLOSED AUTOMATICALLY BY THE ICE DETECTION SYSTEM

ON -

AUTO - IN FLIGHT, BOTH WlNG ANTI-ICE VALVES ARE COMMANDED OPENED OR CLOSED AUTOMATICALLY BY THE ICE DETECTION SYSTEM

* THE ENGINE ANTI-ICE

OVERHEAD PANEL VALVE IS COMMANDED OPEN

ON - IN FLIGHT, BOTH WlNG ANTI-ICE VALVES ARE COMMANDED OPEN

Fig. 25 ANTI ICE CONTROL PANEL OF THE B777

' I

The engine spinner is anti-iced by hot engine oil - which also acts a s an oil heat sink (oil cooler).

Control is via the overhead panel on the flight deck and sections of the system may be in the de-icing or anti-icing mode.

Windscreens,/ - Pitot heads and stall warning vanes are electrically anti-iced. -

Figure 25 shows the flight deck anti ice control panel of the B777. Study the drawing and note the available selections. Note also that the tailplane and fin are not anti-iced. (EEC = Electronic Engine Control.)

A320 Engine Anti-ice

Figure 26 shows the intake anti-icing system for the Airbus A320.

Intake air is taken from a separate tapping from the high pressure compressor of the jet engine. The air is controlled using an on/off valve which is switched from a push button (one for each engine) on the control panel on the flight deck.

The control valve closes when the engine is shut down and automatically opens in the event of electrical power failure (fail safe).

moodull l A-1 194

ANTl ICED AREA

OVERHEADPANEL

ANTl ICE

Fig. 26 A 3 2 0 ENGINE INTAKE ANTI-ICING SYSTEM

- --

When open the N 1 limit of the respective engine is reduced. I

I

Note the 'signals going to FADE$ (Full Authority Digital Engine Control - a computer involved in the control of the engine) and ECAM (Electronic Centralised - Aircraft Monitor - for flight deck CRT screen didplays).

OTHER ANTI-ICED AREAS

These include:

A Air data probes particularly Pitot heads and TAT (Total Air Temperature) probes. Some-times static vents.

* External sensing units such as angle of attack vanes. * Drain masts, in particular external grey water drains. * Door mechanisms - those that suffer from condensation freezing. * Potable water servicing valves and internal domestic water supplies

systems where there is a possibility that they might freeze.

The above are anti-iced electrically with single phase 115V ac, although dc is used on some components. The heater elements may be controlled by a computer with overheat protection being provided by an inbuilt thermostat. If anti-iced slats and flying control surfaces usually use hot air systems.

rnoodull l A-1 195

Slats

Figure 27 shows an example of how the hot air is distributed to the leading edge of the slats. As the slat is moved forward so the telescopic feeder duct extends. This allows air to the leading edge of the slat a t any position. The air is allowed to the inside surface of the slat via drillings in the slat duct.

Fig. 27 HOT AIR SUPPLY TO SLATS

- 48 -

rnoodull l A-1 196

Figure 28 shows an example of the electrical power supply. You should note from this the general layout of the location of the generators, external supplies, bus bars and the supplies taken off the single phase supplies, and the dc supply

Some systems have automatically reduced power when the aircraft is on the ground and heaters such as the TAT heater are turned off completely (automatically). The probe heater control unit takes care of these functions with interfaces to the landing gear weight switch.

Some control units also have interfaces with the engine vibration monitoring system and the Central Maintenance Computer (CMC). On modern systems they will identify and memorise faults.

If current gets too high or too low a warning is signalled to the flight deck warning computer.

GCU-2

AC BUS 2

GLC-2 4000XU

BTC-2

RIGHT WINDSHIELD HEATING

LEFT WIN,WW HEATING

I I L 1 2 D A 2 PROBE HEAT,

I FIO TAT

' 3PU2

EXT PWR-A

EXT PWR-B EPC-B

12XU SIC

1-k K A PROBE HEAT. CAPT PITOT, PROBE HEAT. STBY PlTOT

IDA3 HEAT' STBY ALPHA

C

AC BUS I l X P l

G E 1 &y , I D G I WINDSHIELD LEFT HEATING

GCU-1 12DAl PROBE HEAT, CAPT TAT 4000XU lXU1

2DA1 PROBE HEAT, CAPT ALPHA 4DA1 PROBE HEAT. STBY PlTOT 4DG1 LEFT WINDOW HEATING

Fig. 28 EXAMPLE OF ELECTRICAL POWER SUPPLIES

RAIN MAST (HEATED)

VER TEMPERATURE SWITCH

ELEMENT TURE SENSOR

WASTE WATER OUTLET

Fig. 29 GREY WATER DRAIN MAST

Heaters such as internal domestic pipework heaters (ribbon heaters), door mechb i sm heaters, and drain mast heaters only come on when icing conditidns are sensed.

I

Testing

Always refer to the AMM. On older aircraft testing is done carefully with the fingers by switching the heaters on and noting the temperature (an accurate reading can be obtained by the use of temperature indicator tongs placed around probe type heaters). On modern aircraft inbuilt testing (BIT) is carried out every time system power is switched on, and on command from the central maintenance computer. The tests involve checking computer equipment such as the CPU, RAM etc, also inputs and outputs and the integrity of heater probes.

Figure 29 shows the heater details of a grey water drain mast (called grey water to denote that it comes from the hand basins and sinks of the aircraft (moderately clean) - with blue water indicating that it comes from the toilets which is not allowed overboard of course). Note the heater element, temperature sensor and over temperature switch.

Figure 30 shows the general layout of a probe heater computer. Note the heaters of the various probes. Note also the 115V ac single phase supplies via a dc operated relay and the 28V dc supply. Note the logic inputs from the engines and landing gear and the data connections to the CMC.

moodull l A-I 198

L STAT PROBE (7DAl)

1OlPP STATIC POWER SUPPLY INPUT

BUS 1

11- CASE GROUND HEATING POWER SUPPLY

AOA POWER SUPPLY INPUT -11

901XPC PITOT POWER

34-1 1 4 0 101XPA TAT POWER

SUPPLY INPUT

ENG 7. I ENG 2 RUNNING

I OR FLIGHT INPUT

ENG 1 RUNNING OR FLIGHT INPUT

ElVMU 73-25-00 HEATING A V A I L A B I L I T Y 31-54-00

- - OUTPUT

PIB SW 'ON POSITION' STAT FAULT OUTPUT INPUT I 6DG I POWER SUPPLY AOA FAULT OUTPUT

BUS 3 TATNOHEATOUTPUT RESET,

PlTOT FAULT OUTPUT

LGCIU 2 'FLIGHT' INPUT / VALID LGCIU 2YALIDAT~ON' INPUT CHANNEL A OUTPUT

5GA2 LGCIU-2

LGClU 1 'FLIGHT' INPUT LGCIU 1 VALIDATION' INPUT

5GAl LGCIU-1 451240

ADIRU 34-11-00

Fig. 30 PROBE HEATER COMPUTER

CONTENTS

Page

Landing gear general Nose undercarriages Steering Sliimmy Power steering Undercarriage systems Locking methods -

Indication systems Undercarriage selection Retraction testing Undercarriage configuration Bogie units Shock absorbers Liquid spring oleos Gas oil oleos with separator Gas oil oleos without separator Servicing equipment

-

- --- -

LANDING GEAR

The functions of the landing gear are:

* To provide a means of manoeuvring the aircraft on the ground.

-k To support the aircraft at the correct height for loading, propeller clearance etc.

* To absorb the kinetic energy of landing and control recoil.

* To slow the aircraft after landing.

Once airborne landing gear serves no useful purpose and is so much "dead weight". Because it has a comparatively large mass and plays no part in the actual flying of the aircraft several attempts have been made to replace it. For example for take-off, trolleys which are left behind, have been used but landing is a problem.Other systems have been tried but so far no practical alternative has been found to the undercarriage (or landing gear) as we know it:

The geometrical arrangement of the undercarriage units on the aircraft is not standard and can include: I

- * Tricycle. Two main units and a nose or tail wheeljunit. Helicopters and fixed wing aircraft. - - -

I I ;

J; Four undercarriages - 2 mains and 2 nose units - some helicopters.

* - Multi bogie units housed-within the fuselage and a nose unit. Large

transports. -- - -- -- -

Tricycle Configuration

The majority of aircraft are of the tricycle type with either a nose wheel or a tail wheel. The main units are usually housed in the mainplanes (for lateral ground stability), but on some aircraft these are housed in the fuselage. On some nose wheel type aircraft there may be main landing gear in the wings (wing gear) and in the fuselage (body gear) with a total of 3 (DC10 - two mains and one centre fuselage), or 4 main gear units- (B747 - two wing and two body gear).

Main landing gear is positioned to give ground lateral stability and is near the aircraft's longitudinal C of G - for nose wheel aircraft just aft of the C of G and for tail wheel aircraft just forward of the C of G.

Initially, tail wheeled aircraft gave the propeller of powerful fighters clearance from the ground with out having excessively long landing gear units.

rnoodull l A-1 201

In general, the nose landing gear arrangement as opposed to the tail wheel system is preferred because:

* It provides better pilot vision. * It means the aircraft is already in the normal flying attitude - less

drag, and thrust is in the correct direction (not pointing downwards). * It provides easier loading. * It prevents "nose-over" during heavy braking. * It reduces the possibility of ground looping (a sort of over-steer where

the pilot steers the aircraft a small amount from the dead ahead position and it will steer too far - caused by the fact that the C of G is behind the main wheels).

The requirements for landing gear is laid down in JAR25 and covers the following:

* General requirements. * Shock absorber tests. * - Limit drop tests. ,

L ,* - Reserve energy absorption drop tests. * Retraction mechanisms. * ; Wheels, tyres and brakes. * Nose wheel steering.

The undkrcarriage complete comprises: -

* , A leg, pin jointed to the aircraft structure (retractable). * Wheels and tyres. '

J; For most units, typically main - a brake system. * A means of absorbing the landing shocks and controlling the recoil. * - --kmeans of locking the retractable unit in the u p and in the down

position. * A steering system and castering ability (nose units). * Fore retractable units, a means of retraction and extension - usually

hydraulic. -k An indication and warning system.

It is not usual to incorporate a brake unit in the nose or tail wheel but some aircraft do have them (B727 has a brake unit in the nose landing gear for example).

Loads Sustained by the Undercarriage

* Compressive. Static loads when stationary and dynamic loads when landing.

* Bending. Particularly when braking but also when taxiing round corners, landing in side winds etc.

* Torsional. During taxiing and when turning.

rnoodull ?A-1 202

* Tension. Some of the side stays of the landing gear are designed to withstand tension when braking and when cornering.

These loads can be substantial particularly compressive loads when landing and are transmitted to strong points on the airframe.

Types of Undercarriage

The wheels are mounted on axles attached to the bottom of the leg so that they transmit compressive loads to the shock absorber. There are two basic methods of mounting the shock absorber (figure 1):

Direct Acting. The shock absorber (oleo) may be an integral part of the leg or it may be a separate unit within the leg, but an any rate the compressive forces are transmitted to it directly up from the axle. If it is an integral unit it is designed to take all the additional stresses, bending, torsional etc, as well as the compressive loads. -

- - -

Torque liliks are required to prevent axial rotation of the sliding portion of the leg within the main unit. The torque links keep the wheels pointing in the right direction. Figure 5 shows the use of splines for this purpose - this method is rare.

Levered or Articulated Suspension. The oleo is a separate unit and mounted between the un-sprung part of the leg and the sprung part. Loads (mostly compressive) are transmitted to it by a lever arm connected to the wheel. This means that u p and down wheel movement is larger than oleo movement. As the oleo is subject only to compressive loads it may be made smaller and lighter and is usually easier to change.

- - - -

This type of suspension lends itself to being used on nose wheel and tail wheel assemblies as it can be made to caster to from a trailing unit. (Caster = to turn in the direction the aircraft is going.)

AIRFRAME ATTACHMENT

LEVER

LEVERED SUSPENSION DIRECT ACTING

Fig. 1 TYPES OF UNDERCARRIAGE

- j -

rnoodull l A-1203

Camber (figure 2)

This is defined as the inclination of the wheel with respect to the vertical plane when viewed from the front or the rear of the aircraft. Most large aircraft have a zero angle of camber for all wheels but many small aeroplanes do have cambered main landing wheels.

Positive camber is where the wheel is inclined away from the aircraft and negative camber is where the wheel is inclined towards the aircraft. Camber i s checked by using a clinometer (adjustable spirit level) placed on the rim of the wheel (using a straight edge where necessary). The aircraft should be jacked for this operation - but check the AMM.

Fig. 2 CAMBER

rnoodull l A-1204

NOSE UNDERCARRIAGES

To some extent what applies to nose landing gear also applies to tail landing gear.

The nose undercarriage is usually of lighter construction than the main gear because it normally does not have to take the initial landing shock and on most aircraft also does not have to take any stress related to braking. It does, however, have to carry provision for towing and must withstand bending and shear loads.

It may be a single wheel unit or have two wheels.

QUESTION Can you list four or five features that a nose undercarriage has to have that a main gear does not have? (5 mins)

ANSWER For most large aircraft the nose unit will have:

J; The ability to caster. -k A self centring system. * A steering system - powered for large aircraft.. * An anti-shimmy device.

I * No brakes - usually.

-

Castering --

To enable the aircraft to be manoeuvred on the ground the nose'wheel must have the ability to align to the direction the aircraft is going. It must caster freely to at least the maximum angle the aircraft-is allowed to turn. There will normally be internal stops to limit its angle of travel and markings on the outside will give an indication to personnel towing the aircraft of the maximum limit of travel. On no account must any attempt be made to exceed this limit.

Self Centring

Automatic self centring is essential as soon as the aircraft's weight is removed from the wheels to ensure correct positioning of the wheel/s prior to retraction. If the wheel/s of the nose gear are not lined up correctly on retraction considerable damage will be caused to the structure and the landing gear.

Centring is achieved by either a spring loaded cam or a hydraulic centring jack of through the steering mechanism.

With the spring loaded cam system - at the top of the castering unit there is a V shaped cam down into which is pressed a spring loaded roller. When the leg casters the roller is forced up one side of the V slot against the pressure of the spring.

When the weight of the aircraft is removed from the wheels the spring forces the roller down into the middle of the V slot rotating the leg into the dead ahead position.

To prevent the nose gear retracting should it not be in the dead ahead position the down lock is usually connected to the castering system mechanically so that if it is not dead ahead then the down lock will not disengage.

Centring jacks, either as a separate unit or fitted in the steering motor, rely on hydraulic system fluid pressure when up is selected to centre the leg. Again, the down lock is inter-connected to prevent the leg retracting should it not be centred.

Steering

Some small aircraft do not have a steerable nose wheel (or tail wheel) they are steered using differential braking. Apply the brakes on the left hand side and t h ~ aircraft will steer left - apply the brakes on the right and it will steer right. Many small'airEraft have nose wheel steering systems which are mechanically connected to the rudder bar. When the rudder bar moves so does the nose leg - in the air;- as well as on the gound!

I

Large $rcraft have powered nose wheel steering, this is because the wheels would be imGossible to move manually.

QUESTIPN How much weight is felt through the nose undercarriage? (2 mins)

ANSWER T h e actual weight (or more correctly 'mass') that is felt through the - - - -

nose gear varies considerable - depending on the total mass of the aircraft, but it is usually in the region of 10%. As an example taking a reasonably small large aircraft of 100 tons - eg Airbus A 3 10 (about 100 tonnes - one tonne is about 2% heavier than a ton). If 10% acts through the nose gear that makes 10 tons - too much to be moved manually.

Earlier methods of steering involved differential braking. Its disadvantages compared to steerable systems include:

* Lose of forward momentum when brake is applied. * Increased engine wear when engines are used to get the aircraft

rolling again. * Increased fuel consumption. * Increased brake and tyre wear. * Not so easy to steer aircraft.

Powered steering using hydraulic jacks allows the engines to be set at minimum thrust for taxiing thereby saving fuel, brake and tyre wear.

To allow the nose unit to caster when towing the aircraft there must be some means of disconnecting/isolating the powered steering. This can be done in several ways:

* A by-pass valve fitted in the system (figure 6). When the steering is switch off the valve is opened allowing free movement of fluid between both sides of the jack when it moves due to the nose gear being turned when towing. When steering is selected the valve is closed by hydraulic pressure.

A By disconnecting the torque links during towing/pushback. This allows the unit to caster without moving the steering jack.

J; By the use of an "isolating pin" which is inserted into a special hole in the nose unit by the pushback crew - this effectively isolates the nose gear steering mechanism. The pin is removed after pushback and the usual procedure is to show it to the pilot before he/she , , taxies the aircraft away.

Power steering systems are usually operated hydraulically usink main system to operate either:

(a) A single jack with an equal area both sides of the piston, or

(b) Two jacks of the unequal area type (figure 3). -

-

--- - -

SINGLE EQUAL AREA JACK [e -- STEERING u PLAN VIEW

DUAL UN-EQUAL

When the steering wheel is moved fluid pressure is supplied to the large area end of one jack while the other is connected to return.

Fig. 3 STEERING JACK ARRANGEMENT

rnoodull lA-I207

QUESTION Why does the single jack have the same area both sides of the piston? (2 mins)

ANSWER If the areas were not equal the jack would move faster one way than the other (given the same fluid supply conditions).

QUESTION With reference to the drawing below. The jack is of the unequal area type. Which way will it move the fastest - in or out? (2 mins)

IN 4-----c OUT

ANSWER It would move IN the fastest. The reason is that there is less volume - -

to fill as some of the volume is taken up by the piston ram. Remember, that the rate of movement of a jack is governed by how fast the jack volume can be filled. Given a fixed hydraulic supply and a certain size jack the rate of fluid flow governs the speed of movement and the pressure governs the load that can be moved.

Control I

Steering is controlled from the flight deck, depending on the aircraft, by:

la) A separate steering wheel on the side of the flight deck. (b) The rudder pedals. (c) The aileron control wheel.

Control is provided to both pilot and co-pilot positions.

The steering control will select a hydraulic four way selector to port pressure fluid to one side of a jack whilst opening up the other side to return. The jack/s may be a simple jack as shown above or may incorporate a:

* Self centring jack. J; Shimmy damper. * Relief valves.

Feedback

The control mechanism is not too unlike that of a Powered Flying Control (PFC) system - in principle at least. I t cannot simply be selected one way o r the other, it has to have some form of feedback.

moodull l A-1208

QUESTION There are two types of feedback - can you name them? (2 mins)

ANSWER Positive and negative feedback.

QUESTION PFCs and powered steering use negative feedback in their control systems. Can you describe negative feedback? (5 mins)

ANSWER This is where a signal is put into a control system to produce an output. Part of the output is then feed back (negating signal) into the input to try to cancel the original input signal. The stronger the original signal the greater the negating signal. Incidentally, almost all control systems use negative feedback,.from biological control systems to mechanical control systems to electrical control systems to electronic control systems.

Feedback on a power steering system can be achieved by: - - - -

(a) Mechanically - using a system of cables and pulleys (Boeing aircraft for example).

(b) Electrically - using a Wheatstone bridge. (c) Electronically - using a computer (A320).

More of this later. --

Shimmy Damper

Shimmy is a problem associated with-noseltail wheel units and is a form of vibration of the unit about its rotational axis. More common in non-steered units with incorrect tyre pressures and worn bearings being contributory factors. Because the unit is allowed to caster and the flexible nature of the side-walls of the tyre an unstable swivelling oscillation can be set u p in the castering part of the unit. Excessive shimmy, especially at high speed can cause serious vibration throughout the airframe and can be dangerous.

Shimmy can be damped in several ways:

(a) By the provision of hydraulic damping within the steering motor. (b) By the provision of a separate hydraulic damper. (c) By the fitting of heavy self centring springs. (d) By fitting double nose wheels.

H~draulic Damping. This is achieved by fitting an oil filled damper to the castering part of the gear in such a way that as the unit casters (or is steered) a piston is caused to move through the oil (figure 4). The piston has small oil ways through which the oil passes.

ATTACHMENT

ATTACHMENT TO NON MOVING PART

Fig. 4 HYDRAULIC DAMPER

When shimmy tries to occur and the unit tries to oscillate the piston tries to move back and forth through the oil quickly - the frequency being relative high and of low amplitude. Because of the viscosity of the oil it resists these quick movemen and 'seems! to "lock-up" the damper.

I I

With ordinary castering/ steering the piston moves with the oil flowing steadily through the ports.

QUESTION What does viscosity mean? (2 rnins)

ANSWER It is the resistance to flow of a fluid. Related to friction within the I fluid.

, -

-- -- ppp -

Twin Wheeled Units. These assist in reducing the tendency to shimmy as each tvre has its own natural shimmy frequency (associated with wheel rpm). When <here are two wheels their natural frequencies rarely coincide, one tyre tends to dampen the vibrations of the other and shimmy is eliminated or significantly reduced.

blank

POWER STEERING SYSTEMS

These are normally hydraulically operated and mechanically, electrically or electronically controlled.

The hydraulic supply comes from one of the hydraulic systems. On some aircraft there is no back-up supply as steering can be achieved using differential braking if necessary - and the brake system will have a main supply with a standby.

The steering system comprises:

* A hydraulic steering jack/ actuator/ motor fitted to the nose leg. * A four way hydraulic selector valve. J; A steering tiller/ control wheel. k Control cables or circuitry/ computer to provide control and feedback

and possibly data to other systems (retraction for example). * Hydraulic system to include pressure relief valves etc.

-

When a steering command is imputed the selector valve selects, the jack moves and the negative feedback de-selects to stop the jack movement. When power steering is switched off or isolated the unit is free to caster. Immediately prior to retraction the system is switched off and the unit centred.

-- A Single Jack System (figure 5) I

This shows one method of fitting a jack to the nose gear. In this arrangement the jack ram is fured to the non-steering part and the jack body is connected to the steering part. In other systems it is fitted with the jack ram moving the steering Part.

- - -

Remember that splines are very rare, torque links are far more common - connecting the top part of the rotating unit to the bottom shock absorber part.

The fluid connections to the jack in figure 5 are through the jack ram ends. As fluid pressure is ported to one side of the jack the body moves and through the connecting link steers the wheels to port or starboard. In this aircraft the follow- up linkage (negative feedback) is an electro mechanical system using a drum switch moved by the steering jack.

blank

STEERING CONNECTING

torque links are used)

SHOCK ABSORBER (OLEO)

LE FITTING

TOWING CONNEC1

Fig. 5 STEERING JACK ARRANGEMENT

A Centring Jack

On retraction it is essential that the unit is centred automatically. This is the function of the centring jack.

This jack may be a separate unit or fitted as part of the steering jack. It comprises a piston rod anchored to the undercarriage leg, a jack body connected to the steering part and an inner piston.

- 12 -

rnoodull l A-1212

When the undercarriage is selected up, the steering system is isolated and fluid pressure is fed to this jack. This causes the inner piston to move to the end of its travel in one direction and the jack body to move to the end of its travel in the other direction. In this condition the nose leg is centred.

On some aircraft the steering system is isolated by the weight switch as soon as the aircraft takes off.

A Hydro Mechanical System (figure 6)

In this system the single jack moves the nose wheels via a rack and pinion gear (similar to the Airbus system). Selection is made in the flight deck via the steering cables to operate the steering drum. The feed back mechanism is via the follow-up cables - for example:

1. Pilot steers to the right. 2. Steering drum rotates-anti-clockwise. 3. The left hand side of the beam is pulled down. -

4. Movement of the beam; anti-clockwise causes the selector valve to select porting pressure fluid to the bottom of the steering jack and opening the top to return.

5. The jack moves up to cause ---- the wheels to steer clockwise (to the right). r- -

6. ' This puts tension into the right hand side of the follow-up cable and slack into the left hand side - releasing the tension in that side.

7 . This pulls the beam back to its level position, neutrhlises the selector '

valve and stops the steering action. -

If the pilot continues to move the steering input, the selection will be maintained with the follow-up cables trying to catch up the steering cables - and cancel the input. A s soon as the pilot stops his/her input the follow-up system catches up with the input system the valve goes into neutral, movement stops and the jack is held in that position by an hydraulic lock.

NOTES

1. The selector valve is shown disproportionately large. 2. The cable system is shown simplified - for example there is a gearing

system behind the steering wheel, but the principle is similar to that fitted to some Boeing aircraft.

The throttling valve provides a constant rate of flow of fluid supply

QUESTION Why would the flow vary from the hydraulic system? (10 mins)

ANSWER For 90% of aircraft the hydraulic circuits - flaps, slats, landing gear, brakes etc - are connected in parallel and then connected to the supply circuit. That is, they all experience the same pressure (unless there are pressure reducing valves fitted) but they share the total available fluid flow. If one of the circuits is selected - say flaps - the fluid flow rate to the power steering will be reduced (and incidentally also to the flaps). When no other services are selected the fluid flow rate to the steering circuit is high and the throttling valve will move to the nearly closed position. When other services are selected the fluid flow rate is reduced, the valve is more open and the flow to the circuit is the same as before - this ensures a constant speed of operation irrespective of the selection of other services.

Fig. 6 SIMPLIFIED MECHANICALLY CONTROLLED STEERING SYSTEM

- 14 -

rnoodull lAp1214

The by-pass valve may be actuated hydraulically of electrically and when open allows fluid to pass freely from one side of the steering jack to the other (for towing, pushback etc). In the drawing it is hydraulically operated so when steering is selected pressure to the top of the valve causes it to close. When unpressurised fluid from either side will pass easily through the lightly loaded spring valve.

A two way pressure relief valve allows the relief of any excessive pressure on one side of the jack to be relieved to the other side - possibly caused by the tyres catching a taxiway/runway light.

On retraction the system will have provision to ensure the unit is centralised using a centring link or a centring jack.

An Electrical System

The hydraulic side of the system-is similar to the one described-above but the control and feedback is electrical. The selector valve is electrically operated and the feedback is provided by a Wheatstone bridge (figure 7).

The bridge has four resistors, two of which are a fxed value and two are variable. It is supplied with 28Vdc.

-

In simple terms if all the resistors are the same value then the voltage (and current) in both arms of the bridge will be the ,same. If the voltages between the left and right arms of the bridge are compared'then they will be the same and there will be no voltage difference at the amplifier - and no output to the hydraulic selector. -

- - -

28V dc SUPPLY

RESISTORS

AMPLIFIER

Fig. 7 ELECTRICAL STEERING CONTROL

- l h -

rnoodull lA-1215

If the bridge becomes un-balanced (by movement of the pilots input resistor) then more current flows down one arm of the bridge - there is a voltage difference across it, this is amplified and selects the valve. The undercarriage moves and the valve will be de-selected when the follow-up resister reaches the value of the input resister.

For example, when the pilot moves the steering control wheel it will alter the steering input variable resistor (say UP). This means that more voltage is dropped across that resistor so the right hand side of the bridge has a higher voltage (and has more current). This voltage difference is felt by the amplifier to pull in a relay (left hand one) within the selector valve. The selector valve selects to move the nose leg in the appropriate direction. When it moves it changes the value of its follow-up resister - in this case UP.

While the pilot continues his/her input the follow-up resister will continue to chase the input resister. When the hand wheel is stopped moving, the input resistor stops and a fraction of a second later the follow-up resistor reaches the same resistance value. This balances-the bridge, the amplifier senses no voltage difference, no signal is sent to the selector valve, it de-selects and the nose gear stops moving.

Notes. 1. For more details of ;the electrically operated four way selector valve see the book in this' series entitled Hydraulics. 2. The follow-up resistor (potentiometer) is usually fitted directly on the nose leg and the rest of the bridge is in the aircraft somewhere -

' probably behind the instrument panel.

A Computerised System

Again, the hydraulic side of things is very similar to the system shown in figure 6, it is the selection and control that is computerised.

You will have to take my word for it that a digital computer is basically a very simple machine. It has many things in its favour such as small size, low power consumption, inexpensive, speed of operation etc, but the actual operations the micro processor can do are strictly limited. The digital functions it can perform are :

1. Addition. It can add in binary. All other mathematical functions are performed using an addition process.

2. Store. It can store large amounts of digital data. 3 . Move. It can move this data around. 4. Compare. It can compare digital words.

And that's all. There is no need to discuss this any further but if you want more infomation on the processes then a ten minute 'phone call to your tutor will explain all.

The computer will take the steering input signal and convert it to a digital signal and put in store (2.). It will convert this signal to an analogue signal to select a hydraulic selector valve - this moves the nose gear which will send an analogue feedback signal back to the computer. This is digitised and this feedback signal is compared (4.) to the orignal input signal. If the two signals are not the same the selection is continued and when they are the same the steering signal will stop.

Figure 8 shows a schematic of the A320 system. The pilot's steering input is sent to the BSCU (Brake and Steering Computer Unit). The steering signal can come from the rudder pedals that can be isolated by the rudder pedal isolation button on the steering wheel.

GREEN POWER FROM NOSE GEAR DOORS CLOSING CIRCUIT WHEN -f DOORS ARE CLOSED

POSlTlON FEEDBACK

Fig. 8 THE A320 STEERING CONTROL SYSTEM

QUESTION With reference to figure 8 what other computers have an input into the BSCU? (2 mins)

ANSWER The autopilot and ELAC (flying control) computers.

rnoodull l A-121 7

QUESTION Again with figure 8 - when can power steering be used? (5 mins)

ANSWER It can be used only when:

(a) It is switched on, AND (b) Airspeed is greater than 80kt, AND (c) One engine is running, AND (d) Towing lever is in nonnd position, AND (e) The Weight On Wheels (WOW) switch is made.

With all these conditions made the AND gate will allow steering.

UNDERCARRIAGE SYSTEMS

In general the landing gear system comprises:

(a) A retraction system to raise and lower the gear and doors. ( b ) ' A reliable locking system to operate in the up and the down positions. (c) An indication system to indicate to the crew the status of the landing

I gear. (d) Wheels, tyres and brakes (covered in other books in this series). (e) Systems to interconnect the gear with other services such as low

1 I

I speed warnings etc. , 1

I

Retraction and Lowering Systems I

The most common form of retraction system is hydraulic (see the book in this series entitled Hydraulics), butsome aircraft may use electrical actuation and others may use pneumatics (for pneumatic systems refer to the book in this series by that name).

Provision must be made for the emergency lowering of the gear should the main system fail. Emergency/ back-up systems can include:

* A second hydraulic pump in the power supply system. J; Standby pumps - ram air driven for example (RAT). ~i Duplicate (or even triplicate) hydraulic supply systems. * The use of a compressed nitrogen blow down system. -k Gas operated release locks and free fall of the gear with spring

assisted down locks. * A hand operated wind down system (not common).

- 18 -

rnoodull lA-1218

Locking Methods

JAR25 states that there must be a positive means to keep the landing gear extended in flight and on the ground and to keep it in the correct retracted position.

QUESTION Can you name 3 landing gear locking systems? (3 mins)

ANSWER Mechanical. Hydraulic. Geometric.

Mechanical locks are the most positive and reliable so meet the requirements of JAR25. The actual design can vary from a plunger moving into a machined hole between two links of the landing gear-to a hook type unit locking onto a bar. Nonnally actuated hydraulically and usually fitted with a micro switch/proximity switch for flight deck indication systems. I

A hydraulic lock might also be used - but normally only as a back-up as it is not very reliable.

--

I

QUESTION What is a hydraulic lock and whySis it unreliable? (5 mins)

ANSWER A hydraulic lock is caused by a trapped column of fluid between a jack or hydraulic actuator and a control valve or other valve. This fluid, because it is not compressible (at most working pressures), will prevent the jack from moving.

It is not reliable over the long tern because of pressure changes due to temperature changes and because of the possibility of seepage. In normal hydraulic systems operation some internal seepage is permissible - it is so small that it is not noticeable, but to rely on hydraulic pressure only over a long term (without a continuous supply) would not work.

A geometric lock might also be used - also as a back-up (figure 9). A geometric lock is a series of links - usually three including the undercarriage leg, and so arranged as to try and stay in one locked position. The links can be moved out of the locked position by a jack.

By it self a geometric lock is not reliable as there is the possibility of it being shaken out due to aircraft movement.

The side stays are geometrically locked as gravity will keep them in position (some form of small

MAIN LEG mechanical lock is provided to prevent them being shaken out of position). A hydraulic jack or some other device will be required to 'break' the lock.

0 = HINGE POINTS

Fig. 9 GEOMETRIC LOCK

Selection of the undercarriage rnay be by a direct mechanical connection from the flight deck to the two way hydraulic selector valve or by an electrical connection betwekn the flight deck and the :valve. Mechanical operation is rare these days as electrical systems are much lighter.

Position indication of the landirig gear (UP or DOWN) is signalled to the flight deck by a lamp system; on CRTs on some aircraft, and there is usually a mechanical back-up indication system.

I

Leg and fairing doors are usually moved by separate jacks with sequencing being carried out hydraulically, electrically or mechanically.

Indication and Warning Circuits

Provided to give an indication in the flight deck of the leg position/door/s position - usually in the form of coloured lights and, if a CRT display is provided, by a small display.

With reference to figure 10. Micro switches or proximity switches are operated by the up and down locks and are wired into indicator lamps in the flight deck. There is usually one lamp per indication per undercarriage (with CRT indication as well on some aircraft). Where lamps are used standby lamps may be provided. In basic terms lamp indications are:

Green - Unit locked down Red - Unit unlocked No light - Unit locked up

moodull l A-1220

In some aircraft an amber light will show if the aircraft is on the ground and the fairing door is moved from its normal position (for servicing).

To warn the pilot if the gear is not locked down on approach an additional warning system is provided. This may be via the throttle switch, as in the drawing, or it may be via the ground proximity warning system or wired into a Pitot pressure operated micro switch.

On many aircraft mechanical standby indicators are also fitted (figure 12).

QUESTION Describe the difference between a micro switch and a proximity switch. (15 mins)

ANSWER A micro switch is a mechanical switch that opens or closes when contact is made with its operating plunger. A proximity switch may have no moving parts (depending on type) - and do not rely on contact for their operation, On one type a magnet (target) is fitted to the moving part of-the landing gear - on the other part-a proximity reed switch is fitted., As the target comes into close proxi&ity with the switch (as the leg retracts for example) so the reed switch is caused to move and make or break a dc circuit.

1 -- -

On an inductive type proximi@ switch system the target is a piece of ferro-magnetic material. The switch is a simple coil with an ac supply When the target gets close to the switch the coil's inductance (resistance) increases and a sensing unit in the circuit will pick this u p and give the appkopriate indication on the flight deck.

---

- . - -

Figure 10 shows the gear in the unlocked position with a 28Vdc supply to the red lamp via both micro switches.

If the gear is on the way up then the next thing to happen is the uplock micro switch is operated breaking the circuit to the red light and the light goes out.

If the gear is on its way down the next thing to happen would be that the downlock micro switch operates to switch off the red light and switch on the green.

If the throttle is pulled back past a certain limit (for landing) with the gear not locked down then the throttle micro switch is live and will make contact giving an aural and visual warning as well as signals to the CRTs and CONFIG warnings etc.

Most modern aircraft have a configuration warning system to check that all system are set correctly for a particular operation - eg landing, take-off etc.

rnoodull lA-1221

For signals being sent to lamps the current goes directly to the lamp and then to earth return. For the CRT displays the signals (discrete ON or OFF) are sent to the Symbol Generator Unit (SGU) - a computer with stored symbol data for transmission to the CRTs.

THROTTLE ** To other landing gear units

UPLOCK DOWNLOCK MICROSWITCH MICROSWITCH

FUSE

CRT DISPLAY

INDICATION

I w WARNING SIGNAL

Fig. 10 A SIMPLIFIED INDICATION & WARNING SYSTEM

Figure 11 shows a typical modem flight deck indication and warning system. It shows the A320 system using ECAM displays (Electronic Aircraft Centralised Monitor). This is a computer based system using CRTs and SGUs to show coloured symbols and pictures on the screen. The Landing Gear Page must be selected on the control panel for gear data to be displayed.

\ o r / - 2 P 2 2 2 1 extension

lnr -19 'c G.W. 60300 n; snr * l a 'c 23 n 56 C.G. 26 I :

WHEEL SYSTEM PAGE LANDING ROLL

Landing Gear Brake pressure Seleclor ,Lever lndlcatlon

(allernate system)

Fig. 11 A320 LANDING GEAR CONTROLS & INDICATIONS

- -- -- - -

Screen displays can include, when selected:

* Hydraulic systems. * Cabin conditioning. * Engine indications etc.

Landing Gear Selection

The selector system might be a simple cable or push/pull rod system connecting the flight deck lever to the hydraulic selector valve. More likely on modern aircraft it will be an electrical system.

The selector lever will be tactile (that is shaped like a wheel so it can be found in the dark - if necessary). It will also be instinctive - moved up for up selection and down for down. When operated it will make an electrical contact to operate a solenoid in the selector valve and will have a safety interlock that can be overridden in an emergency.

FREEFALL HANDLE

VISUAL DOWNLOCK

FLOOR HATCH

WINDOWS

HANDLE

Fig. 12 STANDBY MECHANICAL INDICATORS

The safety interlock is provided to prevent inadvertent operation of the up selection-when the aircraft is on the ground. This may be overridden by the pilot in an emergency such as during landing and all the brakes fail and the only way to stop the aircraft before it comes off the runway (say into the sea or nearby buildings) is to retract the gear and letting it slide on its belly - hopefully to a safe stop.

In the simplified system shown in figure 13 the selector lever makes contact in either the up or down position to provide a 28Vdc supply to solenoids in the electro-hydraulic selector valve. This will select to move the gear to either the fully up or fully down position.

When down is selected and the weight is on the wheels the safety interlock (weight on wheels lock) solenoid circuit is broken and it moves out under spring pressure. In this position it prevents up selection being made. When the aircraft takes off the circuit is made by the squat switch, operating the solenoid and withdrawing the weight on wheels lock, the undercarriage can be selected up. The emergency override is a button which allows for gear up selection when on the ground - it is rotated to re-align the two legs' either side of the weight on wheels lock and pushed in.

Note. The squat switch may be called:

* Weight on wheels (WOW) switch (Airbus). J; Ground sensor (tockheed and Douglas). * Ground air sensor (Boeing) . * Flight ground switch (KLM).

Figure 13 is a simplified drawing that shows the principle of how the system works. Figure 14 shows an actual selector as fitted to a B747-400. Note that the selector panel has the following differences:

J: Provision for OFF selection which depressurises the landing gear system.

* Has a lever lock override switch that releases the landing gear lever.

* Has alternate gear extension switches. -- --- - -- -

-

I

Retraction Testing

These will be required after any work is carried out on any part of the gear retraction system and as specified in the maintenance schedule. The tests will verify the correct operation of the mechanical, electrical, electronic, and hydraulic systems. - - -

I I

\

LANDING GEAR SELECTOR 28V DC BUS LOCK OVERRIDE

UP SELECTION

WEIGHT ON WHEELS LOCK

SOLENOID I

, ELECTRICALLY OPERATED

1 Fig. 13 SELECTOR CONTROL SYSTEM

rnoodul l lA-1225

The actual testing to be carried out as per the AMM.

The aircraft must be jacked inside (or outside provided the forecast wind speeds are less than those specified in the AMM, the weather is fine and it is on hard standing).

The electrical and hydraulic systems must be serviceable and barriers erected and warning notices posted to keep people away from the area around the landing gear. Place look-outs.

In general the procedure is:

Check the AMM. Jack the aircraft. Ensure shock absorbers extend to full length and bogies rotate to correct position. Remove the ground locks. Connect external electrical and hydraulic supplies. Ensure hydraulic supply is a t the correct pressure and volume flow rate. Note that some aircraft will have more than one external rig fitted - say one for the green system and one for the blue. If in doubt about clearances, sequencing etc try movement of gear using a hand pump. Select landing gear up and down and check:

* Time up and time down. * Flight deck and standby indications. J; Any warning de;vices/warning configurations. * Correct sequencing - eg, Up selection, door open, leg up, door

closed. Down selection, door open, leg down, door closed. 3; Correct operation of up locks and down locks on legs and

doors. * Correct fitting of doors and legs in the closed position. * Correct clearance between structure, wheels and gear. ~r Correct rotation of any 'folding' part of the gear - bogie ro ta t io~

leg rotation etc. * Correct interconnection of gear retraction system and steering

system. * Nose leg centres before retraction. J; Auto brake is applied before main wheels are retracted. * Operation of retraction system with one supply system on only

and then with the other. J; Test any standby systems - electric pumps etc. * Check any free-fall/emergency systems for correct operation.

Note that if the system is a blow down system using compressed gas the pressure will have to be relieved, the system bled and topped up. Re-set the system and carry out normal retraction test.

x Smooth operation and leaks.

6. Lower the undercarriage and check that there are "three greens" with all other flight deck indications correct.

7 . Ensure all down locks are engaged correctly. Fit ground locks. 8. Disconnect external electrical power and hydraulic test rigs. 9. Lower the aircraft off jacks and chock the wheels. 10. Record all the work done and sign the appropriate documentation.

Steering Testing

Refer to the AMM. Raise the aircraft on jacks, place 2 steel plates under the nose wheels with grease between them. Lower the aircraft ensuring that the nose wheels are on the plates. Connect external hydraulic and electric supplies. Check that nose wheels caster with steering selected off. Select power steering on. Operate the steering wheel over its full range of travel and check that the nose gear follows in the correct sense and at the same rate of movement. I

Check for smooth o~eration of the undercarriage p d that there is adequate clearance between moving parts and no strain on any cables, swivelling pipe connections etc. Check for leaks. -- -

Check steering operates with the rudder pedals -if provided for. Jack the aircraft clear of the ground. Set the nose gear a few degrees to' one side of the ,dead ahead position and select landing gear up. Check nose gear centi-es before the down lock disengages. Lower the landing gear and repeat item 12 but with the nose gear displaced in the opposite direction. With the gear up check that steering is inoperative. Lower the gear. Check any standby supplies to the steering motor including electric pumps, hydraulic accumulators etc. Check any warning/indicator systems. Remove the steel plates by raising the aircraft or towing the aircraft forward a little. Record all the work done and sign the appropriate aircraft documentation.

UNDERCARRLLZGE CONFIGURATION

The increase in size and All Up Weight (AUW) of modern aircraft has led to an increase in wheel loading; this being defined as the static load on each wheel of the landing gear at aircraft take-off weight. This term is not very precise when calculating the actual force per unit area (psi) that each tyre puts on the ground.

rnoodull l A-1227

LANDING GEAR LEVER

UP - releases downlocks - pressurizes up s i d e o f gear ac tuators

OFF - depressurizes land ing g e a r h y d r a u l i c system

DN - releases uplocks - pressurizes down s i d e of gear actuators

.LEVER LOCK CVERRIDE SWITCH

PUSH - releases l e v e r lock

ALTERNATE GEAR EXTENSION SUITCHES

DH ( a l t e r n a t e act ion, guarded) ALTN - releases gear doors and gear

uplocks - GENTER PANEL

\ c-

Fig. 14 BOEING 747-400 SELECTOR PANEL

A more precise term, called the California Bearing Ratio (CBR) indicates the ability of a runway surface to support a load.

I

In general the lighter the aircraft the less pressure is required in each tyre to support the weight, the larger the tyres the less pressure is required and the more tyres there are the less pressure is required. Remember, if pressures are too high runways and taxi ways will be unable to support the load.

- -

So in general for large aircraft a landing gear will need a large tyre at a certain pressure or several smaller tyres at the same pressure.

The actual configuration of the landing wheels will depend on the designer to cot with the problems of stowage, safety, wheel loading and CBR.

Wheel loading can be reduced and CBR improved by have multi wheeled units using twin tyres per landing unit; three wheels; four, six, eight or even twelve.

Some aircraft have many main wheel units (several each side of the fuselage) with twin tyres on each so making up to a considerable number of main wheels.

The advantages of a multi wheeled unit over a single wheeled unit are:

* Reduced wheel loading (though this will depend on other factors also). * Easier to service. Although more complex the wheels are smaller and

easier to handle.

- 28 -

moodull l A-1 228

* Easier to stow. Bogie units can be rotated, turned etc to reduce their stowed volume. Some units are shortened by a shortening link within the leg - though this could also apply to single wheeled units.

k Greater safety factor. In the event of a tyre burst for example there will still be one or more tyres remaining to take the load. Even so, on any landing gear a tyre burst on landing can be critical.

A Better braking. Each tyre is braking on its own piece of runway which might have slightly different conditions to the next tyre so each is braking a t its own best efficiency.

I SINGLE WHEEL DOUBLE WHEEL BOGIE UNIT (4 ,8 OR 12 WHEELS

Fig. 15 WHEEL CONFIGURATION - -

I -- -

The dis-advantages are:

* More complex and more expensive. Will involve more wheels, more brake units, more anti-skid units, more cooling fans, more tyre pressure transducers etc.

* Larger footprint area. When the aircraft is turning around a comer the inside main leg of the corner experiences a significant twisting force about the leg centre line. Because of the large footprint this can put considerable stress in the unit and can cause damage to parts such as the torque links and scrubbing of the tyres. This means that in general bogie main wheeled aircraft have larger turning circles than single main wheeled aircraft - size for size.

An interesting solution to this problem was found by the manufacturer of the Trident aircraft. On this aircraft the four main wheels are d l fitted on the same axle so when carrying out a tight turn the wheels could all rotate independently. I t also had the added advantage of the unit being able to fit into part of the fuselage fairing on being rotated (figures 16 and 17).

rnoodull l A-1229

Note the location of the nose wheel - off centre. For space reasons it had to retract sideways and the additional stresses on the fuselage were found to be negligible so it was fitted as shown.

Fig. 16 THE TRIDENT MAIN GEAR -

L/ MAIN GEAR TURNS ABOUT ITS AXIS TO RETRACT SIDE-WAYS INTO THE WING

Fig. 17 RETRACTION OF TRIDENT MAIN & NOSE GEAR

Bogie Units

Normally confined to main landing gear units, the design of which can vary but in general, comprise a telescopic leg containing a shock absorber and a bogie beam pin jointed at its connection to the shock absorber. This beam carries front and rear axles for four or eight wheels (figure 18). Torque links between the main leg and the beam or sliding portion prevent axial rotation of the unit whilst leaving it free to slide u p and down.

rnoodull l A-1 230

LEG MAIN 6 ATTACH ME^^ POINTS

I- BRAKE COMPENSATING

BOGIE BEAM '&&A\\\-

LEVELLING ---. ..-

Fig. 18 TYPICAL BOGIE TYPE UNIT

/ TO FIT INTO SLIM

Fig. 19 TUPOLEV MAIN GEAR RETRACTION

The beam may be in two parts connected via a bogie swivel hinge which allows better articulation during a turn and the bogie pivot allows the beam to tilt -

usually with the rear wheels trailing down (but not always). This tilt may be assisted by a bogie damper strut attached between the beam and the leg. This tilt is sometimes called assisted articulation.

The beam is usually fitted with a micro switch which prevents retraction in the event of incorrect tilt or articulation.

Figure 19 shows an example of how the bogie can be rotated to, in this case, fit into a wing section and figure 20 shows the main gear of the A300.

ATTACHMENT

TORQUE

w Fig. 20 THEA300 MAIN GEAR

SHOCK ABSORBERS (OLEOS)

Theseiare fitted to absorb the shock of landing. Some shock is absorbed by the tyres $ producing pressure energy, but most of the shock is absorbed by the shock absorbers. Some of the older units used springs or bungee cords to convert the downward kinetic energy into strain energy. Most modern units convert this energy into pressure energy (of a gas, or in some cases of an oil).

QUESTION The oleo works on converting the downward kinetic energy into pressure energy within the oleo. What happens to the fonvard kinetic energy of the aircraft? (5 mins)

ANSWER Most of it is converted into heat energy within the brakes. Some is used u p as drag (airframe, spoilers, flaps etc). Some is taken u p by the reverse thrust of the engines.

The oleo has three main functions:

1. To absorb the downward kinetic energy of the aircraft. 2. To control the recoil. 3. To support the static weight of the aircraft.

The energy absorption process is achieved by increasing the pressure within the unit as it collapses on landing and recoil control is achieved by controlling the rate at which fluid is allowed to pass through a piston.

The support of the static weight of the aircraft is achieved by the pressure within the unit times the area of the unit's piston/cylinder giving a force that is equal and opposite to the downward force of the aircraft on that landing gear unit.

There are three main types of oleos:

1. The liquid spring type. 2. The gas/oil type with separator. 3 . The gas/oil type without separator.

With the liquid spring type the energy of landing is absorbed by the increase in pressure and the compression of the fluid within the unit. Control of recoil is by controlling the rate of fluid flow passed a piston.

- -. - -

With the other two types the landing shock is absorbed by the increase in pressure and the compression oflthe gas. Recoil control is carried out by controlling the rate of fluid flow through a piston.

QUESTION Why is it important to con<rol the recoil? (5 rnins), -

- -

ANSWER If the recoil was not &ontrolled the pent-up e n e r q in the form of very high pressure within the unit would cause the oleo to extend quickly and bounce the aircraft back u p into the air -just like a bouncing

- -b&ll. Not what is required;-In general terms the unit will collapse - quicker than it is allowed to extend. - -

Each oleo consists of an inner tubular or solid member, fitted with a piston, telescoping into an outer tubular member. With the aircraft on the ground the leg telescopes sufficiently to create enough pressure which supports the static weight of the aircraft.

When the aircraft becomes airborne, its weight is transferred to the mainplanes and the oleo fully extends. When the aircraft lands the downward energy of the aircraft telescopes the oleo which absorbs the shock of landing.

THE LIQUID SPRING OLEO

Sometimes called and Oil Compression oleo. It contains only oil.

Not popular on civil aircraft because of its harsh ride characteristics.

In general fluid is considered to be incompressible but this unit works on the principle of compressing the fluid, how is that?

A t pressures up to about 3000psi the amount of compression (reduction in volume) of a fluid is small - for hydraulic system purposes it is considered as negligible. But as the pressures are increased so the fluid becomes more 'squashable' until at very high pressures the fluid behaves very much like a gas. The fluid inside the liquid spring oleo behaves very much like a gas.

A fluid's compressibility is related to its Bulk Modulus - the relationship between its pressure and its reduction in volume.

- al

I

The unitflEonsists of a thick walled cylinder (to withstand the very considerable press&-es that are built up) hoqsing a piston with a large diameter piston rod (figure 2 1).

I

These jv<ry high pressures are diffcult to contain with conventional seals (around the pistdn rod) so a special high pressure gland assembly is used. The idea behind the gl&d assembly is to control the friction between the gland and tlie piston rod whilsti still keeping a fluid tight joint. It achieves this by progressively increasing the seal pressurec'on the rod as the pressure rises.

On lahdini-the piston is caused to enter further into the cylinder reducing its intekGalalvolume by the amount of piston rod that enters. This compresses the fluid and raises its pressure. During compression the valve (flutter plate) is open (figures 2 1 and 22) and fluid passes reasonably freely across the piston head.

On rebound the piston moves out of the cylinder and the valve/flutter plate closes. This reduces the number of holes that the fluid can pass through so the rate of piston movement is reduced.

QUESTION When the piston gets to the end of its travel at its maximum collapse position during landing what are the pressures either side of the piston? (2 mins)

ANSWER They are the same. The pressures are very high but they are the same.

QUESTION If the pressures are the same on both sides of the piston what makes it extend? (5 mins)

ANSWER A force will make it extend and pressure times area produces a force. The area on the rod side of the piston is smaller than the area on the other side (smaller by the amount that is taken u p by the rod) so there is a greater force pushing it out than that trying to push it in.

L -- - , - -- /

Fig. 2 1 LIQUID SPRING OLEO

Maintenance

Variations in design will necessitate different servicing procedures. The following, however, is typical but always, of course, consult the AMM/oleo manufacturer's manual.

During a visual inspection check for security of attachment, damage (particularly to the ram sliding portion) and leaks - from the high pressure gland assembly and the charging valvelbleeder plug. Check for correct deflection.

AIRCRAFT END OF UNlT

PISTON MOVEMENT

PISTON MOVEMENT

TER CYLINDER

HIGH PRESSURE GLAND ASSEMBLY FLUTTERPLATESTOP

FLUTTERPLATE PISTON HEAD

REDUCEDRETURN FLUID FLOW

FLUID PRESSURE INCREASING CONSIDERABLY AND FLUID COMPRESSING

- - WHEEL END OF UNlT

FLUTTER PLATE (CLOSED)

AIRCRAFT LANDING UNIT EXTENDING SLOWLY

Fig. 22 LIQUID SPRING OLEO - OPERATION - -

If excessive defection is found top up the unit as follows:

1. Jack the aircraft to remove the load from the leg.

2. Select the correct charging gun with the correct fluid (eg DTD 585) - - -

and bleed the connecting hose.

3. Connect the flexible charging hose of the universal charging gun (or similar) to the oleo charging valve. Release the oleo pressure slowly.

4. Bleed the oleo my pumping fluid until there is an air free flow from the bleeder plug.

5. Collapse the unit by placing a bottle jack under the landing gear and pumping up. This will expel fluid (to be collected in a suitable container) and ensure free movement of the oleo to the limit of its range. Tighten the plug.

5. Slowly pressurise the oleo, at the same time lowing the bottle jack. The oleo will extend to its fullest extent. Continue charging to the pressure specified in the AMM. {As an example 1700psi (2.7MPa))

6. Disconnect the charging gun. Check for leaks from the oleo charging valve, the sliding portion (gland assembly) and fit dust cap.

No leakage is permitted and if found the unit must be changed, or if allowed the offending part (for example the charging valve) is replaced in-situ (on the aircraft).

In some cases the AMM calls for a 'scragging test'. This involves repeated collapse and extension of the oleo and checking each time that full extension is achieved every time. A duration test may also be called for which involves the unit being kept collapsed for a period of time and then released to check for full extension.

WARNING. Charged oleos are dangerous. Before attempting to carry out any dismantling the pressures must be relieved.

Faults

, Symptoms Probable-Causes Remedy- - - - \ r - i - - \ --, \ - \

Excessive deflection. Lobs, of pressure. Find ldakls - &ti3 Loss of oil. ' and re+charge.

Slow response to deyction but smooth operation.

I ' 1 >

Slow response to deflection 'and

erratic movement.

Too fast a response to deflection.

~ d , n t or damAge1d piston rod. ,,' -- i _- '

Pressure too high. Faulty flutter plate.

!

Check ,reasons ' rectify !and-re-dh&rge.

-- --/'

Replace $nit. i

Correct pressure. Renew unit.

GAS/OIL OLEO WITH SEPARATOR

This unit has two mediums (gas - usually nitrogen and oil - say DTD 585) and each has its own roll to play. In general terms the gas absorbs the energy by compressing and the oil controls the recoil. The oil is at about the same pressure as the gas most of the time but, because it does not work at such high pressures as the liquid spring type unit, the oil does not compress very much - its job is just as a controlling medium.

QUESTION Air will do the job just as well as nitrogen and is significantly cheaper, so why use nitrogen? (5 mins)

- 37 -

rnoodull l A-1 237

ANSWER Nitrogen will not support combustion - air will. It is possible to imagine the compression of the air on landing. I t will be an adiabatic compression so the air will get hot. Should a small quantity of oil be present then there is the same situation in the cylinder of the oleo that occurs in the cylinder of a diesel engine. BANG we couId have trouble. Called dieseling and would not occur if the gas was nitrogen.

Figure 23 shows a typical gas/oil oleo with separator. It consists of an inner sliding cylinder (plunger tube or piston) sliding inside an outer cylinder. The lop of the inner cylinder has a piston fitted with a flutter plate. Inside the inner sliding tube is a floating piston or separator. Above the separator is oil, below is nitrogen.

The drawing shows torque links preventing axial rotation of the inner cylinder within the outer cylinder. There are seals around the floating piston and between the inner and outer cylinders.

T h e ~ e are qil filler plugs in the top of the ,unit (oil head) and,a,gas ,charging valve at thzibofbn of the inner cyliiider; - - -- - - - - - .

I I , : , , ! ! ! , : ,

, ,

MAIN

Ol L STRUCTURE

CY UNDER:

m m n A'

SEPARATOR

SEALS

GAS CHARGING VALVE \

Fig. 23 GAS/OIL OLEO WITH SEPARATOR

Operation (figure 24)

With the weight of the aircraft off the unit the gas pressure forces the separator to the top of its cylinder forcing all the oil into the top of the unit and extending the unit fully. In this condition the oil is a t little or no pressure and the gas is at its normal charge pressure (say 1200psi - 1.9MPa).

On the landing stroke the downward energy collapses the unit forcing the inner cylinder and piston head upwards. This causes the oil to pass through all the ports in the piston head and force the separator down. This causes the gas to compress and the pressure to rise (the pressure rises in both the oil and the gas but the oil volume changes little). The compression of the gas absorbs the shock of landing converting it to pressure energy and hence heat energy.

WITH .UME

Fig. 24 GASIOIL OLEO WITH SEPARATOR - OPERATION

At the end of the compression stroke when all the energy has been converted/ absorbed, the pressure in the unit is greater than that required to support the static weight of the aircraft and the unit will want to extend rapidly. But this must not happen. As the unit extends (recoils) the fluid flows back through the piston and in doing so the flutter plate closes (assisted by a spring). This restricts the flow of fluid to a limited number of holes and permits a comparatively slow rate of oleo extension - thus recoil is controlled.

After landing the static weight of the aircraft is supported by the unit when it has collapsed sufficiently so the internal pressure is enough to just balance the weight of the aircraft acting through that leg.

rnoodull l A-1239

Maintenance

As always consult the AMM first, but in general the maintenance procedures are not too unlike that for the liquid spring oleo. The charging and topping up procedures, however, are different.

Periodically the unit will need topping up with nitrogen, but it is important that the oil quantity is correct before doing this. If there is any doubt the oil quantity must be checked before attempting to charge the gas pressure.

Provided the AMM says so this may be carried out with the aircraft's weight on the wheels. The procedure is as follows:

Ensure that there is no obstruction (jacks, trestles etc) under the aircraft which it could catch when it settles down. Also, if in a hangar, check 'the other side of the aircraft' for obstructions above the structure as this side will go up. Attach an adapter to the nitrogen charging point and completely deflate the unit - slowly - with a look person/s to see that no obstructions are in the way of the aircraft. Remove one of the oil filler plugs (plug 1) and attach a charging gun filled with the correbt fluid. Pressurise the leg with oil (say 1000psi), this ensure that the floating piston is at the bottom of its travel. Release the pressuse. - -

Remove the other filler plug (plug 2) and pump in oil until an air free flow of oil comes out (bleeding). Refit plug 2. Remove the charging gun and refit plug 1. This now leaves the unit full of oil to the correct specification, free of any air, the oil unpressurised and the floating piston at the bottom of its travel. -Connect an adapter (Turner-gauge) to the nitrogen charging valve and charge with nitrogen. The leg will extend and the aircraft will rise. Refer to a pressure/extension graph for the correct amount of extension required (example figure 26). This graph will be in the AMM and possibly on a plate riveted to the landing gear. Check for leaks, wire lock the plugs, record the work done and sign the appropriate documentation.

If the process cannot be done with the aircraft on the wheels then it must be jacked and the following procedure carried out:

1. Jack the aircraft as per the AMM. 2. Place a bottle jack under the unit to be charged and whilst releasing

the gas pressure slowly raise the bottle jack to collapse the unit. Be careful not the raise it too quickly and lift the aircraft off its main jacks.

3. With the leg fully collapsed proceed as for items 3 to 5 above - but remember check continuously that there is no possibility of the aircrdt being raised of the main jacks.

4. Connect an adapter (Turner gauge) to the nitrogen charging valve and slowly charge with nitrogen. The leg will extend so the bottle jack must be lowered - if not, the aircraft will raise off its main jacks.

5. Charge to a specific pressure as stated in the AMM (say 1200psi but check your manual).

6 . Remove the bottle jack. Lower the aircraft off jacks. 7. Check for leaks, wire lock the plugs, record the work done and sign

the appropriate documentation.

Note. In this condition when the unit is fully extended the oil pressure is about zero and the gas pressure is normal charging pressure.

QUESTION Why is it so important when using a bottle jack not to lift the aircraft off its main jacks - after-all, surely it could be put back easily by just lowering the bottle jack and the aircraft would settle back on its main jacking points again?

I -.

I - --- -

ANSWER - There is a very high-chance-thaif the aircraft i i lifted-o&s, (main or I nose) jacking point (adapter) then the lifting jack A d adapt& will

come out of alignme$t. This will put the aircraft iAto a veG unstable and dangerous condition - dygerous for both the +craft and the

, people working u n d q r n e h . So,?* is imperative that the situation does not arise. 1 , i -_ , ,

I \ \\, I I - -- ,' I \ I I

Fault Symptoms, I

-Symptoms -- ---

/ 1

Probable-e'aqses Remedy ,

Excessive deflection Low air pressure. Find leak/ s - rectify and rolling when and re-charge. aircraft turns on the Gas in the oil. Change unit as separator the ground. seals leak.

Insufficient deflection. Gas pressure too high. Check gas pressure and rectify.

Harsh action. Oil quantity too high. Check level and rectify.

Oleo sluggish in action. Adhesion of seals to Usually temporary, should sliding member. rectify itself after taxiing. If

the fault persists change the unit.

Damaged sliding Change unit. portion.

- 41 -

rnoodull l A-1241

Notes. 1. Some seals on some oleos can be changed whilst the unit is still fitted to the aircraft (in-situ) - check the AMM. 2. Some scratches and slight damage to the sliding portion of the unit may be dressed out (blended out) in-situ. The depth and length of the dressing out is limited - refer to the AMM. 3 . On some civil aircraft an oleo is used which is essentially one oleo within another with two floating pistons and two gas chambers and two oil chambers. One set (oil/gas/floating piston) is a low pressure unit and the other is a high pressure unit. The idea is that when the aircraft lands the oleo is so designed so the high pressure set operates to absorb the shock and when the aircraft taxies it moves on the low pressure set, so giving a more comfortable ride for the passengers. Not a common unit - for costs reasons.

GASIOIL OLEO WITHOUT SEPARATOR

This -type of unit is more popular with civil aircraft manufacturers. It is because lL

is simpler-ivith no separator. Without a separator there is a oil/gas interface and the gas must be a t the top with the oil underneath.

I

Fig. 25 GAS OIL OLEO WITHOUT SEPARATOR

Figure 25 shows an example taken from the A310. There should be no need to remember the details (actual pressures etc) but you should remember the principle of operation and the maintenance procedures.

I t consists of one sliding member inside another, with the inner member having a restrictor head. A centre rod or metering tube slides within the restrictor head. On some units the centre rod is tapered with no fluid port in the middle. The taper gves a progressive deceleration of the unit as it collapses as less oil is able to squeeze past the rod.

The centre tube of the unit shown is used to brake the end of the oleo's extension after take-off. Total travel for the shock absorber is 17.7 inches (45mm).

Valves are fitted (main valve and restrictor valve in the drawing) to control recoil. These allow a reasonably fast collapse of the unit on landing but close to control the rate of extension on recoil.

1

With the weight of the aircraft off the leg the gas pressure for& down &, the oil which fully extends the unit. On 'lahding the h i t is caused t o telescope w'hich forces fluid from the bottom chamber to the top. The fluid will: pass through the port in the centre rod and through ,bzhppen restrictor valvesi ~ 0 t h - t h e fluid and gas will increase in pressure witd the gas compressing in volume- this absorbs

! 1 the shock. , , I I I ' I I

I I

On recoil the fluid is pushed by the gas downwards to move into the bottom of the unit. This will cause both restrictor valves to hose and f l u i d ~ a n only pass thro&h-the-small holes in the buter-restri6tor valve and the the centre rod - thus recoil is controlled.

In the static weight condition the internal pressure times the circular internal area of the unit equals the weight acting through that undercarriage leg.

Maintenance

The unit is of the direct nitrogen fluid contact type and is filled with oil and nitrogen through one common filling and charging valve. This is located a t the top of the unit and is connected to a short stack pipe which automatically limits the amount of oil that can be in.

Nitrogen pressures and compression distances are given on a chart attached to the leg and in the AMM (figure 26).

Ramp servicing is limited to visual inspection for damage, leaks and correct extension.

Oil level is checked with the unit collapsed. The aircraft's weight can be on the wheels or the aircraft can be jacked and a bottle jack used.

After the oil level has been topped up to the bottom of the stack pipe the unit is charged with nitrogen.

SERVICING CHART

AMORTISSEUR

i,wITn AIRPLAY WEIGHT OM MAR I,AKIRTISSEUR SOUS

M€ASU(IE STRUT M E S S U R E - CHARGE SlAlIQVE W W E R

w l r n PRESSURE OABE SA PAESUOW

I-DETERNME CDRMCT DtYLMKm 2-EN D E M E L A L O W U E W

'WFOII T I l n WESfCIRt: F I O l U* Sm LE M U R A W

SERVawB C U I V L

3 - F W C L S S A I V ADO DR1 1-9 M C E S Y I M W L E R A

MIM)(;EM TO OITAlM CORMCT L'AZmE EEC hWI

STRUT M w a c

Fig. 26 A310 NOSE GEAR OLEO CNARGING CHART

blank

SHOCK

.--- - ---

SPHERICAL BEARING

i i

, Fig. 27 SHOCK A B S O ~ E R ASSEMBLY I ' ' I 1 !

Charging-Procedure - Aircraft on Ground / - --- I

1. Carry out the same precautions as with the gas/oil oleo with separator.

2. Deflate. Connect a selector valve and pressure hose to the filling and charging valve. Open the valve and release the pressure until the oleo rests against its inner stop. Check extension (dead travel 34mm).

3. Oil Chargng. Connect the oil pump (filled with the correct specification oil - as per AMM/data plate) to the selector and charge the oleo with oil to raise it by 20mm. Check the new dimension is 54mm.

4. Bleed. Wait 1 minute. Open the selector and release the fluid allowing the unit to settle on its inner stop again. Check fluid for bubbles. If necessary repeat steps 3 and 4.

- 45 -

rnoodull l A-1245

' 3 1. DEFLATE

Fig. 28 DEFLATION

, I -

5. Pressurise. Connect a nitrogen pressure source to the selector. Pressurise to 5 bar (72.5psi). Check there is no movement of the oleo.

6. Bleed. Open the selector to bleed the fluid. The 5 bar pressure acts inside the unit against the fluid level to force any surplus fluid up the

; stack pipe and out of the unit. This happens until the fluid level gets to the lower end of bhe stack pipe indicating the correct fluid level.

7. I Nitrogen Charging. Connect the nitrogen pressure source to the , selector and charge slowly to obtain the correct pressure ' corresponding to "D" on the seivicing chart 9 (fi&re 32).

2. OIL CHARGING

Fig. 29 OIL CHARGING

- 46 -

moodull lA-1246

3. BLEED

Fig. 30 BLEEDING

- -- - -

Fig. 3 1 PRESSURISING

Pressure/Extension Check - Aircraft on the Ground (Figure 32)

An example check with ambient temperature at + 10°C is as follows:

1. Connect a pressure gauge to the filling and charging valve. 2. The gauge reads lOObar (1450psi). 3. Checking pressure against dimension "D" on graph and the 10°C

curve gives lOOmm (3.9in). "D" is to be measured between the base of the shock s tmt and the upper surface of the torsion link attachment.

4. Add more nitrogen if necessary.

- 47 -

rnoodull l A-1247

, Fig. 32 OLEO DATA PLATE

SERVICING EQUIPMENT

The AMM will specify the equipmeritpto be used. In some cases it will be "special to type" and supplied by the aircraft manufacturer, in others the equipment will be standard equipment. In the case of standard equipment the aircraft manufacturer will specify the requirements the equipment has to meet - such as threat adapter sizes, pressure requirements, fluid flow rates etc. What follows is a description of some_"standard equipments".

Universal Charging Gun (figure 33)

This hand held gun is used to charge liquid spring and gas/oil oleos with oil. It consists of an adapter block fitted with a Bourdon tube type pressure gauge, a non-return valve, a bleeder screw and a flexible hose - one end of which connects to the equipment to be charged. The other end connects to the gun itself.

The gun is filled with the fluid to be pumped and the banjo union is screwed directly into the charging valve of the oleo to be charged (after removing the dust cap).

Before use fill the gun with the specified oil then proceed as follows:

1. Screw the flexible charging hose onto the adapter block and screw the adapter block onto the charging gun.

- 48 -

moodull l A-1 248

2. Ensure the bleeder screw is screwed down and prime the flexible hose by pumping the handle until an air free flow emerges.

3 . Remove dust caps of banjo union and oleo and screw banjo union to charging valve. Do not tighten fully.

4. Pump fluid to issue from the loose connection and tighten connection a t the same time. This ensures that there is no air in the charging hose.

5. Operate handle to pump fluid into oleo noting the pressure on the gauge. If oleo charging valve not of the non-return type slacken to allow pumping.

6. When charging is complete tighten oleo charging valve (not of the non-return valve type), release pressure in the charging hose by slackening the bleeder screw. Remove hose.

7 . Fit dust caps. Lock wire as appropriate.

PRESSURE GAUGE

' I

~ .-

.

CYLINDER CONTAINING FLUID TO BE PUMPED

SPECIAL BANJO UNION (To connect to equipment to be charged)

Fig. 33 UNIVERSAL CHARGING GUN

Inflation Adapter

Used when inflating components with gas and also used for deflation, pressure checking etc. Is fitted with a control valve to control the rate of inflation and a pressure gauge. Is also fitted with a pressure release valve. It has an adapter on one end to connect to the high pressure hose from the nitrogen charging trolley and an adapter on the other to connect to the oleo.

moodull lA-1249

High Pressure Nitrogen Charging Trolley

May have two wheels and fitted with one high pressure nitrogen bottle. May be a four wheeled trolley with four bottles and provision made for towing. Controls are provided for reducing the bottle pressure and for regulating the flow. Pressure gauges are provided to monitor the bottle pressure (fully charged normally 4500psi) and the reduced supply line pressure. It is important to note that when charging this line pressure gauge reading should correspond to the adapter gauge reading on the component.

The controls on the trolley can be set to reduce the bottle pressure down to the line pressure required for the component.

To charge an oleo the adapter is connected to the oleo charging valve and the hose from the high pressure nitrogen charging trolley is attached to the adapter. The trolley controls are set to OFF. The bottle key is inserted into the end of one nitrogen bottle and it is turned ON. The first gauge on the trolley controls will show bottle pressure (say 4500psi). The trolley regulator valve is adjusted to give the korrect line pressure to the component (say 1200psi) - shown on the second gauge 'on the trolley. When correct the third valve is opened to supply this pressure to the adapter on the oleo. The gauge reading on the adapter should be the same as the second gauge on the trolley.

I I

CONTENTS

Page

Tyres Tyre wear Markings Pressures Inspection Mounting tyres Inflation

s torake Inner thbes Wheels I

Inspection Brakes

Drum brakes Disk brakes

Skid cont!rol , Mechanical &ti skid Electron$ anti skid

-A320-system --B7 67 -

AIRCRAFT TYRES

Until recently these were usually of cross-ply construction. However, radials are becoming more popular as they offer a reduction in weight with increased landings.

The cross-ply tyre is made up of plies of weftless nylon fabric, each ply laid on the bias and at 90' to the previous ply.

The radial tyre is made u p of weftless plies aB laid radially across the tyre one upon the other.

JAR25 states the requirements that a tyre has to meet, in particular the downward and forward loads in terrns of "g" and relate to main wheels, nose wheels and single and multi wheeled units. Tyres must have an approved speed and load-rating and have sufficient-clearance when retractep-to-aUow..for tyre growth: Tyre growth is the incrkase-in-size of the tyre due to-centrifug&forces at

1 : I high speed. I I I

I I / I I I I i I

I I I

1 1

I I Cross ply Construction - - I I

--I\ I --

Tread. Made of natural rubber, cbr$poun&k,d\for toughness y d durability. The tread pattern is designed in accordance with qrcraft operationaf requirements. The cir~urnferentid ribbed tread i s widely used today to provide good traction

I under v*ng Wnway conditions. , /' 1 < -

I - 1

,/ L

Sidekall~~protective layer of flexib1e;wezdher-resistance nhttural-rubber covering The outer carcass ply, extending from tread edge to bead area.

Tread Reinforcement. One or more layers of nylon fabric that strengthen and stabilise the thread area for high-speed operation. Also serves as a reference for the buffing process during the re-treading of tyres.

Breakers. Reinforcing plies of nylon or Ararnid fibre placed under the tread rubber to protect carcass plies and strengthen and stabilise the tread area. They are considered an integral part of the carcass construction.

Plies. Alternate layers of rubber-coated nylon fabric (running at 90" to one another) provide the strength of a tyre. Completely encompassing the tyre body, the carcass plies are wrapped around the HTS wire beads and back against the tyre sidewalls (ply turnups).

Beads. High tensile copper-coated steel wires embedded in rubber, the beads anchor the carcass plies and provide a firm-mounting surface to the wheel.

- 1 -

rnoodull l A-1252

TREAD PATTER

REINFORCING

RUBBER LINER TUBELESS TYRES

- - -- STRIPS

Fig. 1 CROSS SECTION OF A TYRE

Apex Strip. A wedge of rubber affmed to the top of the bead bundle, serving as a filler.

Flippers. These layers of rubberised fabric help anchor the bead wires to the carcass and improve the durability of the tyre.

Chafers. Protective layer of rubber and/or fabric located between the carcass plies and the wheel to prevent chafing.

I i

~ i ~ . a BADWL TYRE \ I . 1 - --_. ,

I \ '\ ,

Bead Toe? The inner bead edge ciosest to the tyre centre line. I i --

; I / ' I I ' Bead Heel. The, outer bead edge $hat fits ag&st the wheel f l ~ g e .

- , , - -' \ -

InnetLi~et~In-tubeless tyres, this-inner-l5ier of low permeability (air proof) rubber acts a s a built-in tube and prevents air from seeping through the casing plies. This liner covers the whole of the inside of the tyre (carcass) from just before the heel area one side to just after the heel area on the other. For tube type tyres a thinner liner is used to prevent tube chafing against the inside ply layer but this is not air proof.

Fitting Lines. Lines moulded around the lower sidewall to check for concentricity of tyre to wheel.

Radial Construction

The basic difference between a cross-ply tyre and a radial tyre is the lay of the plies within the tyre carcass. As stated previously, the plies are laid radially across the tyre, but the remainder of the tyre is very similar in construction to a cross-ply tyre.

rnoodull l A-1254

SIDE

INNER LINER

WIRE :BEADS

Fig. 3 CROSS PLY TYRE - . x

I

~ a d i a l tyres feature a rigid belt Bnd a flexible carcass, providing an increase in the number of landings and a reduction in rolling resistance. The efficient use of high strenhh materials results in a lighter weight tyre with improved performance.

Tread-Patterns

The tread pattern on a tyre is usually designed to suit specific operating conditions (runways, grass strips etc), aircraft weights and aircraft take-off and landing speeds.

Ribbed (ie circumferentially grooved) treaded tyres are probably used more than any other and there are a number of variations on the basic pattern such as the number of ribs and the width of grooves. A ribbed tread provides a good combination of long tread wear (long tyre life) and good traction and directional stability, particularly on hard surfaced runways.

Diamond pattern (or 'all-weather') tyres are also used and give good performance on all types of surfaces. They are particularly suitable for unpaved (eg turf or packed earth) airfields.

- 4 -

rnoodull l A-1 255

The plain tread (smooth tyre) was common at one time, particularly on British aircraft, but has gradually been replaced by ribbed and diamond pattern treads. I t is, however, still used on some older fixed wing aircraft.

Some nose wheel tyres are fitted with a water deflector (or 'chine') on the upper sidewall, to deflect water away from rear-mounted engines when landing/taking off on/from wet runways. This deflector may be on one side for twin-wheel installations or on both sides for single-wheel installations.

Water dispersing treads, which have many small holes incorporated in the crown and shoulder rubber, are also fairly common as a means of helping to prevent aquaplaning.

SINGLE CHINE NRE - I ESTION: Describe aquaplaning. (5 mins)

- ! I

-1

TYRES 1

ANSWER: It is a condition where the tyre is travelling along a wet runway so fast that the water has not got time to be moved out of the way by the tyre tread (normally a sort of 'squeezy' action). This means that a thin film of water gets in between the runway surface and the tyre reducing the grip so much that the effect is similar to being on sheet ice. Aquaplaning is a function of speed, standing water depth, aircraft weight and tyre tread pattern (a smooth tyre aquaplanes at a slower speed than a patterned tyre).

TYRE WEAR

It is important to note that different manufacturers may have different wear indicators and reference must be made to the specific manufacturer's manuals for actual wear limits. What follows are typical examples.

- 5 -

moodull l A-1256

Tyres with Marker Tie Bars

Some rib pattern tyres embody equispaced marker tie bars in the groove designated for use as a wear indicator. The limit of wear is reached when the tie bar is worn to the base of the groove which contains it.

Tyres must be renewed a t the following wear stages:

* Recommended when the tread is worn to the top of the marker tie bar.

* Mandatory when the tie bars are worn to the base of the groove which contains them.

MARKER TIE BAR

-

1

. -

I Fig. 5 MARKER TIE BAR

I

Tyres Without Tread Reinforcement --- -

-

Assessment of tread war on these tyres must be based on using the centre groove as a wear-indicating groove. If the tread pattern does not incorporate a centre groove, use the grooves outside the centre rib. The tyre is worn to its limit when the tread is worn to the base of the wear indicator groove.

WEAR INDICATOR GROOVEIS

Fig. 6 WEAR INDICATOR GROOVES

Tyres with Tread Reinforcement

Identified by the letters DRR in the Code Panel and the words REINFORCED TREAD on the sidewall. They may or may not have marker tie bars. The limit of the wear is reached when the depth of the wear-indicating groove is reduced to 2mm. If the wear is not even, the tyre must be discarded when a total of 25% of the groove circumference is reduced to 2mm depth, irrespective of the depth of the remainder of the groove. Tyres thus worn must be withdrawn from service for possible remoulding.

Ai r Worthiness Notice

AWN number 5 states that wear indicators must be fitted that indicate when wear has occurred that impairs wet braking efficiency. This must be part of the

i certification process for the tyre. \ 7-- .- -. - - - --

In thb absence of a particular wear-indicathr the tyre must be withdra'nm , i from service when the following applies:,

I I

* If any groove has a d e ~ t h of less than 2mm for more than 25F of the , I tyre circumference. , /I I

\ ', I

* If the tread has worn to less than:2mrn across the entire width of the treadjin contact with the r u n ~ a y ~ a t any place on ithk circumference.

/ - - /' -

- \ - I

~ o r n e ~ i a t i o n s - k a y be allowed, eg if-one$roove is less than- 2mm and the others are 3mm or more.

Twin Contact Tyres

These have two points of contact with the runway either side of the crown. Not in common use and were designed to reduce the problem of shimmy. Fitted to nose wheels only.

Can remain in service until the centre of the crown shows signs of having been in contact with the ground. (The centre of the tyre between the contacts can be as high as 25mm from the ground on a new tyre).

Helicopter Tyres

Are fit for further service when they contain tread cuts, which expose but do not penetrate the casing plies, irrespective of the number or size of cuts.

moodull l A-1258

Chinned Tyres

Inspect for cracking along the chine/sidewall junction. Cracking up to 0.4mm (0.0 16 inches) in depth is permitted irrespective of length. Remove a chine tyre having:

(a) A crack deeper than 0.4mm and exceeding 25.4mm (1.0 inch) in length.

(b) A crack deeper than 2.29mm (0.09 inch) at any point.

1 I

LOAD RATING - -

TUBLESS APPICABLE SPECIFICATION

APPLICABLE

SKID 8 SPEED RATING

PART NUMBER &

CODE LOAD RATING

TREAD CODE

Fig. 7 TYRE MARKINGS - GOODYEAR

- 8 -

rnoodull l A-1 259

Tyre Markings

All commercial aircraft tyres approved under FAA Test Requirement TSO-C62c (AWN 93 refers) are marked clearly with the following minimum information:

Manufacturer's name Size Load Rating Speed Rating. (Air Worthiness Notice 93 states high speed tyres as rated at above 160mph.) Skid Depth Manufacturer's Part Number Serial Number Manufacturer plant identification along with the TSO marking.

r In addition, Goodyear tyres are marked with the ply rating and other marhngs as t requiredb? I airframe manufacturers-or other -.. organisations. i __ - ,

-. -\ ' -

Military tyres carry markings required by the, appropriate militaky s6Scification. I 1

In addition, the Association of European Ajrljhes (AEA) requires an AEA code which defines both tyre carcass h d tread cpnstruction. 1 1

; I

/ --- - \.. --

All retread t v e s carry the symbql 'b' or tf$e,dprd 'RETREAD'. i It also ifidicates the number of times $he. tyre has been retreaded :(Some tyres are\ rk-treaded many times).

Tyre sixes are usually identified by three dimensions, eg 26 x 7.55 - 13. The markings are in inches but the inch mark is not shown. If the size is in millimetres then the symbol mm is used.

26 - equals overall diameter of inflated tyre. 7.55 - equals the cross-sectional width of the inflated tyre. 13 - equals the bead diameter of the bead seat on the wheel.

NOTE. When only two figures are given it is the first figure, the overall diameter of the tyre that is omitted.

Serial Number

Moulded into the tyre wall to identify a particular tyre. Each tyre will have its own unique serial number to identify it against its records.

SERIAL NUM

EQUIPMENT IDENTIFICATION

TYPE NUMBE

DGAC STANDARD

FAA STANDARD

MOLDED SKID

PART NUMBER SPEED

AEA CODES

SERIAL NUMBER TYRE SIZE

I

RATING

Fig. 8 TYRE MARKINGS - MICHELIN

Date of Manufacturer

Moulded into the tyre wall.

Tyre Pressures

Range from 25 to 350psi gauge pressure. Always consult the AMM. Pressures may be given in psi, bar or kPa. Tyres with greater than 100psi pressure are considered to be high-pressure tyres.

Notes

1. Gauge pressure is the pressure that is indicated on a gauge. For absolute pressure the ambient atmospheric pressure must be added. So at sea level with ICAO standard conditions add 14.7psi or 1.03bar to the gauge reading to obtain the absolute pressure. In general all tyre pressures are given in gauge so normally the problem does not exist - this note is included as the CAA are known to ask questions re absolute and gauge pressures.

2. 14 .5ps i= lba r lpsi = 6894Pa

3. Pa = Pascal

Ply Rating --- -- -- -- - - \

- - \ I -

The term ply rating is used to identify a Greys maximum loadmarid prk.ssure. It is an index and does not represent the actual number of cord plies in th&, constnicdion. 1

I 1 I

Conductivity I

I -, Some tyres are qianufactured wilh; tread rubber with conduc$ng compounds to

I permit ea'rthing of static charges!. Wsually nose and tail w h e e l ~ ~ e s are conducting and have markings to indicate this such as a lightening strik,e symbol or the word "conducting".- - - -- - -

IN-SITU TYRE INSPECTION

Regular inspection of tyres is recommended for safety and tyre economy. The frequency of the inspection should be determined by the use and normal tyre wear of the particular aircraft involved and is specified in the maintenance schedule. With most aircraft, tyre inspection after every landing, or at every turnaround, is required. With all aircraft, a thorough inspection (usually in the bay) is advisable after a heavy or overweight landing.

Tread Wear

Tyres should be removed when tread has worn to the base of any groove at any spot, or to a minimum depth as specified in the AMM or tyre manufacturer's manual. Tyres worn to fabric in the tread area should be removed regardless of the tread remaining.

moodull l A-1 262

Uneven Wear

If tread wear is excessive on one side, the tyre can be removed from the wheel and turned around, providing there is no exposed fabric. Gear misalignment causing this condition should be corrected.

Cuts

Inspect treads and sidewalls for cuts and remove if:

1. There are any cuts into the fabric. 2. Cuts extending across more than 50% of the width of the rib.

To check for depth of cuts the cut must be probed by a tool something like an ice pick. This requires considerable sideways force to prize the rubber to one side to see the bottom of the cut. Once the bottom of the cut can be seen the plies can be checked for damaged and the area can be checked for foreign objects.

I

WARNING. Deflate the tyre before this procedure is carried out.

Bulges -- .

I

Bulges in any part of tyre tread; sidewall or bead areas, indicate a separation of the plies or damaged tyre. Mark the area with a crayon and remove the tyre.

1

Fabric Fraying/ Groove Cracking

Tyres should be removed from service if groove cracking exposes fabric or if crachng undercuts the tread ribs.

Flat spots

Generally speaking, tyres need not be removed because of flat spots due to side or hydroplane burns unless fabric is exposed. If excessive unbalance results, however, remove the tyre.

Beads

Inspect bead areas next to the wheel flanges for damage due to excessive heat - especially if brake drag or severe braking has been reported during taxi, take-off or landing.

INFLATION PRESSURES

Keeping aircraft tyres a t their correct inflation pressures is the most important factor in any maintenance program. The problems caused by under inflation can be particularly severe. Under-inflation produces uneven tread wear and shortens tyre life due to of excessive flex heating. Over inflation can cause uneven tread wear (in the centre of the crown), reduced traction, makes the tread more susceptible to cutting and increases stress on the aircraft wheels.

Tyre pressures should be checked with an accurate gauge on a daily basis. Ideally, pressures should be checked before each flight. Check only cool tyres - at least 2 or 3 hours after a flight. Use an accurate gauge, preferably the dial type. Inaccurate gauges are a major source of improper inflation pressures. Gauges should be checked periodically and re-calibrated as necessary. Cross check gauge reading with flight deck indicators - if fitted.

The inflation pressure as stated in-the AMM should be used for each-tyre. I t must be determined if loaded' or 'unloaded~infl&ion pressure hak+beenspe?ified.

I 1 ' , I

I ' When 4 t$re i s under load, the q r chamber volume is reduced due to tyr; deflection. Therefore if unloaded pressure has been specified; that pre&sure should be increased by 4% to obtain the - ehuivalent loaded inflation presiure.

~d jus t ind for Temperature

When tyres a5,subjected to ground tempera<ure changes in exdess of 80°F (27°C) because of flights to a different-climates;-fnflation pre~sures~should be adjusted for thkworst-c-xse prior to take=off.------' --

The minimum required inflation must be maintained at the cooler climate and the pressure can be adjusted in the wanner climate. An allowance must be made for the inflation drop in the cooler climate. An ambient temperature change of 5°F or 3°C produces approximately a 1% tyre pressure charge.

NOTE. Excess inflation pressure should never be bled off from hot tyres. All adjustments to inflation pressure should be performed on tyres cooled to ambient temperature.

- 13 -

rnoodull l A-1264

Cold Pressure Setting

The following recommendations apply to cold inflation pressure settings:

1. Minimum pressure for safe aircraft operation is the cold inflation pressure necessary to support the operational loads as determined by the formula under 'Unloaded Inflation' or as specified in the AMM.

2. The loaded inflation must be specified 4% higher than the unloaded inflation.

3. A tolerance of minus zero to plus 5% of the minimum pressure is the recommended operating range.

4. If tyre-in-service pressure is checked and found to be less than the minimum pressure, the following table should be consulted (of as

- specified in the AMM/tyre manufacturer's manual). In service is defined as a n aircraft taxing, taking off or landing but does notinclude hangared aircraft.

I

~ y & Pressure Recommended Action ' I

100 tb 90% of service pressure Re-inflate to specific service pressure. ..

89 'to:80% of service pressure Remove tyre from aircr&t.

1 '

79% or less Remove tyre and axle mate from aircraft. - -

Blown fuse plug -Scrap tyre. If blown while in servicing (rolling). Scrap axle mate also.

NOTE. Any tyre removed because of low inflation pressure should be inspected by an authorised re-treader to verify that the carcass has not sustained internal degradation. If it has, the tyre should be scrapped.

MOUNTED TUBE TYPE TYRES

A tube-type tyre that has been freshly mounted and installed should be closely monitored during the first week of operation, ideally before every takeoff. Air trapped between the tyre and the tube at the time of mounting could seep out under the beads, through sidewall vents or around the valve stem, resulting in an under-inflated assembly.

- 14 -

rnoodull l A-1265

MOUNTED TUBELESS TYRES

A slight amount of diffusion through the carcass in tubeless tyres is normal. The sidewalls are purposely vented in the lower sidewall area to bleed off trapped air, preventing separation or blisters. A tyre can lose as much as 5% of the initial inflation pressure in a 24-hour period and still be considered normal.

NYLON STRETCH

The initial stretch or growth of a new nylon tyre results in a pressure drop after mounting. Consequently, nylon tyres should not be placed in service until they have been inflated a minimum of 12 hours, the pressures re-checked and tyres re-inflated if necessary.

NYLON FLAT SPOTTING - --- - - . 1 -

"1 \

-- - -\ - - Nylon tyres on aircraft left statiorary for &$length of time d l 1 develop temporary flat spots, The degree of this flat-spotting dep~nds on the load, tyre defledtion and ternperathe. Flat-spotting is more 'severe and more difficult t& work odt during cold wesather. Moving a station+ @craft regularly will lessen this condition. If possible, an aircraft parked for longj%Gods (30 days or more)/ should be jacked up to remove the weight from the tyre;under normal conditions,-a fldt spot will

1

disappem by the end of the taxi h n . i

I I '

1 ' 1 I

I I 1 I

COLD WEATHER PRECAUTIONS -- - I

/ - / .

-- -/ - -

Much of the following also applies to tyre maintenance in general.

1. Use new O-ring seals with best cold weather properties, properly lubricated and installed.

2. Use an accurate dial type pressure gauge.

3. Be sure that wheel bolts are properly torqued as per wheel manufacturer's instructions.

4. Aircraft parked and exposed to cold soak for a period of time (1 hour or more), should have tyre pressure checked and adjusted accordingly. Tyres will have taken a nylon 'set' and experienced a pressure drop.

5. High-speed taxis and sharp turns should be avoided to prevent excessive side loading.

6. Remember that for every 5°F (3°C) change in temperature there will be a corresponding 1% change in tyre pressure.

Tyres subjected to above normal braking energies during an RTO (Rejected Take- off) should be removed and scrapped. Even though visual inspection may show no apparent damage, tyres may have sustained internal structural damage that could result in premature failure. Also, all wheels must be checked in accordance with the applicable Wheel Overhaul or AMM after an RTO.

When new and/or re-treaded tyres are installed on the same landing gear axle, the diameters should be matched within the Tyre and Rim Association inflated dimensional tolerances for new and grown tyres.

It is recommended that tyres mounted on dual wheels have similar inflated outside diameters to ensure that each tyre will carry an equal share of the load. The outside diameter of tyres (new or retread) should be measured at operating pressure. . ..

I - -- ..

Tyres 'should be kept clean and .free of contaminants such as oil, brake fluid, grease, far and degreasing agents which have a deteriorating effect on rubber. Contaminants should be wiped off with alcohol, then tyres washed with soap and water immediately. When aircraft are being serviced, tyres should be covered.

1 -- --

Aircraft kyres, like other rubber products, are affected to some degree by sunlight and egremes of weather. While weather-checking does not impair performance, it can be reduced by protective covers. These covers (ideally with light colour or aluminised surface to reflect sunlight) should be placed over the tyres when an aircraft is-tied down outside.

- ---

Regardless of the excellence of any preventive maintenance program, or the care taken by the pilot and ground crew in handling the aircraft, tyre damage will result if runways, taxi strips, ramps, hangars and other paved areas of an airfield are in a poor condition, have debris or are improperly maintained.

These areas should be kept clean of stones, tools, bolts, rivets and other foreign materials at all times. With care and caution in the hangars and around the airfield, tyre damage can be minimised and ingestion damage to engnes can be reduced.

MOUNTING PROCEDURES

Deflate all tyres before wheel disassembly. A clip-on chuck, an extension hose and a safety cage are recommended for inflation. A direct reading or dial tyre pressure gauge should be used. Use of excessive air pressure to seat beads can cause failure and result in serious injury.

TUBELESS TYRES

1. Check for word 'Tubeless' on sidewall. 2. Make sure tyre is clean inside. 3 . Inspect the wheel halves (split wheel, divided wheel, loose flange etc)

for damage and corrosion. Check all seals for security and damage Inspect all fittings for security - inflation valve, drive blocks, fusible plugs, bearings etc. Make sure that the "halves" of the wheel are the correct halves and fit correctly oriented.

4. Inspect the tyre. Check that it is undamaged and within the wear limits if a used tyre. Check that it is the correct tyre and is the correct size.

5. Clean the bead base with a cloth darnpened with denatured alcohol. 6. Align red balance dot (light spot) on tyre with wheel valve or wheel

heavy point, if indicated on the wheel. Lower the tyre onto the first halve of the wheel.

7. - Ensure that tyre fits first halve of wheel correctly andglace second r

halve on the top of-ther first-->nsuring they mate, colrzcsyy 8. Be sure that wheel bolts are\pro$erly torqued as per the'iaheel

manufacturers i n s t r p d t i o n s / ~ ~ ~ . I I I

9. Inflate tyre to rated bressure using inflation gauge and s&ety cage. 10. Check for any obvious! signs of,&- leakage, partic)du-ly from the

valve. I - -- -\ --

11. After a 12-hour strejch perio'd; &-inflate to rated Fnflation/bressure. I ' ',

I ' I I I I

1 1

Air Retention Check I / i /' , I

-- .. - -- / - 1 - If pressure-has-dropped more chan-5%-in-tfie next 24 hours:-

1. Check for loose or defective valve, valve core or seal. 2. If OK, release pressure and disassemble tyrelwheel assembly. 3. Check wheel '0' ring seal for condition, proper size and type and

lubricant. 4. Check wheel for cracks, porosity, fuse plug or pressure release plug

leakage.

TUBE TYPE

1. Check the mating tyre and tube are specified and correct for the wheel-tyre assembly (size etc).

2. Inspect the wheel and tyre as for tubeless tyres. Inspect the tube also. 3. Clean inside of tyre and lubricate lightly with talcum powder. 4. Place tube in tyre and inflate to slightly round out the tube in the

tyre .

- 17-

rnoodull l A-1268

5. Align the yellow stripe on the tube (heavy spot) with red balance dot on the tyre (light spot). Assume the tube valve is the heavy spot if no stripe on the tube.

6 . Move the hand around the inside of the tyre between tyre and tube to ensure there are no creases in the tube.

7. Lower tyre and tube assembly onto first half of wheel, ensuring valve stem lowers into wheel valve slot correctly.

8. Lower second half of wheel onto first half. 9. Fit wheel bolts and nuts and torque to wheel manufacturers

instructions/AMM before inflating. 10. Inflate tyre to rated pressure using safety cage. 11. Deflate to equalise stretch in the tube. 12. Re-inflate to rated pressure. 13. Check for obvious signs of leakage particularly the inflation valve.

Air Retention check - _--1

If has dropped more than 5% in the next 24 hours:

'1. Check valve core for leakage. 2. If OK, release the pressure, disassemble tyreltube from wheel and I check tube for leaks. 13.1 Replace tube as required. I

I

ASSEMBLY OF TUBELESS TYRES AND TYRE/TUBE COMBINATIONS

- - - -

Wheel Sealing Ring Examination

Ensure that the sealing ring is free from deformation, permanent set, ageing and general damage; lightly grease the sealing ring with silicone grease.

Assembling Tyre to Wheel

Stretch the sealing ring evenly on to the wheel and ensure that it seats correctly in its groove. It is imperative, however, to ensure that the tyre beads do not become contaminated with the grease. The clearance between thetyre beads and the fitted wheel-sealing ring is adequate to avoid contamination if care is exercised in fitting.

It is recommended that the tyre is fitted on the wheel with the 'red spot' of the tyre lined up with the valve location. (Early issue wheels may have a heavy spot indicated by two concentric rings; such markings should be ignored). Ensure that the sealing ring has not been disturbed and assemble the wheel.

moodull l A-1269

(b) Connect a suitable nitrogen* pressure supply to the valve and inflate rapidly to spread the tyre walls against the flanges. Stop pressurising once this condition is achieved. If a seal cannot be affected, stand the assembly vertically and apply a circumferential load around the crown.

* Tyres inflated with air and subject to excessive heating due to braking action can experience a chemical reaction releasing volatile gases. In the presence of oxygen this can cause an explosion. AN70 (data now transferred to CAP747) requires that tyres on aircraft over 5700kg with retractable landing gear should be inflated with a suitable inert gas with oxygen levels no greater than 5% by volume. Nitrogen is widely used.

(c) Disconnect the pressure line, allow to deflate and fit the appropriate valve core using the relevant torque spanner.

(d) For assemblies requiring inflation up to 175psi (12.06 bars) inflate to the - --

r_e'quired pressure as $lEly asspracticable and fit the-ve cap. -- \ 7 --, ,

I

(e) 'For assemblies requiridg pressures greater than 1 7sp;i, inflate 'the tyre to a pressure of 140~si (10 .33 bars) then deflate completely. check that the tyre beads are properly seated on the wheel flanbe8. If not, re-inflate to 150psi and deflate completel~.~~inally, if the beads I__ are - corrkctly seated, inflate the tyre slo%fly-t~to-wdqking pressure arkd fit th_e,valve cap.

I -

1 ' I ', (f-) 'After a period of time, the pressure may decrease dde to tyre stretch, and

'up to 10% can be regarped as normal.

-- -- --.-

Tyre and Tube Combinations

With tyre and tube combinations, tread separation and blistering can be caused by incorrect inflation procedure. This is due to air being trapped between the tube and the tyre. This condition is normally caused by inflating the tube too quickly. Any trapped air will find its way through the tyre casing and finally build up between the outer casing and the tread or sidewall. The resultant blister is easily detected on the sidewall but not on the tread. A leaking tube can cause the same kind of defect. If a blister does appear a s a result of trapped air, it will usually occur during the first 48 hours after inflation. If the blister is located in the lower half of the sidewall and does not exceed l.Oin (25.4mm) in diameter, relieve the blister by puncturing the rubber only with a sharp instrument (called an awl) held parallel to the sidewall. If it is greater than 1 .Oin or is located in the upper half of the tyre, the tyre must be removed. Small ridges are moulded on all high-pressure tubes to facilitate the escape of the trapped air and tubes without such vent ridges must never be fitted in high-pressure tyres.

moodull l A-1270

(g After a period of time, the pressure may decrease due to tyre stretch, and up to 10% can be regarded as normal.

Tyre and Tube Combinations

With tyre and tube combinations, tread separation and blistering can be caused by incorrect inflation procedure. This is due to air being trapped between the tube and the tyre. This condition is normally caused by inflating the tube too quickly. Any trapped air will find its way through the tyre casing and finally build u p between the outer casing and the tread or sidewall. The resultant blister is easily detected on the sidewall but not on the tread. A leaking tube can cause the same kind of defect. If a blister does appear as a result of trapped air, it will usually occur during the first 48 hours after inflation. If the blister is located in the lower half of the sidewall and does not exceed l.Oin (25.4mm) in diameter, relieve the blister by puncturing the rubber only with a sharp instrument (called an awl) held parallel to the sidewall. If it is greater than 1 .Oin or is located in the upper half of the tyre, the tyre must be removed. Small ridges are moulded on all high-pressure tubes to facilitate the escape of the trapped air and tubes without such vent ridges must never be fitted in high-pressure tyres.

Types,up to 175psi -

1 Inflate tyre slowly to position beads fully on flanges. -

2. ' Fully deflate to relieve local stretching and creasing of tube.

3. Check that the tube inflation valve is lined up, not under stress and -- seated correctly.

After servicing operations that involve tyre deflation, wheel assemblies must be tested iaw the relevant instructions. These tests include a duration pressure test, when the assembly is allowed to stand inflated for a period of 12 hours during which period checks are made on the pressure. An immersion test (in water) can be carried out when time does not permit the duration test. (The bearings must not be submerged).

(Tyre pressures are given for cold tyres based on ambient conditions of 20°C [63"F]. Should the ambient temperature rise to 23°C (73°F) it will increase the pressure by about I%.)

CREEP MARKS

The purpose of creep marks are to indicate circumferential movement of the tyre relative to the wheel. After the initial movement of tyre to wheel caused by the landing (dependent on tyre load, inflation pressure, degree of braking, shimmy on noseltail wheels, speed of landing etc) any subsequent movement of the tyre around the wheel takes place at a much reduced rate, which if allowed to continue beyond a certain limit, can, on a tyre and tube combination, result in the valve stem being tom from the tube - and total deflation. Hence creep marks are painted in white on the sidewall of a newly fitted tyre and over on to the wheel flange. The width of the mark represents the maximum permissible tyre creep and when the paint marks on the tyre and the wheel flange become misaligned by the full width of the mark, the wheel must be removed from the aircraft for bay servicing.

-- - -- -1

Width of Creep Marks -

\

, 7-1 ', i '\

Wheels Outside Diameter (inlmm) ' Width of Tyre hark ( i n / ~ m ) I

I I 1 I Up to 24 (609.6) I

' ,/ 1.b (25.4,) I Over24(609.6) 1 - --_ 1.5-(38A) / Tubeless I '\"\ 3 inches '

I 1 \ -. I ' I

I I I , I

CARCASS VENTS ' __,' ,' , I %

/

-2 - -- , i

All tubeless-tyres, 8-ply rating and above; fiave been ventedLthe' lower sidewall area. These vents prevent separation by relieving pressure build-up in the

. carcass plies and under the sidewall rubber. These vent holes (marked by green coloured dots) will not cause undue air loss. Covering them with water or a soap solution may show an intermittent bubbling, which is normal.

Sometimes called awl holes.

AIR RENTENTION TEST

When no leaks can be found on the prior checks, an air retention test must be performed. The tyre should be inflated to operating pressure for a t least 12 hours before starting the test.

This allows sufficient time for the casing to stretch, but can result in apparent air loss. The tyre must be re-inflated after the stretch period to operating pressure. Allow the tyre to stand a t constant temperature for a 24 hours period and recheck the pressure. A small mount of diffusion is considered normal. However, an inflation pressure drop of more than 5% of operating pressure indicates excessive vent leaking.

Since there are only two reasons for air loss in a tube-tyre, a hole in the tube or a defective valve or valve core, finding an air leak is usually simple. As with a tubeless tyre, the first step is to check the valve and replace the core if it is defective. If the valve is air tight, demount the tyre from the wheel, remove the tube, locate the leak (by immersion in water if necessary) and repair or replace the tube.

When inspecting a tube to decide whether or not it is the cause of the leak use only enough pressure to round out the tube. Excessive inflation strains splices and-may cause fabric separation on reinforced tubes.

I - --

TYRE AND TUBE STORAGE 1

whenever possible, tyres should be stored vertically on tyre racks, kept in their original wrappings, if wrapped. The surface of the tyre against which the weight of the tyre rests should be flat and, if possible, 3 to 4 inches wide to minimise distortion. Stacking of tyres is permissible. However, care must be used to p e v e i t distortion of the tyres on the bottom of the stack. The maximum recommended stacking height is:

- T v e Diameter Maximum Recommended Stacking Height

Up to 40 inches Over 40 inches u p to 49 inches Over 49 inches

Tubes should be stored in their original cartons whenever possible. If stored without their cartons, in bins or on shelves, they should be dusted with talcum powder and wrapped in heavy paper.

Tubes can also be stored in matching tyres. Tyres should be clean and dusted with talcum powder with tubes inflated just enough to round them out.

Under no circumstances should tubes be hung over fittings, pegs or over any object that mighty form a crease in the tube. Such a crease will eventually produce a crack in the rubber.

Ideally, both new and re-treaded tyres should be stored in a cool, dry place out of direct sunlight. Temperatures should be between 32°F (0°C) and 80°F (27°C).

Particular care should be taken to store tyres away from fluorescent lights, electric motors, battery charges, electric welding equipment, electric generators and similar equipment. They create ozone, which has a deteriorating effect on rubber.

Tyres and tubes should be stored on the "first in first out" basis and records should be kept detailing each item's part number, serial number, wheel fitted to, tube fitted, aircraft fitted to, number of landings going on the aircraft, number of landings coming off the aircraft etc. Records should cross refer to other documents such as aircraft log books, J A R form 1 (to be known as EASA form 1) etc.

Care should be taken that tyres do not come in contact with oil, gasoline, jet fuel, hydraulic fluids or similar hydrocarbons. Rubber is attacked by these in varying degrees. Be particularly careful not to stand or lay'tyres on floors that are covered

i with these contaminants.

--- -- I-- -

All tyres and tubes should be inspected i&ediately upon ieceipt-fo;>hipping \ i and hapdling damage and correct referencing to JAR/EASA f o e 1. ,

1

I I

Storage pressures for assembled wheels and tyres are given bb the tyrd manufacturer. Sometimes quoted , as - 20 -- to 3bpsi and Michelin State 25% of normal pressure. , '\, I -A ,'

I \ \ I I - I'

1

I , INNER TUBES ; I I I I

- -- \-_-, , - -

The manufacturer of an inner tube -is done by an extruding .process which forces a compound of hot rubber through a circular die, thus producing a continuous length of rubber tubing. The requisite length, according to the size of the tube required, is cut off, the ends are then butt "welded" together and a valve is fitted. The tube is placed in a mould, inflated and vulcanised, so producing a finished tube to the required dimensions with all the markings on.

During braking, excessive heat is generated in some types of brake unit, which could cause damage to the standard tube. Thus some tubes are manufactured with a thickened base or reinforced base.

Size

The size indicated on the tube is the same size as indicated on the tyre.

Balance Mark

A red line or red dots indicates the heavy spot of the tube. Otherwise assume the valve is the heavy spot.

Examination of Tubes

Tubes and tube base supports may be cleaned with a warm soapy wash prior to examination and dried thoroughly. Slight creasing or surface cracking may be disregarded. On fabric-based tubes, splits in the reinforcement area or minor blisters in the rubber covering around the base of the valve stem, and small pinholes may be ignored. Tube compound properties tend to deteriorate with age and this is often indicated by hardening of the rubber.

A method used in examination of tubes requires the marking of two parallel lines, one-inch-apart, anywhere in the crown area of the tube and stretching the tube until-the lines are approximately two inches apart. The stretched surface of the rubber must then be examined.

If theie are no signs of cracking on the surface of the rubber and if, when released, the tube assumes its original condition without evidence of permanent set in tHe stretched area, then the tube can be considered serviceable.

I

NOTE'. Tubes which have been subjected to excessive heat are scrapped.

STANDARD THICKENED REINFORCED BASE BASE

Fig. 9 TYPES OF INNER TUBE

Tubeless Tyres - Advantages

* Weight saving (7%). * 10% cooler running. * Less deflation risk. * Less normal air seepage. * Less maintenance time. * Valve damage due to creep eliminated

AIRCRAFT WHEELS

AWN 93 states that new wheels brought into service for transport category aircraft as from December 31 1980 must comply with design standard TSO C26c.

While there are many different designs of wheel and sizes and weights vary, the classification of wheels falls into four main headings:

1. Well Based: The well-based wheel is in one piece and is recognised by the deep well in the centre of the hub.

2. Divided Wheel: The divided wheel comprises of either two half hubs bolted together (two piece divided wheel) or a centrepiece with a half hub bolted either side (three piece divided wheel).

3. Loose Flange: The loose flange wheel has a single flange retained by a _- lockring, either round, wedge shaped or a three-piece flange with the

' pieces bolted together. - \ -, - ', '. \ I \

4. ' Detachable Flange: he detac$able flange is bolteid on to the wheel hub. 1 I I ,

I 1

~i rc ra f t wheels are made from alumiKtum &loy or magnesium alloy. Cast wheels, after mladhining, are subjected to impre&ation with a bakelite solution, to prevent air perkgating through the poro+sl wheel. @oSt wheels these dabs are made from aluminium alloy; 1 1 I I 1

/ ,' I

Brakg drive blocks are either mqdel.integrd yith the wheel or,htied on after casting. The centre of the wheel-is-machined to take the whgel-bd&ngs.

Well Base Wheel

Similar to those fitted to most cars, bicycles and motorbikes. The centre of the bead seat area is recessed to allow fitment and removal of the tyre. When easing the bead of the tyre over the flange of the wheel (say at the 12 o'clock position) the bead at the 6 o'clock position is positioned in the wheel well.

As anyone who has ever changed a tyre will know - it is not easy. It is difficult to change on a bicycle let alone a car or an aeroplane. The bead coils are made of high tensile steel and wound around many times, so they will not stretch and forcing the bead over the wheel flange is a difficult task. This is why many wheels have provision to "remove the flange" and bolt it back on once the tyre has been fitted.

moodull l A-1 276

BRAKE DRUM BI-METAL STEEL

BALL BEAR1

WHEEL

LOCATING COLLAR

COUNTERSUNK

- - -

I---

Fig. 10 WELL BASED WHEEL

Divided wheel

~ o n s i ~ t d of two half wheels rnatcied up and connected by bolts which pass through the two halves. The bolts are fitted with self-locking nuts. There is a seal between the two halves on tubeless wheels to prevent air leakage between the halves. ' - _

- - - - -

The wheel may not be exactly in two halves and one side might be bigger than the other.

Loose and Detachable Flange Wheels

Wheels of this type are made with one flange integral with the wheel body and the other a loose flange and machined to fit over the wheel rim on the other side. The difference between the loose and detachable flange type is the method by which the removable flange is secured. The loose flange is retained by a lockng device (similar to a large circlip) on the wheel rim and the detachable flange is secured to the wheel body by studs/bolts and nuts. A loose flange may be a single piece, or consist of two or three pieces bolted together.

The loose flange is secured by a lockring (circlip), which is fitted in a groove in the wheel rim. When assembling the wheel, the loose flange is fitted over the rim and pressed down so that the lockring can be engaged in the rim in the groove.

rnoodull lA-1277

NYLOC NUT SEAL '. \ BOLT

OIL SEAL HOUSING

LOCATING

ROLLER BEARING

--.,

.OUTER--, I D ~ I V E '.

jBL0CKS i 1 I

i

Fig. 11 BIVIDED TYPE WHEEL - -, l - - .

\ i , I

- \

\ ' I

When the tyre is iGflated, the loose flange 4oies outwards an& covers the lockring, thus trapping it in the groove. This prevents disengagement of the ring, which- is locked to the flange and wheelhim by the side of the tyre. For addiiionalse~urity some lockrings-are-joined at the ends by!a lo& plate and screws after fitting.

WARNING. There have been serious injuries with incorrectly fitted loose flanges coming off when the tyre is inflated. Always inflate a newly assembled wheel and tyre in a cage and always double check the fitting of the loose flange, lockring and any additional locking devices during assembly.

To prevent rotation of the single loose flange on the wheel rim, a locating collar on the lockring engages in a recess in both the wheel rim and the flange. To prevent possible disengagement of the lockring during tyre deflation a multi-piece flange was designed.

The multi-piece flange may consist of two or three pieces, each bossed to accommodate securing bolts. Each piece has a beading of semi-circular section on its inner face. This engages in a corresponding groove in the wheel rim.

- 27 -

rnoodull l A-1278

Tubeless Wheels

Where these have loose or detachable flanges suitable "0" ring seals are fitted between flange and wheel to prevent air leakage. The wheel is also fitted with a fusible alloy plug, an inflation valve, no knurled flange (this would reduce the sealing capability between tyre and wheel flange) and a sealed bead seat region to prevent air leakage.

FUSIBLE SEALED WHEEL LOOSE ALLOY PLUG AREA SEAL FLANGE

Fig. 12 LOOSE FLANGE WHEEL

Fusible Alloy Plugs (Tubeless Wheels)

Under hard braking conditions sufficient heat may be generated to weaken the tyre structure causing the tyre structure to fail and blow-out, particularly when it is "baking" in the stowed position within the wheel well.

To prevent this most tubeless wheels have one or more fusible alloy plugs. These plugs screw into the bead seat area of the wheel and are drilled to accommodate a piston seal and fusible alloy plug. When excessive temperatures are generated the alloy melts which allows the air pressure to blow the piston seal out. The tyre will deflate safely.

Creep

In service the tyre has a tendency to rotate (creep) around the wheel. This creep, if excessive, will cause damage to the tyre bead region and tear out the inflation valve causing tyre deflation.

Creep can be kept to a minimum by correct inflation, but design methods can include the following:

* Knurled flange. The inner face of the wheel flange is milled so that the side pressure on the tyre locks the bead to the flange. Not on tubeless wheels.

* Tapered bead seat. The wheel rim is tapered to increase its diameter towards the flange. As the tyre is inflated and the bead moves outwards towards the flange, so it is caused to form a tight fit with the wheel. --

- - -, \ \ i

I / \ \

'0' RING SEAL FUSIBLE

Fig. 13 FUSEABLE ALLOY PLUG

Protective Treatments

All wheels are given a primary anti corrosive treatment followed by a paint treatment. Other treatments are applied but the most popular primary treatments are:

Magnesium ,dloy wheels - chromate treatment Aluminium alloy wheels - anodic treatment

- 29 -

rnoodull lA-1280

INSPECTION OF WHEELS

Check wheels for damage. Wheels that are cracked or damaged should be taken out of service for repair or replacement in accordance with the manufacturer's instructions.

Valves

Before deflating and removing the tyre from the wheel, check the valve. Put a drop of water on the end of the valve and watch for bubbles indicating escaping gas. Tighten valve core if loose. Replace valve core if defective and repeat leak test. Check the valve stem and its mounting for leaks with a soap solution. If a leak is detected, the wheel must be dismantled and a new valve stem installed.

If the valve stem threads are damaged, stems can usually be re-threaded inside or outside by the use of a valve repair tool without demounting the tyre. Make certam that every vaive has a cap to prevent din, oii and moisture from damaging the coke.:

, , , ,

I

I

Fusible Plug I

- --

The fdsible plug may also be defettive or improperly installed. Use a soap solution to che'ck fusible plugs for leaks before removing tyre. Leaks c'an usually be pinpointed to the plug itself (a poor bond behbeen the fusible material and the plug Body, or to the sealing gasket used. Be sure the gasket is one specified by the manufacturer and that it is clean and free of cuts and distortion.

If excessive heat has caused a fusible plug to blow, the tyre may be damaged and should be replaced. After a fuse plug in a wheel blows, the wheel must be inspected to ensure there is no mechanical damage (using NDT equipment as appropriate) in accordance with the manufacturers instructions. This will include bearings and all wheels on the bogie.

Pressure Release Plug

The inboard wheel half may contain a pressure release plug, a safety device that prevents accidental over inflation of the tyre. If the tyre is over inflated the pressure release plug will rupture and release the tyre pressure. A soap solution can be used to check a release plug to determine whether or not it is defective.

Wheel Base

Air escaping through a cracked or porous wheel base is best found using an immersion test.

A porous wheel base can often be fuzed by proper painting or by an impregnation process. Generally, air losses because of holes drilled into the wheel base for mounting purposes can also be resealed in accordance with manufacturer's instructions.

Cracks in the wheel-well area, in most cases, cannot be repaired.

"0" Ring Seal

A defective seal between the wheel halves can usually be detected in an immersion test by bubbles emerging through the centre of the-wheel. Check to see that wheel bolts are properly torqued. hleaking "0" ring mustbe changed.

1 \ \ I

1

Beads and Flanges I

Check the bead and flange areas of afire . , f6r leaks before demounting This can be done either by immersion or by ,using a soap solution. Any of the- following factors Icdn cause-air loss: 1 I

I \ I j i I

' I

1. Cracks or scratches lin wheel bea,d ledge or flange, area. -2. -Ekceptionally dirty or borrodeh surface on whee1,bead seating - - - - - - - - - surfaces. --

3. Damaged or improperly seated tyre bead.

Bead Stick

Tyre levers may be used to break tyre bead stick on wheels equipped with a tyre and tube combination only when no tyre removal machine is available or when the design/size of the wheel and tyre make the use of the machine impossible. Tyre levers MUST NOT be used on wheels designed for tubeless tyres if the sealing between the wheel and the tyre are to remain unimpaired. Tubeless tyres must only be freed from the wheel bead seats using a tyre removal machine.

Flanges, Brake Drums and Drive Blocks

There is a need to consult the relevant bay servicing procedure for comprehensive details relative to a particular type of wheel. In general the procedure is as follows:

1. Clean all metal parts with trichloroethane and dry with compressed air. Ensure bead seat areas are free from grease, oil and foreign matters, eg rubber smears. Clean roller bearing component with white spirit, thinners or trichloroethane. Bearings corrode rapidly if left un-greased. Ensure cleaning, examination and re-greasing completed within 30 minutes. Alternatively, coat bearing with light oil, eg DTD585 temporarily until required for assembly, then clean and re-grease.

2. Examine wheel flanges for chips, dents and flats. Such damage may be dressed out with a smooth file and emery cloth, blending out any sharp edges. Dressing must be kept within permitted limits for

- ----1:-.l-1- -I------ I--&--- 4.- 4.L- --I + ---L--l ----L1:--+:--i TL:- ~lcgllglulc ucullasc ( I C L C I - - L U - L I I C L C L C V C U L L W L A C C I pu U L I L a L l u L I j . L L L L D

damage must not be of sufficient extent to break into the surface of the tyre from where it contacts the flange. Examine the lock-ring (if fitted) and groove for damage and corrosion. Brake drums must be inspected in position in the wheel. Removal following inspection will depend on type of brake d r u k and extent of damage. Heat may cause excessive distortiod ahd the appropriate gauge must be used to check the internal diameter of the drum at several points for ovality and irregular wear, iaw the specified limits. Examine for excessive scoring and also for transverse fissures caused by 'crazing', a condition liable to occur in bimetal brake drums, due to the effects of differing temperatures at the inner and outer periphery of the drum when the brakes are applied.

Wheels designed for disc brakes have only drive blocks fitted and no other brake components fitted - usually.

3. Check the drive blocks for wear, security, damage and corrosion.

After any repair or inspection of a wheel or its parts, particular care must be taken to restore any protective finish.

Wheel Bearing End-Float

Tapered roller bearings fitted to aircraft wheels require careful adjustment when the wheel is fitted to the aircraft axle. The predetermined clearance 'End-Float' ensures the bearing is free running and the bearing takes u p its normal path. It also allows for expansion because of heat generated by the brakes.

- 32 -

rnoodull l A-1283

GENERAL

An inflated tyre is a potentially explosive device. Mounting and demounting of aircraft tyres to/from the wheel is a specialised job that is best done with the correct equipment and properly trained personnel.

The following precautions are advisable in handling both tube-type and tubeless tyres, especidly those with high inflation pressures.

1. Inspect fusible plugs. Fusible plugs are used on the tubeless wheels of high performance aircraft to relieve excessive pressure created by excessive brake heat. Fusible plugs are generally not removed during a routine tyre change unless they are defective or the wheel is subjected to degreasing and cleaning. They are, however, always removed and inspected during wheel assembly and overhaul.

2. Prior to removing the wheel/ tyre assembly from- the &craft, compieteiy defiate .the tyre*<n\,a deilation cap.-it isrgoo~a~ractice to deflate the tyre before removing the axle nut. When all pressure has

8 ' been relieved, remove the valve core. Valve cores under presslure can

, be ejected like a bullet, so be careful. If the wheel ot tyre is sQspected, approach the tyre from -- the - front or rear, not from the side (facing the wheel). ,

I -- - -// ' ', ",

I -- _,

3. Take special care wien encoudtefing difficulty in frkeing tyre beads i froq wheel flanges. in^ to p b beads free incorrectly may cause an accident. Even with special tyrk i;krnoval tools, care imust be taken to '

damage t o beads-or-whekl flanges. On s.all_&res, successive ---pressing with a two-foot length of wood close to1 the-bead or tapping

with a rubber mallet is generally sufficient. On large tyres, a hydraulic or mechanical bead-breaking press may be required. If using a 'bead-breaking' press, some method should be used to prevent further movement of the tyre bead after it is broken away from the bead seat area.

WHEEL CHANGE

Always consult the AMM, but the following points should be noted:

1. Jack aircraft with wheels clear of the ground using (Bottle) jacks or conventiondl jacks.

2. Deflate the tyre. 3. For a disc brake assembly lock brake rotors in position by applying

parking brake or adjusting the brake pack assembly to fully on. 4. Remove axle nut, wheel speed transducer, tyre pressure transducer

etc.

5. Fit threat protector and use wheel transporter if necessary. 6. Slide wheel off axle. Retrieve bearings. 7. Inspect brake pipes for damage, security, locking and leaks. Inspect

the brake assembly for wear, damage and security. Inspect the landing gear leg and axle for wear, corrosion, damage and security. Check cables and equipment relating to tyre pressure indicators, anti-skid units and brake cooling fans.

8. The fitting of a wheel is similar to removal with the addition of greasing the bearings.

9. After fitting, checks should be made on:

(a) Wheel end float. (b) Brake operation. (c) Anti-skid and fan operation. (d) Tyre pressure indication.

-

---La T - n mr\ A rrr-T3 A - TT7T T P P T JsbrmJsa 1 u i-uJsLmr 1 vv nnf i~a -

1

corrosion Removal

The procedure to be used when removing corrosion from a particular wheel would be laid down in the relevant servicing manual. The following is an example of a typical procedure. Slight damage or corrosion may be removed by polishing out with a smooth hone or grade 00 carborundum cloth. Restore the protective finish.

Blend Out-Slight Damage - - . . . -

This can be affected by careful use of a smooth file, hone or carborundum cloth. Lightly polish away the damage taking care to stay strictly within the limits laid down in the particular manual.

Restore Surface Finish

Certain chemical pre-treatment processes may be applied where surface finish has been removed locally by dressing or chemical stripping, eg ALOCHROM 1200, an acidified chromate, is used extensively for the pre-treatment of aluminium and magnesium alloys. It produces a chromate film which increases corrosion resistance and provides an ideal surface for subsequent painting. It is not suitable for use prior to etch priming.

NOTE. It must be stressed that before commencing any pre-treatment process, any non-metallic components and fusible alloy plugs must be removed. Strict adherence to the correct procedures and precautions when using Alochrom 1200 is of the utmost importance (please see JAR modules 617 books for more detailed information).

Replacing Bearings and Seals

Replacement of bearings and seals must be canied out strictly in accordance with the Wheel Manual. Details of the pressing and removal tools for a particular wheel are given. A typical example procedure for pressing in an outer race is as follows:

1. Grease the bearing housing in the liner, steel sleeve or hub with the approved lubricant.

2. Set u p the wheel on the hydraulic press table with press tools in position.

3 . Position the distance piece over the mandrel, then place the hub over the mandrel.

4. Position the outer rat-e on-the lip of the bearinghouskg with the I - - smaller diameter of-thesleeve'in the outer race! ~ocate~,he\collar over

the mandrel. I I I '\

1 1 1 '! I

5. Ensure the head of (he mandrel is central beneath the jack. Apply the I minimum of hydraulic pressure necessary to push the outer race in -

its housing. , , -- - - - / ,

I I I

I I FIXED STRU,CTURE OF HYDRAULIC,

--

- -

OUTER RACE EXTRACTOR

BOTOM OF HYDRAULIC PRESS \ END CAP

Fig. 14 BEARING PRESS

- 35 -

moodull l A-1 286

WARNING

It is important that no excessive pressure is applied to the outer race during this operation. Investigate the cause if the outer race will not seat correctly. Failure to adhere to this instruction could result in irreparable damage to the wheel structure.

Remove the press tools and check with a 0.0015in feeler gauge that there is no gap between the base of the outer race and its seating in the bearing housing. If the gauge fits into a gap, set up the wheel press and tools as before and apply further pressure.

The inner race would normally be placed in position after being lubricated and retained, with the seal in its housing by a circlip.

Replacing-Fusible Plugs I

The follolwing is an example of the procedure to be followed when fitting new fusible plugs:

:1. Smear the new seals and new '0' rings with the type of silicone grease laid down in the servicing manual, eg silicone grease. Insert the seals into the fusible plug casings and fit the '0' rings around the casings.

I I

2. ' Grease the threads ofithe plugs with the same grease type used on the seals and screw them into the wheel well to a torque loading of 16

- lb in (or as laid down). Secure plugs with locking wire to adjacent lock - - - screws.

NOTE. The '0' ring seal of a tubeless wheel should be carefully cleaned and inspected for defects before being lubricated and installed. If its condition is at all questionable, it should be replaced.

Bead Lubrication in Mounting Both Tubeless and Tube Type Tyres

It is often desirable to lubricate the toes of inner edges of the beads. This is done with an approved talcum powder only. This will facilitate mounting and seating of the tyre beads against the wheel flanges. Care must be used with tubeless tyres, however, to ensure that none of the talc gets on the sealing area of the bead.

rnoodull l A-1287

WHEEL/TYRE ASSEMBLY BALANCING

It is important that aircraft wheels and tyres be as well balanced as possible. Vibration, shimmy or out of balance is a major problem. However, in most cases tyre balance is not the cause. Other items affecting balance and vibration are: installation of wheel assembly before full tyre growth, improperly torqued axle nut; improperly installed tube; improperly assembled tubeless tyre; out of balance wheel halves; poor gear alignment; worn or loose gear components; flat spotted tyre. In addition, twin tyre inflation not equal and dual tyre diameters not matched can cause an unbalanced condition.

Balance marks are placed on many tubes to indicate the heavy spot of the tube. These marks are often paint stripes about ?4 inch (12mrn) wide by 2 inches (50mm) long. When a tube is installed this balance mark must be aligned with the light spot' balance mark of the tyre (red dot).

When-mounting tubeless tyres,-the-balance mark on the tyrre-k-aligned with the wheelvalve, unless otherwise specffied-bx the manufacturer. - , , ,

I i I i

With some split wheels, the light1 spot of thb wheel halves is iddicated with an 2' starnpgd on the flange. In assembling the& &heels, position thk 'L's 180 /degrees apart. If additional dynamic or static b d ~ ~ i ' n ~ is required after,tyre mounting, many +heels have provisions for attachiqgaccessory balance /w&ight$afbund the circumference of the flanges. 1 \,, -,' I -

I \ I

\ I

1 ~ n b d a n d e d wheelk can also proddce v ib ra th . Be sure to have wheels balanced according to in/structions specified :by the aircraft or wheel manufacturer. They will nom'ally be statically and dynamicalw balanced, ie balanced in two planes.

I - - I - A' --

Goodyear manufacturers a lightweight, portable, low cost balancer to assist operators in achieving proper assembly balance.

AIRCRAFT BRAKES

Brakes are designed to convert the kinetic energy of the moving aircraft into heat energy by friction. They must produce enough friction and have enough heat capacity for the weight of the aircraft and must also be able to dissipate the heat generated.

They are not allowed much time to perform this task. In the case of an RTO when the aircraft weight and speed is high and conditions are much more severe the brakes are tested to their upper limit. This can be shown in the fonn of a Power/Time Curve where the kinetic energy absorption rate is depicted.

JAR25 specifies the braking ability standards to be met.

TYPICAL FULLY BRAKED LANDING BAC 1-11 (39,500kg

O c 87,0001b)

200 -.:--*---- ------- * . I * 45 SECS 30 MlNS

TIME -5 , ' LANDING COOLING

Fig. 15 TEMPERATURE AND POWER CURVES

In the early days drum brakes were used, operated by air supplied by a compressor, or the aircraft engine. These brakes suffered from brake fade and distortiondue to heat and were-soon replaced by the copper disc brake.

L. - .- - --

~hese':brakes operated a t higher temperatures and were made of copper (to dissip&t+ the heat and allow a Aore even temperature around the disc) and coated with nickel chrome to provide ahard wearing surface. Brake design developed and the copper disc brakes were replaced with rnulti disc (segmented) brakes. These consis$ of a torque plate housing the hydraulic pistons, which act on the pressure plate which in turn forces the stators and rotors together, with the reaction force being applied by the thrust plate. This whole ,assembly is known as a heat pack. Concoi-dl stators and rotors are made of carbon because of the intense heat generatdd. /

-

-

Friction

The amount of friction developed in a brake will vary depending on many factors but in general the amount of friction generated is given by the equation:

where F = Friction generated.

p = a coefficient given for the type of surfaces in contact.

RN = Reaction Normal. The force pushing the brake pad against the disc or drum.

- 38 -

rnoodull l A-1 289

It can be seem therefore, that in general the greater the force (RN) the greater the friction. One of the reasons why discs are better than drums is that the drum distorts with the high radial forces created by the brake pads or shoes. This means that with a disc brake greater RN can be applied without fear of distortion of the discs. Also with modern heat packs the total surface friction area is greater than with a comparable sized drum brake.

DRUM BRAKES

Although used extensively on earlier aircraft, drum brakes have largely been superseded by hydraulically operated disc brakes on most modern high performance aircraft. Pneumatically operated drum brakes may still be found in service, however, and the construction, operation and maintenance of a typical brake unit is described in the following paragraphs.

- I -- --- --- -

The main components of the br+e-unit-@the back plate, brake-drud, expander tube (pressure bag) and brake livings. \ \ \

I ' I 1 I

1

RNETS SECIJR~G EXPANDING BRACKETTO BRAKE LINING TUBE \*< BRAKE LINING i

/BRACKET i ,

t BACK PLATE

I

NUT AND BOLTS ,

TO AXLE 61-&1

SPRING

a SPRING CLIP SECURING BRAKE UNIT

AXLE FLANGE

ROTATING BRAKE

DRAWING FROM CAP 562

Fig. 16 DRUM BRAKE

Back Plate

This unit is cylindrical in shape and is attached to a flange on the axle. It houses the expander tube, brake linings and pneumatic connections.

Expander Tube

This is a circular, reinforced rubber tube of flat cross-section and is fitted around the back plate. It has a pneumatic connection leading through the back plate to the aircraft pneumatic system.

Brake Linings

The complete brake lining assembly is made up of a number of segments of heat- resisting friction material which-form a ring around the expander tube and are sha6ec.I ti, confurrll tu tilt: irisidt: radius of tilt: brake dr UIII. Each seg~rlerri is bonded or riveted to a metal fitting which protrudes through the back plate and is secured by a spring clip.

separ+tors

Phosphor-bronze gauze separators are fitted between the ends of the brake lining segments to reduce heat penetration to the expander tube and to exclude carbon particles.

Brake Drum

The brake drum is a heavy steel cylinder, attached to and rotating with the wheel and against which the brake lining segments expand to produce the braking action. I t may be of bimetallic construction to rninimise heat distortion.

Normally fitted so there is an air gap between it and the wheel around its circumference - to allow for better cooling and help prevent heat transfer to the wheel and tyre (heat sink into the tyre will cause a pressure rise but more importantly can cause deterioration of the tyre structure and possible tyre failure.

When the pilot's control is operated, air pressure is applied to the inside of the expander tube which expands and forces the brake linings out against the brake drum. When air pressure is released the expander tube collapses and the brake linings are withdrawn inwards from the brake drum by the action of the return springs.

Removal/ Installation

Before attempting to work on the brake system or to remove a wheel, it is important to ensure that all air pressure is exhausted from the brake system. In many pneumatic systems a pressure maintaining valve is used to safeguard the brake pressure in case of a leak elsewhere or failure of the compressor, so that lack of pressure in the brake system must be confirmed from the brake system pressure gauge and not be reference to the general system pressure.

When the wheel has been removed, the brake unit can be removed by disconnecting and blanking the air pressure connection and removing the bolts attaching the back plate to the axle flange.

When installing a new brake drum, the protective treatment applied for storage purposes should first be removed with a suitable solvent such as methylated spirits. Petrol or paraffin should not be used.

i/ , . -- - - - -. .. --

When-installing the brake unit,-care mGt'bq taken to ensuie,that oilor grease r-7 does not Come into contact with the linings; Operators should jalao a~oid 'handl in~

the linihgb as the natural oils froF! the ski: n)ay have an advirs!e effect. 1'r brake linings idq become contaminatedi they must ?je considered unLerviceable; !no atternpd should be made to clean the surf$e:with solvents. ~ ;. ;

! L-.- ~/' j i / J :

i i----.,, ,,

i j -.,' /

lnspectio+ I j ,!'?

! c I i I Drum qrakes are not norrnally aycessible fyk gisual inspectiori dhen installed on

the aircrdLDdring a preflight inseectionthk back plate and wheel should be examihed for signs of overheatlngand-theflI'exible pneumatiE h o d between the brake units and the landing gear leg should be checked for damage, security or leaks. Operation of the brakes may be checked by means of the brake pressure gauge and also by checking that air is discharged from the brake relay valve when the brakes are released.

At the times specified in the maintenance schedule and whenever unsatisfactory operation is suspected, the brake unit should be removed for inspection and overhauled. Disassembly, which should be carried out on a rubber or felt covered bench, is normally straight forward, but reference should be made to the approved Maintenance Manual for details of any special procedures or tests required.

Brake segments should be exarnined for wear by measuring the thickness of the remaining material, the minimum thickness permitted for replacing the linings being laid down in the AMM. Any carbon deposits should be removed with a stiff bristle brush.

- 41 -

moodull l A-1292

Test After Reassembly

Following reassembly the complete brake unit should be installed in an appropriate sized test brake drum and submitted to pressure tests as prescribed by the manufacturer. No leakage should occur and the linings should return to the 'off position a soon as air pressure is released. The most suitable means of detecting a leak in the expander tube connection is by applying a solution of non- corrosive soapy water which, subsequently must be washed off.

DISC BRAKES

Most modem aircraft are fitted with hydraulically-operated disc brakes (also known as plate brakes). Light aircraft generally have a single-disc type and larger aircraft, a multi-disc type (also known as segmented plate brakes, heat packs etc).

- - - - - - . -

DRUM BRAKE (PNEUMATIC) DRUM BR~KE(HYDRAULIC)

SEGMENTED HYDRAULIC DISCS (ROTORS)

1000 HP KE 1% ft Ib COPPER DISCS

DRAG 2 TONS ROTORS (FOUR)

BRAKE PADS

PISTONS (TWO) 2000 m KE 30n f DRAG 4 TONS

COPPER DISC BRAKE MULTI DISC HEAT PACK

Fig. 17 EXAMPLES OF BRAKE UNITS

OPERATING CYLINDERS

ROTOR

STA

25 HP KE 50M ft Ib DRAG 4 . 8 TONS

-- - i---- -- -

,- Fig. 18 CONCORDE~'~AKE UNIT (CARBON) - , \

I ' , \ ' 1 'i

\ '\ 1 1 I 1 I '

I I I

D U N L ~ P HYDRAULIC BRAKE NIT (COPPER DISC) i / 1 /' I ; 1 1 __,

An old; unit and included here o,nly&givt(& historical background-io the developnient of the disc brake unit. ~ a & ~ o f \ c o ~ ~ e r to allow &en heat distribbtion and I hence , less ch&ce of dist&!ion.

I I I

I I I ' I I

1 i The cob$er dis,e'k be re driven bd t$e wheey amd chromium plz@d to provide a hard

1 wesu;idg kuda<e./ (The discs were bapagable o 'being re-plated)..' !,--

:__.. -- .d C ___ . _-, j

The torque plate was bolted directly to the axlal/landing gear unit and hydraulic pistons would clamp the whole assembly, during the braking operation, to provide the retarding friction.

This type of unit required regular adjustment to allow for wear as no wear compensators were fitted.

The maximum permissible wear of the friction pads fitted to this unit is indicated by a dimension taken from the face of the torque plate to the friction face of the first brake plate, with the piston rods screwed fully home. When the limiting dimension is reached, all the friction pads must be renewed.

No direct measurements are involved when checking the copper brake plate. Scoring of the plates is permitted until the plating at the base of the scores is broken. If any lifting of the plating is evident, the plate is classified as unserviceable and the inner and outer plates must be changed as a matched pair. Tenon minimum permissible width is measured with a wear gauge special tool.

rnoodull l A-1294

SINGLE-DISC BRAKE UNITS

A simple single-disc brake unit is shown and is of a type found on many light aircraft. A single operating cylinder is shown but two or three are often used for increased braking performance and larger aircraft may have brakes using five or six cylinders. The brake unit consists basically of a light alloy torque plate shaped for attachment to the landing gear leg or axle flange, housing a calliper- type hydraulic jack unit and a pair of friction pads. A solid steel disc is slotted into the wheel and rotates between the friction pads.

When the brakes are operated, fluid pressure is applied to the cylinder and forces the operating piston towards the disc, thus squeezing the disc between the operating and fixed friction pads and thus resisting wheel rotation. When the brakes are released the disc is free to rotate between the friction pads.

SEALING PISTON - - - - -

m ..,m- s,. m . .- -\ - fin-.-. .L -.-..,A -.--.--I, n T UWULIL u r c w I IIUU 1-13 I UIU - --


Fig. 19 SINGLE DISC BRAKE

The brake unit should be examined periodically for fluid leaks, damage or corrosion, the friction pads for wear and the discs for scoring or pick-up of surface plating. The single discs used on light aircraft brakes are prone to corrosion and pitting and this may lead to rapid wear of the friction pads. Discs in poor condition should be replaced or machined to give a clean surface as appropriate. Discs worn below their minimum width dimensions should be replaced. Replacement of worn pads is normally a simple procedure once the wheel has been removed and often does not necessitate breaking down the hydraulic system.

rnoodull lA-1295

THE MULTI DISC BRAKE UNIT

In general they are made u p of several plates (constructed in interloclung segments to prevent distortion). These plates are called rotors and rotate within stators. The rotors are driven by tenons grooved into drive blocks in the wheel and the braking is achieved by several hydraulic cylinders spaced around the unit.

The friction pads may be made from sintered metal or more conventional pad material.

A typical multi-disc brake unit is shown. In this unit a torque plate and torque tube assembly fits over the axle and is bolted to a flange on the axle; alternative designs are often similarly mounted but prevented from rotating by means of a torque arm attached to a suitable furture on the landing gear leg or bogie. A number of cylinders are spaced around the torque plate, connected to the hydraulic brake system and house pistons which apply force to the pressure plate. _ _ -- - - - - -

! 1- \ - \ ', -- 7 r---, \

The dlsc pack contains alternate stationary and rotating discs the-Gtationaq discs being keyed to the torque ;tube and !he rotating discs dei'Ag keykd fo drive blocks in the wheel hub. In this1 ubit the stationary discs house the brake pads and tHe rotating discs are segmented -_- to p5e&ent head distortko: and brake drag. ~ o r r e d t working clearance in thk disc packjs maintained by 'means-df adjuster assemblies. Pins attached to the presshe plate and protruding-through the torqud plate on t q s brake unit, indicate the kmount of wear which has taken place in the disy'pack. I I I I

I

I I / I I i ' ,' / I ' ~i A hpher- typ6 of'multi-disc br&e 'is-kno&nis a tri-metalli~~br*. Construction /' is similarto-the-brake unit described-except that the rotatihg-discs have a metallic

compound sintered to their faces and steel segments, known as wear pads, are riveted to the faces of the stationary discs. Alternatively, the faces of both sets of discs may be sintered, or the stationary discs may be plain.

When the brakes are applied, hydraulic pressure is admitted to the cylinders and moves the operating pistons against the pressure plate. The friction loads generated between the stationary rotating members provide the required braking action. When the brakes are released, springs in the adjuster assemblies move the pressure plate back to maintain a working clearance in the disc pack and permit free rotation of the wheel.

Maintenance

Contamination of the friction surfaces of a brake unit by fluids used in aircraft servicing operations is highly detrimental to brake operation. I t is essential, therefore, to protect brakes from contamination by fuel, oil, grease, paint remover and de-icing fluid, etc. If the brake unit is contaminated then it should be changed.

Installed disc brakes may be inspected for signs of fluid leakage, external damage, corrosion, disc pack wear and overheating, and the associated hydraulic pipes for security, distortion, chafing or leaks. Brake disc pack wear can be checked by measuring wear pin protrusion, the limits being specified in the AMM.

ROTOR DRIVE AUTOMATIC WEAR BLEED SCREW PRESSURE BLOCKS

ADJUSTER \ / STATORS /


Fig. 20 MULTI-DISC HEAT PACK ASSEMBLY

In some installations a worn disc pack may be exchanged after removing the wheel and thrust or back plate and without disconnecting the hydraulic connections, but in order to carry out a detailed inspection the brake unit must be removed from the axle.

At the periods specified in the maintenance schedule the brake unit should be removed for inspection and overhaul. The wheel should first be removed and the hydraulic pipe couplings should be disconnected and fitted with suitable blanks.

rnoodull l A-1297

In some cases fluid will drain from these pipes and bleeding will be necessary after re-connection, but in other cases connection is by self-sealing couplings, which isolate the hydraulic system from the brake unit. The brake unit attachment bolts (and, where fitted, the torque link) should then be removed and the unit withdrawn.

Following its removal, the brake unit should be dismantled, cleaned and inspected. All metallic components should be thoroughly cleaned and dried; if chemical solvents are used they must not be allowed to come into contact with the seals. Inspection will include:

Rotating discs should be checked for excessive scoring, corrosions, distortion and wear on the friction surfaces and driving slots. Light surface damage which would not cause excessive wear of the friction pads may be acceptable, but deep scores or corrosion should be ground out within prescribed limits specified by the manufacturer for the disc to be re-used. -- -. - - -.

L '\ -- ,

beybnd limits - --

or

I I , I

i j 1 / i

The ,torque plate, , t d r & e u b C ~ d thrust plate-shobid be examined forcracks, corrosion5-distorti6n and damage, p&-iedlar attention being paid to bolt holes and other highly stressed areas. Cylinders and pistons should be inspected for scores or other damage and springs inspected for corrosion and given a load/ compression test as specified by the manufacturer.

(d) Operation of the self-adjusting mechanism should also be checked and the friction force applied to the retraction pin measured.

Protection treatment should be applied to the metal components and the unit reassembled and tested for leaks and correct operation. It is normally specified that new seals, gaskets and self-locking nuts should be used for reassembly and all fasteners torque loaded in accordance with the manufacturer's manual. The unit should be primed with hydraulic fluid and blanks fitted.

When re-installing the brake unit on the axle, care must be taken not to spill fluid on the disc pack. Jointing, sealing or anti-seize compounds should be used where specified and all fasteners and pipe connections should be torque loaded and locked to the manufacturers requirements.

Adjuster Assemblies

A sectioned view of a typical adjuster is shown. At least two adjuster assemblies are fitted to the majority of disc brakes, their purpose being to maintain a suitable brakes off clearance between the stators and rotors. In a single-disc brake the retraction pins are often attached directly to the operating pistons but on multi- disc brakes they are usually attached to the pressure plate. In operation, movement of the piston or pressure plate is transmitted via the retraction pin and friction bush to compress the adjuster spring and move the guide until it abuts the torque plate.

FRICTION BUSH

SPRING HOUSING \ W 1

!?ETP"A..CT!n!'! PIN

PLATE

DRAWING FROM CAP 562 - - -

Fig. 2 1 AUTOMATIC WEAR ADJUSTER

When the brakes are released the adjusted spring pulls the guide back until it contacts the spring housing, the clearance between the guide and torque plate being the designed running clearance. As wear takes place in the disc, the pressure plate has to move further forward, thus pulling the retraction pin through the friction bush by an amount equal to the disc wear, but maintaining the design clearance when brakes are released. On some brake units wear may be assessed by measuring the protrusion of the retraction pin.

On initial assembly a special tool is used to position the retraction pin a t the position of maximum protrusion through the friction bush. The pin takes up its initial operating position when the brakes are first applied.

Correct operation of the adjuster assemblies must be checked whenever the brakes are tested and should result in free rotation of the wheel when brakes are released.

CARBON BRAKES

These are similar to the conventional multi-disc brake unit except the rotors and stators are made of carbon and are of single piece construction. They are lighter, run at higher temperatures and are more efficient.

BLEEDING THE BRAKES

The method of bleeding the brakes will depend on the particular aircraft system and reference should be made to the AMM. However, the normal method of bleeding is to pressurise the brake system, open the bleed screws fitted to the brake units and apply the brakes from the flightdeck. Allow hydraulic fluid to flow through until air free fluid is discharged; the bleed screws are then closed and brake operation tested. Bleed fluid should be piped to a suitable container and

(' - must not be allowed to come into contact with the disc pack and must not be re-

\ 1 7

"\ \,

7- -\ - -, On lo&-bressure brake systems, as fitted'to hany small aircr*, thgpressure is generated by the pilot pushing fhe appropeate brake pedal d o h .

1 I

On high'pressure brake system? (asfit$ed t6 most large airc!afk), the (associated /' hydraulic accumulator is pressured-and as the brake pedal is depresyed, fluid is ', \ forcedi out of the bleed screws uhder pressune. In this type ob systemHit is \ sometim~es recommended that only a specified quantity of fluid1 is discharged and

it may be necessary to bleed ot$e$ parts ofithe system such bsj where fitted, the servo system frbrh the brake pedqs to thelsl#ve units (four small systems each with-its 0-3-eservoir) and the normal and,emergency accy~ula tors , before bleebing-the-brakes. After bleeding the-appropriate re~ervoi~shoilld be topped u p as necessary.

Testing the Brakes

Brakes are normally tested after overhaul and after installation on an aircraft, while the aircraft is still jacked up. The wheel is spun u p by hand and the brakes applied several times and released; there should be no leakage and the brakes should stop the wheel when applied and allow wheel rotation when released.

Operation of the emergency and parking brake-system should also be tested. Special care should be taken to ensure that the hydraulic systems are correctly connected.

BIT tests may also be called for depending on the system.

Brake Temperature Monitoring System

On some aircraft, in order to inform the pilot of excessive build-up of heat in the wheel brakes, a brake temperature monitoring system is fitted. A typical system includes a temperature sensor at each wheel, which supplies information to a central monitor and warning devices on the flight deck. The monitor contains a temperature gauge and a selection button for each wheel. The gauge normally records the temperature a t the hottest brake and a button illuminates when the associated brake temperature exceeds a predetermined amount. When any button is pressed, the gauge records the temperature at the associated brake.

On aircraft with CRT display systems brake temperatures (and tyre pressures) are displayed in colour on the screen.

For testing purposes, operation of a BIT test switch will cause system decals to illuminate and gauge readings to move to the test signal range when all circuits are serviceable. - -

\

- -

Routine maintenance includes inspection'of the sensors and asSociated wiring for security and damage and functional tests of the system using the appropriate test switches and BIT systems.

With the 'Glass Cockpit7 type of aircraft su6h as the Boeing 747-400,-B777 and the Airbus aircraft, flight deck indications are via CRTs. , - -

The CRTi~anding ~ e a r page will show details such as: I

~r Brake temperatures * Tyre pressures --

* Landing gear and door positions, etc.

(See the books in this series entitled Hydraulics; Landing Gear and Instrumentation.)

SKID CONTROL

The braking systems of most modern aircraft are provided with a means of preventing the wheels from skidding on wet or icy surfaces and of ensuring that optimum brakmg effect can be obtained under all conditions, by controlling the hydraulic pressure to the brakes automatically. Anti-skid units/ transducers sense the rate of change of wheel deceleration, decreasing the hydraulic pressure applied to the brakes when a high rate of increase in deceleration occurs (ie consistent with an impending skid) and restoring the pressure to the brake as the wheel accelerates again.

- 50 -

rnoodull lA-1301

A modulator is often fitted in conjunction with the anti-skid unit, to restrict the flow of fluid to the brakes after initial brake application and to conserve main system pressure. There are basically two types of anti-skid systems in use, the mechanically controlled system and the electronically controlled system.

The mechanically controlled system (rare) can use either a wheel mounted unit or an axial mounted unit

The anti-skid control valve (automatic) is positioned in the hydraulic brake line between the brake control valve (controlled by the pilot) and the wheel brake unit. When autobrake is selected, it also supplies fluid to the brakes via the anti-skid valve.

MECHANICAL SYSTEM (MAXARET)

The anti-skid unit is mounte$on_the_bra_ke unit torque platepor_within the axle bore-andfis connected into thd-brake hydraulic circuit at the brake.&it. The anti- skid u p t consists of a valve assembly conndcted to a flywhedl &hicQis driven by

I ] its asqociate wheel. \ I

,' 1 I I I ' I I I I I I

1 1 I I I I I 1 I )

Fig. 22 MECHANICAL ANTI-SKID UNIT (WHEEL RIM DRIVEN)

The Wheel Rim Driven Unit (Not common these days)

This unit is mounted either on the wheel brake unit or on the landing gear assembly so that the rubber tyre rests on and is rotated by the rim of the aircraft wheel.

The unit consists of two main parts: a hydraulic control valve (operated by the flywheel); and the flywheel assembly. The flywheel is housed inside the rubber tyred shell which is rotated by direct contact with the wheel rim.

moodull l A-1302

The control valve has four connections:

* Pressure supply - from the pilot's brake control valve. * Return - to the hydraulic reservoir. * Brakes 1. * Brakes 2.

Fig. 23 FLYWHEEL UNIT - EXPLODED V I ~ W I

NOTE. With-this-type of unit the spring acts as a sort of memory- when the wheel decelerates too quickly, as determined by the spring, then the valve is caused to operate and reduce fluid pressure to the brake.

As shown in the exploded view the drive from the rubber tyred shell is transmitted to the flywheel via a drive ring, drive spring and drum. The drive pegs of the shell engage with the slots on the drive ring to drive the drum through the spring. One end of the spring is connected to the drive ring while the other is connected to the flywheel. The spring is coiled in such a way that as the unit is being driven the spring expands and drives the drum (and flywheel) by a positive friction drive.

The flywheel, mounted on the outside of the drum, is driven through its spoke by two segmented bosses attached to the drum. A lightly loaded main spring links the flywheel to the drum and tries to keep the flywheel spoke in contact with the driving faces of the segmented bosses.

rnoodull l A-1303

Two thrust balls are located in the flywheel spoke and are kept in contact with a cam on the drum by a thrust plate attached to the end of a spring-loaded thrust rod.

The thrust rod is connected via a system of levers to the hydraulic control valve.

With the brakes applied and when the landing wheel starts to slow down too quickly - that is, starting to skid - the rubber tyred shell slows down at the same rate, as does the drum (because of the friction drive of the spring). But the flywheel, with its momentum does not slow down and starts to overrun its position against the load of the main spring. This will cause the flywheel spoke to make contact with the faces of the two-segmented bosses, which will drive the drum against the frictional resistance of the drive spring.

. ~ .

/--

PRESSURE SU,PPLY

I

i

!

I I I

1 , 4 NORMAL BRAKING CONDITION '

- .- ' ; , L--' ' ,,,' ; WHEELRIM

r-- L-.

8 ANTI SKID CONDITION 1-1

Fig. 24 SCHEMATIC OF THE UNIT

This relative rotational movement between the flywheel and drum causes the two thrust balls to ride u p the face of the cam and push the thrust rod axially to release the brake pressure.

Movement of the thrust rod will shut off brake pressure supply and open the line from the brakes to return.

After the unit operates the flywheel is slowed down by the drag of the drive spring and with the brakes off, the landing wheel regains speed. This will cause the drum to begin to drive the flywheel again and the thrust balls and thrust rod will regain their positions (with the aid of the return spring), and normal brake pressure will be re-established.

Should the wheel bounce clear of the ground during landing the wheel will stop instantly. This will cause the anti-skid valve to release the brake pressure completely and re-apply it some 4 seconds later, which should be enough time to allow the wheel to make contact with the ground again. -

Installation

The mounting details of the various types of mechanical units vary considerably and reference should be made to the AMM:for details of any particular installation. The whole unit is spring-loaded, or the mountings shimm'ed, to maintain satisfactory driving contact with the aircraft wheel rim. The tyre loading is normally checked after installation by measuring the flat produced on the rubber tyre at its point of contact with the aircraft wheel. i

AXLE MOUNTED UNIT

An axle-mounted unit is driven by means of a shaft, which is splined into the anti-skid unit a t one end and into a drive housing bolted to the wheel hub, a t the other. All types of units are marked with the correct direction of rotation and this must be checked before installation.

The unit contains components similar to those found in the rim driven type. Additionally the unit also includes a sun and planet gear assembly - to increase its speed ratio with the wheel and a clutch to absorb sudden start up loads.

Figure 30 shows the position of the components during normal braking,

If the landing wheel decelerates too quickly (an impending skid) then the flywheel will run-on relative to the input and main shafts. This relative movement between the main shaft and the drive ring will cause the two thrust bearings to ride up the cam profile and move the thrust rod against its spring to open the control valve and release the brake pressure.

moodull l A-1305

MAIN SHAFT DRIVE RING CLUTCH FRICTION

MAIN SPRING VALVE THRUST ROD

\ CLUTCH SPRING

Fig. 25 AXLE MOUNTED ANTI SKID UNIT 1.

THRUST BEARING CLUTCH

->

- -

@ NORMAL BRAKING CONDITION

Aircraft landing wheel decelerates

The inertia of the flywheel causes consistant with an approaching skid

2 the drive nng to advance 30 deg and the main shaft decelerates i n O i n r e l a t i o k the shaft proportion

RETU

The thrust rod opens \ The 30 deg movement of the drive

4 the valve and relieves the pressure in the brakes ring forces the thrust bearings to @ ANTI SKID CONDITION @ ride up the cam face

Fig. 26 SCHEMATIC OF THE AXIAL MOUNTED UNIT

moodull l A-1306

Bleeding

Bleeding of the anti-skid unit is normally achieved when bleeding the main brake system but independent bleeding may be necessary after installation. This is carried out by fitting a drainpipe at the exhaust connection, rotating the drive smartly in the direction of rotation, then bringing it to rest.

Each time rotation is stopped, fluid will be discharged from the exhaust port and bleeding should be continued until the discharged fluid is free from air, then the pipe connections remade.

Inspection of Mechanical Anti Sid Units

A t the periods specified in the maintenance schedule the unit should be inspected as follows:

\ --

(;;; TL- ---:+ .-.L---lA 'ha ~ \ l a o + - , ~ A -mil ; ~ P ~ P P + P A fni- P P P I I Y ~ ~ J ~iornc ~f ~ r r c . ULUL OALVUIU ~b UAU-I~U -IU I A A V ~ U U C U U r v z u-urn--- J 2 --?--- corrosion, external damage and leaks (with brakeg applied)

(b) The pipelines should be checked for damage, distortion and the connections for security of attachment, and leaks.

- - - -

(c) The driving tyre and wheel should be inspected for correct loading and alignment and the tyre for excessive wear. ,

1 NOTE. It is possible to lock the spring-loaded type units out of contact with the wheel track by inserting a pin in the mounting stud. This is done to allow wheel removal, and it is recommended that a red streamer-should be attached to the pin as a visual reminder that the anti-skid unit is out of operation.

Maxaret Indicator or Pin Position

With the brake on the indicator rod should be approxim.ately flush with the surface of the unit.

ELECTRONIC/ELECTRIC ANTI-SKID SYSTEM

The system comprises a wheel speed transducer, a control unit (computer) and an anti-skid valve in the brake pressure line, together with associated switching, BIT and check-out and warning lamps. The wheel speed unit may supply either dc or ac depending on the type of system used. Operation is basically similar to the mechanical system but the use of logic circuits in computers enables much finer control to be exercised.

moodull lA-1307

Further refinements such as strut oscillation damping circuits, touch-down protection and locked wheel protection and auto brake may also be incorporated and some systems automatically de-activate a t low speed to prevent interference with normal taxing manoeuvres.

Wheel s h d control is achieved by sensing the wheel rate of change by a tachometer or generator. If the wheel slows down a t too high a rate compared to some parameters (eg IRS computed speed) then a signal is sent to the anti-skid valve to reduce the pressure. If the wheel continues to slow down too quickly (approaching a skid) then pressure is reduced still further.

If this fails to prevent the wheel from skidding then a further signal will cause all remaining brake fluid pressure to return to the reservoir. Only when the wheel regains its speed will the pressure be allowed to the brake. (All this takes place in a fraction of a second.)

Depending on the system the wheel s ~ e d may be compared to: - I I . -

I \ -- - \

- c - - -

JF Memory in the computer. ,, \

-k The inertial reference ,speed of the aircraft. , ,

I , , i I I

I I I

~ e n e r b t & r System

This is a voltage sensitive system, 1 with the, voltage being generated by an axle I 1 mounf ed 1 I

/ I I 1 ' It is ,drivkn-bfdrive bosses attadhkd-to-thb $heel bearing c&er lPlate and the output-voltage-is directly proportional--to-the wheel speed.

GENERATOR WHEEL BEARING COVER PLATE (ROTATES WITH

E C

AXLE SPRING

Fig. 27 AXLE MOUNTED GENERATOR

- 57 -

rnoodull l A-1 308

When the wheel slows down the voltage drops. This information is used by a control unit. If the rate of slowing down is too great (voltage drop too quick) then the control unit will send a (dc) signal to a solenoid on the anti-skid valve. When energised this will release the brake pressure.

With the brakes off the wheel will regain its speed, the generator voltage will be re- established, the control unit solenoid will be de-energised and the brakes will be re-applied.

Tacho Probe System

This consists of a tacho probe (a magnet around which is wound a coil) which is caused to produce a pulsed dc (a sort of ac) output by being close to a ferro magnetic exciter ring. The frequency of the output of the probe is dependent on the speed of the exciter ring (attached to the aircraft wheel). This signal is sent to a comparator/computer for processing. -- .-

BRAKE UNIT

EXCIT'ER RING

Fig. 28 TACH0 PROBE WHEEL SPEED SENSOR

The probe may be housed in the axle or fitted as shown in figure 28.

As the wheel starts to slow down so the frequency output of the probe drops. This rate is compared by the control unit/computer with either a known maximum frequency reduction, or the actual rate of aircraft deceleration (from the Inertial Reference System).

If the deceleration rate is high enough then a signal will be sent to the anti-skid valve solenoid to reduce/stop the pressure to the brakes.

If the wheel retardation rate is being compared to the aircraff retardation rate then the wheel can be slowed a t a rate which is a percentage of the aircraft slowing rate - thus maintaining maximum braking efficiency.

This system is more sensitive than the others and therefore the brakes can be kept a t just the right level of braking for maximum efficiency.

I HYDRAULIC PRESSURE

BRAKE CONTROL VALVE

Fig. 29 ELECTRONIC ANTI-SKID CONTROL SYSTEM

ACCELERATION RATES FROM - - HYD~T', ---,.A. ,-- R ~ J R N

INERTIAL UNIT - - rnraaunr , SIGRAL +O ANTI SKID

I ' VALVE \

Control units/ computers normally contain circuits which provide warning failure in the system and a self-test facility which enables the serviceability of the various components to be checked. Controls for the operation and testing of the anti-skid system are contained in the control unit/computer and on the flight deck.

Maintenance

/ I

--- ' /' 1 - - --

I I

The inspection, testing and maintenance of any particular anti-skid system will vary considerably between different installations and details should be obtained from the AMM.

I , I '

I I

- 59 -

rnoodull l A-1310

WHEELSPEED AND , , ACCELERATION I COMPUTER 1

I . . ' \ \ I /

\ ' ', 4 \ I I PRESSURETO

TACHO,;SIGNAL OF WHEEL I SPEED AND DECELEFXTlON

I , I BRAKES

1 1 ,

/ \ - ' I m COMPUTER ,-,

2 - - - 1' WHEEL AND BRAKE ASSEMBLY L -

However, the self-test facility normally enables complete testing of the system to be carried out and the test circuit is designed to facilitate location of faulty components. A visual inspection of the system should include the following:

(a) The various components should be examined for damage, security and where appropriate, fluid leaks.

(b) Pipelines should be examined for security, chafing and fluid leaks, particularly at connections.

(c) Electrical cables should be examined for security, chafing and damage by fluids or heat.

The following is a general description of the A320 and B767 anti-skid systems. There should be no need to remember the details of each system but the principles should reinforce those already learnt.

The anti-shd system is based on the optimisation of the aircraft deceleration rate a s measured by the ADIRS (Air Data Inertial Reference units): Maximum braking is provided by maintaining the wheels a t the limit of an impending skid. ,

- - I I -

1 r-

, HYDSUPPLY I

AMPLIFIER

Fig. 30 A320 SIMPLIFIED ANTI SKID SYSTEM

Brake release orders are sent to the four normal and to the four alternate servo valves as well as to the ECAM system which displays the released brakes. An ON/OFF switch in the flight deck activates or deactivates the anti-skid system and nose wheel steering.

The speed of each main gear wheel (given by a tachometer) is compared with the aircraft speed (reference speed). When the speed of a wheel decreases below 0.87 times reference speed, brake release orders are given to maintain the wheel slip at that value.

The reference speed is determined by a BSCU from the longitudinal acceleration given by ADIRU 1 or ADIRU 3. If ADIRU 1 and ADIRU 3 are not valid, reference speed equals the maximum of the 4 main landing gear wheel speeds.

Deceleration is limited to 1.7m/s2. (BSCU = Brake & Steering Control Unit).

Figure 31 shows a simplified schematic ofthe Boeing 757 sydtem and'figures 32 and 33 show the systems in mole detail. Study the drawings, carefully' noting all the inputs/outputs to and fromthe - syste$ - knd the relationspif, and cokrol of the various valves. , --. 1 I _ _ / ,

---, '\ , \ rL'

The brakes are supplied with d o independetit hydraulic systems - left and right. When braking qnder pilot contrbl ithe normal! brake control val4e is used to supply

I pressurk fluid to the anti skid vhve. L.. /' , \ '

/ / . - I

--A -- - 1

Anti-skid Overview

The rate of wheel retardation is compared with the aircraft retardation rate (picked off from the IRS) and should the wheel start to slow too much (approaching a skid) then a signal will be sent to the anti-skid valve to reduce the pressure to prevent the wheel skidding. Once wheel speed has been regained then the pressure signal is re-instated. This process is carried out in a few millseconds.

The anti-skid can be switched off from the flight deck. When off, a warning light is ON in the flight deck.

The alternate system is similar to the main system and the supply is fed in via shuttle valves. These are pressure operated.

- 61 -

moodull lA-1312

RIGHT HYDRAULIC

Fig. 31 SIMPLIFIED SCHEMATIC OF THE B757 ANTISKID/ AUTOBRAKE SYSTEM

SYSTEM SUPPLY CONTROL SIGNALS TO AUTOBRAKE PRESSURE CONTROL VALVE

FEEDBACK SIGNALS

NORMAL BRAKE -)R -)R ANTl SKlD B AUTOBRAKE CONTROL COMPUTER CONTROL VALVE AUTOBRAKE

CONTROL

Autobrake Overview

PARKING BRAKE VALVE ----) MAIN & ALT HYD SYS ----)

28V DC ----) AUTOBRAKE SELECTOR -----)

WEIGHT SWITCH ----) THRUST LEVER POSITION -----)

SPOILER POSITION ----)

The pilot can select (prior to landing) 1 of 5 brake settings with deceleration rates from 4.5ft/sec2 to 1 lft/sec2. This allows for "feet-ofl" automatic operation of the brakes.

VALVE

AUTOBRAKE SHUTTLE VALVE

+ PRESSUREFEEDBACK LEFT HYDRAULIC

SYSTEM SUPPLY PRESSURE

To operate the autobrake has to:

* Have a no faults operational anti-skid system. * Have IRS input. * Be armed. * Have all throttles not advanced.

IRS ANTl SKlD CONTROL SIGNALS

rnoodull lA-1313

ANTI SKlD ONlOFF SWITCH - - - GEAR LOCKED DOWN -------)

BIT ----) SKID VALVE

TO FLIGHT DECK HYDRAULIC FUSE

DISPLAY COMPUTER SGU (EICAS)

-----) SIGNAL TRANSMISSION DIRECTION SHUTTLE VALV

FLUID FLOW DIRECTION * CONTROLLED PRESSURE

R = RETURN FLUID

SPEED SIGNALS LBRAKE I !

ANTl SKlD CONTROL SIGNALS

-k Have weight switches made for more the 0.2 sec. -k Have correct spoiler lever position. ~r Have wheel spin-up circuited activated. Activates when all

wheels are 60kt or greater and deactivates when average velocity drops below 30kt.

The autobrake supplies brake pressure to the brakes via the anti skid system entering the system via shuttle valves and the systems are tested using BITE.

The anti-skid/autobrake system consists of the following components:

1. Transducer (8) 2. Anti-skid module, normal system (2) 3. Anti-skid module, alternate system (2) 4. ANTISKID fault light on flight deck (1) 5. Anti-skid shuttle valve module (2)

-6.- . ANTISKID ON/OFF s w i t h o n flight deck (1) -_ . - 7 . Anri-skid jauiobdke corliroi dr+ii ivi i 02 j i j -\

8. Autobrake selector &witch on',(l) 1 \

9. Autobrake module (;I) I

10. Autobrake shuttle v/alve assembly (2) I I 1 ' I 1 1. AUTOBRAKES light! (1) 1' / 1

I 6' 1-1 .' - --\\ /

A N T I S ~ D ON/OFF switch I , ', '

I ' I I 1

/ 1 I I

'I \ 1 j I I

I I An ANTISKID O~('OFF switch provides 28,* dd power to the cbAtro1 unit when ON. / I

An -b&light d t h OFF legen& on-thes6itdh comes on tor- the pilots that the anti-skid-system has been- turned-offkwith the ANTISKID s&itch in the OFF position, the ANTISKID OFF advisory message appears on the EICAS display. The ANTISKID OFF message will inhibit the ANTISKID/AUTOBRK maintenance message unless ANTISKID/AUTOBRK has existed 10 seconds prior to the display of ANTISKID OFF.

ANTISKID Fault Light

An amber ANTISKID fault light illuminates to signal anti-skid system fault when any of the following faults exists. At the same time the ANTISKID advisory message appears on the EICAS display.

1. 28v/5v anti-skid control unit power supply out of tolerance. 2. Anti-skid transducer wiring open circuited or shorted circuited. 3. Anti-skid module (normal or alternate) valve wiring open or shorted

circuited, or valve driver failure. 4. Anti-skid control unit card failure. 5. Parking brake valve not fully open when the parking brake is

released.

Five anti-skid fail relays control ANTISKID light functions and the related EICAS message display. When the control unit lamp driver output signal is a ground (no fault in anti-skid system), the relay energises. The light extinguishes and no EICAS message appears. When the lamp driver output signal is an open circuit (fault in anti-skid system), the relay de-energises. The light illuminates and the related EICAS message appears. Four relays each control anti-skid failure indication for two wheels. One relay controls the alternate anti-skid failure indication.

The relays also control ANTISKID OFF light function and the ANTISKID OFF advisory message on the EICAS display. When the four relays controlling paired wheel fault indication are all de-energised (ANTISKID switch in OFF position), the ANTISKID OFF light illuminates and the ANTISKID OFF advisory message appears on the EICAS display.

Fig. 32 B757 ANTI SKID SYSTEM

Autobrake Selector Switch

The switch is a rotary, magnetic-latching seven position switch, or, on aircraft with RTO, an eight position switch. It performs the following functions:

* Provide 28v dc power to the anti-shd/autobrake control unit. * Selects 1, 2, 3, 4 or MAX AUTO aircraft deceleration level. * Arms or disarms the system. * Turns on or turn off the AUTOBRAKES light.

A positive detent between the DISARM and OFF positions ensures that when the switch moves to the DISARM position it does not overshoot to OFF and prevent the AUTOBRAKES light from illuminating.

Fig. 33 B767 AUTO BRAKE SYSTEM

AUTOBRAKES Light

The AUTOBRAKES light is an amber light located near the selector switch. The light, controlled by the selector switch, comes on when:

* The switch is a t DISARM position. * The switch is a t OFF position and the autobrake module solenoid

valve output pressure switch shows presence of high pressure. x The switch is a t 1, 2, 3 , 4 or MAX AUTO position and a system fault is

detected.

When the selector switch is at 1, 2, 3 , 4 or MAX AUTO, the light illuminates for a moment as the switch is moved through the DISARM. The light then goes out when the unit confirms that arming requirements are met. When the system disarms, the light illuminates until the switch is placed to OFF or the system is re-armed.

I - r

-

Anti-skid/Autobrake Control Unit

The control unit compares each wheel speed with the IRS ground speed for touchdown and hydroplane protection. A change in speed causes a change in control signal to increase or decrease hydraulic pressure to tkie brakes.

- - ,'

The unit, located in the main equipment centre, contains conFrol, BITE and display circuit cards and a front control/display panel with alphanumeric readout and BITE test switches.

The unit and-the circuit cards are LRUs. Data links in the unit provides means of communication between cards. Power of 28v dc to the card supplies the 26v dc and the regulated 5v dc source required for all logic circuits.

The main wheel card primary function is to control braking pressure and prevent wheel skid or lockup. Each wheel card provides skid protection for two wheels in tandem. Each wheel circuit shares a common power source line and is combined by a locked wheel cross over function. A driver circuit in each card provides a signal to the EICAS display via the EICAS computers when the anti-skid system fails.

The autobrake card performs all autobrake functions, including control, logic, interface and BITE. The card shares with the wheel cards for wheel speed and system test information. A driver circuit in the card provides signals to the AUTOBRAKES light and the EICAS computers when the autobrake system fails.

Both wheel cards and autobrake card contain self-test and status circuits. The circuits check the system for fault and provide status to the BITE card.

- 66 -

moodull lA-1317

The BITE card communicates with all other cards via the data link in the control unit. Its primary function is to monitor the operation of the four main wheel cards and the autobrake card.

The function includes analysing received data, examining analogue voltages, performing tests required and providing fault signals to the memory. Logic circuits in the card determine BITE test switch inputs required for test. The interface/display card contains circuits for BITE and ANTISKID light functions.

Anti-skid Module (Normal)

Two 4-valve anti-skid modules are used in the normal brake system. Each module contains four identical anti-skid valves, four hydraulic fuses, a shut-off valve, two

i inlet filters, a check valve and a restrictor. The module provides individual wheel

\ control-to-each main gear. Each-module_<s an LRU and therfuses,shutoff valve anci iniet rilrers are separate c b m p o n e n r i k ~ s . Tne vaives 'anci fiiters can be

\ removbd for inspection without disconnecfilig hydraulic lines. \ '1

I I I

1 I _ - _ . - \

1 L---

Two 2;vdlve anti-skid modules +e used&,this system. Each module'6ontains two I I identidalanti-skid, valves, two hjrdbaulic fuyes, one inlet filter,, a check valve in a housiag. The module provides laterally paired wheel control to each main gear. In genkrkl the,altkrnate 2-valve medule fuiictions similarly td t$e main 4-valve

-- -1 / module. ' ' - , -/ i, i I

Anti-skid Valve

Each valve in the normal or alternate module consists of two stages. The first stage (servo valve assembly) develops hydraulic pressure proportional to the input current. The second stage (slide and sleeve assembly) repeats this pressure at the lower level required for brake control.

The first stage servo valve assembly consists of a torque motor and a hydraulic circuit. The motor, an electromagnetic device, produces an armature deflection proportional to input current, which positions the armature between two nozzles, pressure and return.

The second stage slide and sleeve assembly consists primarily of a stop, a quill and the slide and sleeve. Slide position is controlled by first stage output pressure on one end and pilot's brake metered pressure on the other end.

The anti-skid valves in the normal hydraulic system operate in the unigain mode. System pressure at 3000psi is supplied to the first stage providing a constant pressure source by which the second stage is controlled. Through unigain operation, the valve is able to maintain consistent stopping performance throughout the brahng range, independent of applied brake pressure.

The anti-skid valves in the alternate hydraulic system operate in the multigain mode. Pilot's brake metered pressure is supplied to the first stage rather than system pressure. A s a result, the first stage output pressure is a function of both metered pressure and valve current. Multigain valves are used in the alternate system to simplify the system and reduce weight.

Hydraulic Fuses

The fuse prevents hydraulic fluid loss from the brake system if a line should rupture a t any point downstream of the fuse. The fuse automatically_shuts off all '

n ---- :r L L - ;:^I-..--^^ ^C n.-:r3 ---- :-,.. +Lwn ,-.- L ;;+ ;.-. rl\IJ, ;- PPlf-rocP++;nrT l l U W 11 LLlC V U I U I I I C ; Ul 11UlU p a r 3 o l l r f ; L ~ L L uu&&l 11. lo c v v 5 s ~ U L . I I I ~ \IUV~-LV.YUII L IVUCCLLA b

and can be reset within 5 seconds against a pressure differential of 18,30psi without the aid of reverse flow. I

Fuse resetting is done manually by rotating the reset knob. This equalises the pressure on both sides of the fuse and allows the fuse to reset. When the reset knob is released, a spring forces the knob back to normal position. A slit in the knob allows visual verification of the bypass valve position. i

Hydraulic Filters - -

Two module filters, one at the metered pressure inlet and the other at the system pressure inlet, to protect module components against debris ingress.

Shut-off Valve

A shut-off valve installed in the main module prevents system pressure from being supplied to the anti-skid valve without pilot metered pressure input of 260psi or more. When the metered pressure > 260psi, the valve ports fluid from the system pressure line to the anti-skid valve. When the metered pressure < 240psi, the valve switches and the flow is from the metered pressure line to the anti-skid valve. The shut-off valve contains a check valve to allow metered pressure supply to the first stage of the anti-skid valve if the shut-off valve closes.

rnoodull lA-1319

Check Valve

A module return line check valve provides free flow from the metered pressure port to the anti-skid valve via the shut-off valve, and prevents residual pressure in the brake return line from feeding back.

Restrictor

A 2-way restrictor is fitted in the system pressure line upstream of the shut-off valve. The restrictor limits hydraulic flow during certain failure conditions.

Normal Anti-skid Module Operation

The torque motor armature in the first stage of the anti-skid valve (servo valve assembly) sits itself between twopnozzles as a function of input current. One , -. nozzie;is,sAppiied with sysrem-prepsure,dcj rhe other is connected~o\reiurn. With rio command applied, the motor mobes',the annature agaihst thq 2eturn nozzle a n d control pressure equhs supply pressure. With full current supplied, the rnoto,r moves the armature against the pfessure nozzle 4 d lcontrol pressure equals rCturn pressure. For each ih-between value of input curkent there is a

/' chara~teristic armature positiod apddc&tro< pressure value. ~ k e - v a l ~ e ,varies pressure from a high of pilot metered pressure to a low of no pressure'(re1ease).

1 ' I I I '\ I, ' 1

The s2cohd stag; islide and sleeve assemblb) is a spool valve &ven by control I I pressure (frorn/{he first stage) 04 one end df !he spool and b r a e pressure on the

other.~When-ontrol pressure e-xceeds-the'pilotys metered pressQe, the spool r ' movesto-port metered pressuhedireetly-to'the brake. In the metering region, the

second stage spool works as a pressure follower such that the brake pressure equals the control pressure. The bias spring on one end of the spool holds the spool down to allow full-applied pressure to the brakes when the first stage is de- energised. A control orifice slows the second stage valve movement to control input. At the brake port an outlet orifice slows the application rate of the brakes in response to pressure input. When the metered pressure is removed, the return check valve allows free flow return from the brakes.

Alternate Anti-skid Module Operation

The alternate module functions the same as the normal module except that the pilot metered pressure is used. A characteristic of using metered pressure input to drive the first stage anti-skid valve is a slight decrease in output pressure.

The alternate anti-skid module requires no shut-off valve and restrictor

Anti-skid Shuttle Valves

The module contains four identical valves. Each valve is independent and pressure operated. Each valve shuttles pressure between normal and alternate systems. The valve consists of a n LRU valve assembly and an LRU filter. The valve has a manual override plug feature.

The basic three-way, two-position shuttle valve consists of a normal (input) port, an alternate (input) port and a brake (output) port. Under normal operation the normal port connects to the brake port. If the hydraulic system switches from normal to alternate, a detented slide in the valve moves to block the normal port. This slide shift allows fluid flow from the alternate port to the brake port. In this manner, one input port always connects to the brake port while the other is blocked.

In the event of shuttle valve failure, the slide plug on the face of the valve is removed. A, flight dispatch plug (a fly-away ground maintenance tool consisting of - ---I1 +L=, J ,A -l-.m\ ;c. ;-otollorl in i tn n1-e- Tho i n c t ~ l l ~ r l p ~ g fnrC& the a ~ L L L C U L LILLLUCCLU ~ I L A ~ , IlluLculbu rLu-yruvu. . rrrurruAu-r

in the valve to shift, thus blocking the normal port and opening the alternate port. This condition remains with the plug installed. The plug has no [moving parts and is equipped with a ring to allow an indicator tag to be tied to it whilst being used.

Anti-skid Wheel Speed Transducers

This is a speed sensing device fitted to each main gear wheel $.nd contains only one moving part, a rotor which rotates inside a fxed stator. The stator attaches to a support inside the main wheel axle. The rotor, using a four-arm dog rigidly attached onthe rotor shaft couples to the transducer drive in-the hub cap. The drive, consisting of a bellows-type coupling and related mounting hardware inside the hub cap, turns the rotor when the wheel rotates.

The stator comprises a permanent magnet, a 150 tooth soft carbon steel pole piece and a pickup coil. The magnet sets up a magnetic field around the coil, whereby the mating 150 toothed rotor produces dynamic discontinuities when rotated. Turning of the rotor provides field changes as the mating teeth come in and out of alignment to produce a series of voltage pulses (1 50 times per rev). The voltage, related to the speed of the wheel, provides the control unit with wheel speed data.

Autobrake Module

The autobrake module is connected to the normal brake lines and is located within the wheel well. The module contains an Electro Hydraulic pressure control Servo Valve (EHSV) (pressure control valve), an upstream three-way solenoid shutoff valve and two pressure switches, located one each at the outputs of the solenoid valve and the servo valve.

The module is a n LRU as are the valves and switches. Solenoid valve, pressure control valve and pressure switches can be replaced without removing the module from the aircraft.

The module develops brake pressure in response to selected deceleration for all required autobrake functions. The solenoid valve provides on-off control of hydraulic power to the valve module and the pressure control valve controls output pressure from the module as commanded by the control unit. Pressure switches on the module monitor the pressure outputs from the solenoid valve and the pressure control valve and provide the logic to the control unit.

Autobrake Module Operation

The solenoid valve is a 2-stage, 3-way operated shut-off valve. When de-energised,

(- the valve moves to the right or closed position. The mechanical spring acting on the spool-and the presence of ~upply_pre<sure (admitted via an intemal passage of the spool)-provides the closing force. \

1 --, \ -\ r ---.. \ I

I I I ' ', \

This force drives the piston to the 'right anp kolates supply pressure from the rest of the kn-nbdule. The module in trim ports butput flow to retuh.

I /' 1 1 L, 1 i I

Input of 28v dc to the solenoid yal~e-~a<ses supply pressure toibe-applied to the I i piston. This drives the shut-off lspool to &fied, open positiob and-ap6lies

pressure! to the rept of the module. The valbe pressure switch tpen closes to show high pkeksure (above 100Opsi). %the selectbriswitch is a t OFF and the solenoid

I I pressurcj switch shows high pressure, the AUTOBRAKES light will come on. , 7 -- ,' --_ l' ,I

I -' r- <--

The pressurecontrol valve consists-ofajef-pipe first stage, an-in-between pressure feedback stage and a slide-and-sleeve second stage. Supply fluid entering the pressure control valve second stage also feeds the jet pipe through a first stage filter. The jet pipe directs a jet of fluid from a nozzle into two ports. The change of kinetic energy of the jet into static pressure in the two ports provides the pressure required to drive the second stage.

A torque motor in the first-stage electrically controls jet pipe position and the amount of flow to two receiver ports. One ports the pressure to return and the other controls the pressure.

A feedback spring attached to the jet pipe on one end and the pressure-feedback spool on the other counters the input of the torque motor. The spring returns the jet pipe to its steady state position (for that particular pressure) when the commanded pressure is reached. The feedback spool moves and compresses a spring until the spring force equals the first-stage control pressure acting on the area of the feedback spool.

Brake pressure and the first-stage command pressure act on equal areas a t opposite ends of the second-stage spool which ports fluid either in or out of the brakes as required until brake pressure equals the first-stage command pressure. Without first-stage command pressure applied, a spring biases the second-stage to return. A 0.070-inch diameter orifice in the valve return port limits brake pressure release to ensure smooth brake release during autobrake disarm.

The pressure control valve pressure switch checks module output pressure. When brake application conditions are met and the commanded deceleration exceeds the actual aircraft deceleration by more than one foot per second squared for more than three seconds, the system shuts down and the AUTOBRAKES light comes on.

Autobrake Shuttle Valve Assembly

The assembly consists of a valve and a pressure-sensing syitch. The basic three- \

~ r * ~ r e &-&P~s t2 by either g ,,12,! c r , zihtghh~kp,:ypt~m-- T?X?A valve issemblies are located one each on left and right wheel well. ~ o t h valves and switches are LRUs. I

The valve has two input ports (one normal, one auto) and one'output port (brake). Under normal operation, the normal port connects to the brake port. ' y h e n the autobrake system applies pressure, a detented slide located i i the valve moves to block the normal port. This slide shift allows fluid flow from the auto port to the brake port.

The pressure switch, connected to the normal input port, checks pressure downstream of the normal brake metering valves. When manual-braking effort exceeds 750psi on either the left or right pedal, the switch opens to provide an input to the control unit to disarm the system.

THE ANTI-SKID SYSTEM

Circuit breakers provide 28v dc power through the ANTISKID ON/OFF switch to the control unit. Four main wheel cards in the unit compare IRS speed data inputs with wheel speed. This results in a control signal to the anti-skid valves, which limit hydraulic pressure to the brakes. Two shuttle valve modules shuttle pressure between normal and alternate systems.

The AUTOBK ANTISKID TEST IND 1 AND AUTOBK ANTISKID TEST IND 2 circuit breakers provide 28v dc power to the interfaceldisplay card for BITE functions. A data bus provides all the tie-ins between wheel cards, autobrake card, BITE card and the interface/display card. The control unit provides fault signal to the EICAS computers and the ANTISKID light.

- 72 -

moodull I A-1 323

Two IRS speed data inputs (L and R) provide aircraft speed signals to the four wheel cards in the control unit. If the left or right IRS input is not present, the captain and the first officer can place the IRS select switch to the ALTN position. This allows L or R IRS select relays to energise to provide the alternate IRS speed data from the centre IRS system.

The PSEU (Proximity Switch Electronics Unit) provides gear down signal to all wheel cards for anti-skid system operation.

Individual Wheel Control

The microprocessor in each wheel card provides control for the fore and aft wheel pair. The card receives and processes wheel speed signals, receives and decodes IRS data, provides valve driver signals and communicates with the test and fault

i inputs from the BITE card. A digital/analogue converter in the card provides the valv~command to drive the re,spective-valve for each wheel, .

Paired wheel Control 1 I I I

The ndrmal valve driver generates a secori$;signal for the altdrnate valvq driver. The s{cd,nd signal is OR'd with th<p&dxwheel signal from dhd-rnatihg wheel card. The higher signal of the two drive>th&,altemate valve $or-the-laterally paired wheel. , I 'i i I

I 1 , Valv,e Driver Logic

L--

The wheel speed compares the applied brake pressure in three modes. These are Proportional, Integral and Derivative. The Proportional mode involves applied brake pressure with proportional wheel speed as the brake pressure changes to maintain a deceleration rate short of a skid. The Integral mode checks the past performance of wheel speed. The Derivative mode checks rate of change of the wheel speed. The three modes provide the data input required to produce a driving signal to the normal and alternate valve drivers. An alternate brake selector valve determines which valve driver is used.

Anti-skid Locked Wheel Protection

Locked wheel protection, a secondary anti-skid function, prevents lockup of individual wheels during all braking above 25kt. The tandem pairing of lock wheels compares the wheel speed of paired wheels. It provides a full brake release signal to the anti-skid valve when the speed of the controlled wheel is less than 30% of the paired wheel.

Hydroplane/Touchdown Logic

Hydroplane protection provided to the aft wheels to protect against hydroplane- induced wheel lockups and also provides touchdown protection. The control unit compares the IRS speed data with the wheel speed to generate a full brake release signal.

The signal goes to the respective anti-shd valve when the speed of the controlled wheel is as least 50kt below the IRS ground speed. The hydroplane /touchdown protection requires valid IRS inputs and that landing gear is down and locked. Loss of IRS signal, however, does not affect other anti-skid functions. Hydroplane/ touchdown protection for the forward wheel is provided indirectly through locked wheel protection.

Gear Retract Braking Logic - - -

The gear 11p signal (when the left and right landing gears are not do-F and locked) inhibits the alternate valve drivers for about 12 seconds., This is to allow the alternate system brake pressure (left system pressure) to stop the wheels before gear is retracted.

Low Speed Dropout Logic

A low speed dropout a t 7kt inhibits the valve drivers in the wheel control circuit. Below the dropout speed, the anti-skid system provides no b r k e release signal.

Anti-skid Hydraulic Operation

Brake pressure input of 260psi or more to the normal anti-skid module opens the shut-off valve to allow system pressure supply to the anti-skid valves. On alternate system, pilot metered pressure is supplied to the valve. The valve varies the output pressure to the brakes by moving the slide valve between apply and release, using an electrical signal from the control unit to the servo valve torque motor. The pressure flows through the fuses to the shuttle valve module and then on to the brakes.

Wheel speed signal from the transducer provides the required data input to the control unit. The unit compares the wheel speed with the aircraft ground speed to provide a control signal to the normal and alternate valves. The valves limit the pressure to the brakes.

- 74 -

rnoodull l A-1 325

AUTOBRAKE SYSTEM

The AUTOBK ANTISKID TEST IND 1 and AUTOBK ANTISKID TEST IND 2 circuit breakers provide 28v dc power to the AUTOBRAKES selector switch to the control unit. An autobrake card in the unit compares IRS deceleration data inputs with the selected deceleration (encoded in the control unit). This results in a pressure command to the autobrake valve. The valve provides the pressure via the normal anti-skid valves to the brakes. Two autobrake shuttle valves shuttle pressure between the normal brake metered pressure output (manual) and the autobrake pressure output. When the metered pressure > 750psi, either the left or right shuttle valve pressure switch opens. This results in an open circuit signal to the control unit to disarm the system.

Both AUTOBK ANTISKID IND 1 and AUTOBK ANTISKID TEST IND 2 circuit breakers simultaneously provide 28v dc power for the BITE functions. The

(' control unit provides fault signal to the AUTOBRAKES light and the EICAS

.\, display-vial the EICAS computers -. ~

\ I \ , ', !' \ '\

) r--~

Input4 tb the control unit affecthd a u t o b r w functions incldde I R S ~ ? ~ ~ , air/gr+uhd signal and thrust 1ei.e; and sp+il$r handle positiob +igna14 A working anti-s$id system is required for lautobrakq oieration. , ,

I / 1 1 I ; : , , I L.---' ,, / 1 ' I 1 i

, ,

' / A switch; on the panel provides the-piLo~'s 6.Qoice for autobr* e f u-nct%ons. Prior to

/' landing, the pilot arms the syst4rq by p?hing the switch at o r MAX AUTO.! dt touchd~wn, the brakks apply ad~$o~atically. i ~ / / 1 1 I I i i Should A failure occur, the systfni disarm$ +d the selector ~&tch automatically movpsl t o ' ~ 1 - S ~ R M . At DISARY,' tficsystkm,teleases the aut~brake pressure, an A U T O B R P ~ K E S - - ~ ~ ~ ~ ~ comes on 'and-t-he-~~d~~ display will s h o w d e AUTOBRAKES advisory message. The pilot can turn off the AUTOBRAKES light and remove the message by placing the switch to OFF. The light will not go off and the message will remain if the solenoid valve on the autobrake module is faulty (solenoid valve output pressure switch shows high pressure).

Autobrake deceleration level inputs selected by the pilot are:

Autobrake Deceleration Selector Pressure Level

Switch At (ft/sec/sec)

1750 5.0

I I - . -

MAX AUTO 3000 11.0

Autobrake Arming Logic

The system arrns and latches with a magnetic latching switch within 100 milliseconds when the following conditions are met:

1. A decel level (1, 2, 3, 4 or MAX AUTO) has been selected. 2. No autobrake failures detected. 3. No anti-skid failure detected on the normal system except that

failures on a wheel whose indication has been deactivated will not prevent arming.

4. All thrust lever switches show not advanced when either airlground signal indicates ground mode.

5. IRS signal available. 6. Brake pressure switches (both left and right) show low pressure.

7. When conditions 1 to 6 are met, hydraulic pressure is metered to the brakes. Establishing an initial low-pressure level. The system ,

- then holds the brake pressure a t this level until the-pitch angle of the - zircrzft is r~di-~ced to a -~~rnx imate ly one degree,-as measured by the

IRS. A s the aircraft de-rotates through the one degrke refkrehce attitude, brake pressure is increased to achieve the deceleration value which corresponds to the chosen autobrake setting.

If aircraft de-rotation is delayed, thesystem will still command brakeapplication. However,'the system will pause ljefore transition to the selected deceleration rate. For the lower autobrake settings, the system will transition t o sdlected deceleration rate after about 5 seconds, independent of aircraft pitch attitude. With the higher autobrake settings, transition to the selected deceleration rate will commence when 8.0 seconds have elapsed from main gear touchdown. In either case, the preceding time delays will be overridden when the pitch-attitude reaches one degree.

The control unit provides an arm hold signal to the switch magnetic latch to keep the switch in the selected position when the above requirements are met. If the arm hold signal is not present, the switch moves to DISARM.

Autobrake Application Logic

The control commands brake pressure by opening the solenoid valve and modulating the control servo valve as required when:

1. Autobrake is armed. 2. All thrust lever switches are not advanced. 3. Aircraft on the ground, indicated by one airlground signal

continuously for 0.2 seconds. 4. Wheel spin-up circuit activated.

Loss of conditions 3 or 4 above after autobrake application causes auto-braking to be removed but not disarmed and the time delay resets.

Wheel Spin-Up Circuit & Brake Application

The spin-up circuit includes a detection circuit and a latch circuit. The detection circuit activates when the average velocity of all wheels is 60kt or above and deactivates when the average velocity drops below 30kt. The latch circuit latches three seconds after the aircraft is on the ground and the detection circuit is deactivated. The latch circuit resets (unlatches) when a ground-air transition takes place or the autobrake system is turned off or disarmed.

Upon initiation of autobrake control, the control commands an initial brake application to provide brake application.'

r -- - -\ . -- -\

~ u i d b r a k e rjisarm h g i c \

-.\ ' - -. i I \ I I

The adtobrake removes the power from the sblenoid and control1 servo:vqve driver$ and from the selector sd tdh latch &hen: I

, / I I -- " / ' I 1. 1 System is selected off .r__-- < , ___ / '

2 . Either left or right rhetered p?ys+ure switch indicbtFs-pres<ure 1 1 (manual brake application) of 75ypsi. 1 3. ~ n d t h r u s t lever switch indicates;advanced on the ground, except that i any thrust lever s q t c h indicating advanced for Jp 3 seconds after

- - -t&chdown will not idibarm-tlie,dystern. ' '7 4r -Autobrake fai1uredeteeted;-iniluding failure to~applyl pressure

(indicated by the pressure control switch) when application conditions are met and the commanded deceleration exceeds the actual aircraft deceleration by more than one foot per second squared for more than three seconds.

5. Anti-skid failure on normal system detected except for failures on a wheel whose indication has been deactivated by the control unit selector switch.

6. Spoilers stowed after having been deployed on the ground. 7. IRS signal not present or faulty.

Operation

The autobrake solenoid valve provides on-off control of normal hydraulic system power to the autobrake valve module. When a valve driver signal from the control unit is received, the valve opens and admits system pressure to the pressure control valve. The servo valve varies the output pressure from the module as commanded by the control unit to maintain the selected deceleration.

rnoodull l A-1328

The solenoid valve pressure switch senses the solenoid valve position (open or close). When the valve opens and admits system pressure to the module, the switch closes to show high (> 750psi) pressure and provides an appropriate signal to the control unit. The control valve pressure switch senses the output pressure and provides an appropriate signal to the control unit for autobrake control.

Brake pressure output from the autobrake module positions the two autobrake shuttle valves to allow fluid flow via the normal anti-skid modules to the brakes. Two pressure switches on the metered pressure port of the shuttle valves monitor manual brake pressure downstream of the brake metering valves. When manual- braking effort exceeds 750psi on either the left or right brake pedal, the switch opens and provides a signal to the control unit to disarm the system.

SYSTEMS TESTING

Extensive BITE on the control unit provides for anti-skid and autobrake system ( - testing I?,stk ir, fl,ight ~ q d on t h e gci~nd. The RITE dcE! n r n ~ n ' A e c r* - - ---- S ~ C ~ ~ I T , t,rcllhle shooting at LRU level. If the unit is the failed LRU, it isolates the fault to the particular circuit card with in the unit.

A BITE control and display panel and an instruction placard are on the front face of the unit. The panel face consists of the following:

-- -

J; One alphanumeric display - provides readout and identifies the failed channel.

* One BRAKE TESTIDISABLE rotary switch - allows selection of 1 to 8, normal system operation test position and an A/B brake test position. In addition it disables fault indication on a selected-wheel. On aircraft dispatched with one wheel deactivated, the switch inhibits the inputs to the related channel, thus preventing the display of anti-skid EICAS messages and the illumination of the ANTISKID light.

-k One ENABLE/VERIFY momentary pushbutton switch - used with the VERIFY switch to enable a system test or brake operational test.

* One VERIFY momentary pushbutton switch - performs a complete system test or brake operational test when used with the ENABLE /VERIFY switch.

* One PRESSITEST momentary, 3-position (PRESS/TEST-OFF-BIT) toggle switch - performs lamp test (switch in PRESSITEST position and recalls faults (switch in BIT position). The lamp test includes all segments of display, ANTISKID light and AUTOBRAKES light.

* One RESET pushbutton - clears memory of stored fault information.

Three levels of BITE test are provided. These are system test, continuous monitor test and brake operational test.

System Test

With both the ENABLE/VERIFY and VERIFY switches pressed and the BRAKE TEST rotary switch in the NORM position, the control unit performs a complete system electrical test and sends its results to the BITE. When a fault is detected the test identifies the failed LRU on the display. When the VERIFY switch is pressed and released again, subsequent faults will be shown until TEST END is shown, indicating no remaining faults.

Continuous Monitor Test

The control unit checks the valves and transducers for continuity and the control circuit for proper functions.

I Y The control unit checks the anti-skid valves (normal and alternate) by monitoring

a small-voltage, which it applies-to-the-valye torque motor. This voltage causes the anti-skid//Gdve torque motor %a~ure-to,be \ I biased slightly; away-from'the hydradli& return port. \

, I ' I I

I

I ! I I

I i I 1

With full' system pressure availble, the biaqlcan be fully blocked. This allows full prbsbure to't~e brakes.

I ---<\ With ohl$ accumulator pressure: available,,,pl;essure be suf~cient to overcome the bids h d preskuke degradnk brakel&forrnance. 1 1 1 , I

I / / / / 1 / I I

~hereforkm"thout right or ce?tke-iystei6 prbssure, the BITE rnbnitor function of the anti-~~tIEGdves is suspended,---' - _-A

The test stores the following detected faults (continuous and intermittent):

* Failed transducers. * Control circuit failure. * Faulty valves. * Parking brake control/parking brake valve disagreement. * Wheel deactivation. ~r Loss of IRS signal.

Faults are stored in non-volatile memory for readout during ground maintenance. The memory can store data for at least 250 hours without power.

Memory recall for fault readout requires placing the PRESS/TEST switch at BIT and releasing; this shows the first fault. Subsequent BIT selections display any remaining faults until TEST END is shown. The RESET switch clears memory of all fault data.

Brake Operation Test

The test checks the skid release feature and allows a visual brake check. The visual check is on a single wheel for the normal system and on a wheel pair for the alternate system. The test starts with the BRAKE TEST switch selected to one wheel and the ENABLEIVERIFY switch pressed. Pressing and releasing the VERIFY switch causes a one-time brake release and re-application for the selected wheel (normal hydraulics) or wheel pair (alternate hydraulics).

With BRAKE TEST switch at NORM and ENABLEIVERIFY switch pressed, pressing the VERIFY switch performs an anti-skid system check. The display reads TEST END when:

* Transducers, valve circuit, associated aircraft wiring and power supply voltage are valid.

* Parking brake control and parking brake valve are in agreement. * - Gear retract braking test (with landing gear lever in OFF)_is valid. I ,

* C=ds i~ i ~ ~ i t z r ~ nn~ratincr -r--I'- nnrrnnllv --------- - -

J '

With selector switch a t 1, 2, 3, 4 or MAX AUTO, BRAKE TEST s h c h at A/B and ENABLE/VERIFY switch pressed, pressing the VERIFY switch performs an autobrake system check. The display reads TEST END when: '

I

a) ' Airlground sensing shows aircraft in ground mode.---- *

b) Brake pressure meets the selected deceleration leiel. c) Thrust levers are in retarded position. I

BITE Logic - -

The communication port in each wheel card provides the tie-in between cards. A BITE status circuit processes the incoming signal in and out of the self-test circuits and provides a BITE status output to the driver circuit. A second BITE status output is OR'd with a failure signal input from other wheel cards to provide a driving signal to the driver circuit.

Anti-skid/Autobrake BITE and Fault Annunciation

Circuit breakers provide power to the autobrake, BITE and interface/display cards in the control unit.

The interfaceldisplay card receives front panel rotary switch position discrete signals to the panel switch input port. The display card also receives the autobrake discrete signals, the system fault signals, through the BITE/display port.

rnoodull l A-1 331

Display Card Signal Processing

The display card, through data management, processes the discrete and fault signal inputs into the microprocessor.

The rotary switch position discrete signals allow the test initiation circuit to generate request commands. The commands are appropriate to its mode of operation (test level) to the wheel cards and autobrake card. On receipt, each card responds with the status information requested.

When a fault is detected, the display card stores the fault in the memory and provides driving signals to the display driver and the fault light driver. The display driver enables the illumination of the display on the control unit (upper half of the display comes on immediately and the lower half follows after three seconds). The fault light driver receives the signal from the fault memory and turns on the ANTISKID light.

I -- - - : - - ~

I - - ------..\ - ..~ - +,, /- . - \\, \ . '\

\ i \ \

~ n t i - s h d Wheel Card Signals \ \, 's, i , 'I, j,,

~.

i ' 1 j i i I

' i I i

The B$E self-test circuit in each %heel c$djProvides a corndleie electkick self- / test to t h e anti-skid system. Th< tesl-includes component check for continuity

and i4pedance, driver check foi circuitfdlqre and a brake rkldase-t&s>; Results of thes!e tests are passed on to the BITE c\ard\via the data link

I i ,A: . 1 , I ' $ 1

', ', : I The BITE status,;t,est circuit in ehch wheel k d d provides a rndnitoring test on

/ completibn of t h e self-test. The jebt monl;tbr9 the following &dl provides its i _--/ I , , i-l ' 1 , results to the BITE card: .- ,-I _.I \.-

- - -p - - - L," L

* Gear position switch - up/down. ~r Transducer interface - operable/non-operable digital check sum -

active check with memory. ~r Wheel velocity - transducer continuity and impedance. * Valve current bias - electrical response to valves. J; Failure light drivers - continuous operation of light drivers. ~r Digital parity - check digital work parity for correctness.

The analogue signal inputs from the wheel card to the BITE card are the wheel speed transducer voltages, the 26v dc and 5v dc regulated power source voltages and valve voltages.

Autobrake Card Signals

The BITE self-test circuit in the autobrake card provides a compete electrical self- test to the autobrake system (aircraft in the air with the AUTOBRAKES selector switch a t 1, 2, 3 , 4 or MAX AUTO). The result of the test is passed on to the BITE card via the data link.

The BITE status test circuit in the autobrake card provides a monitoring test on completion of the self-test. The test monitors the following and provides its results to the BITE card:

j: Deceleration selection - a deceleration selected. * Solenoid and servo valves - continuity and impedance. * Failure light drivers - continuous operation of light drivers. * Digital parity - check digital word parity for correctness. x Digital checksum - active location check. * Throttle switch - continuity. * Airlground sensing - continuity.

BITE Card Signal Processing

The multiplexer in the BITE card selects the analogue voltages to be monitored. A continuous monitor test circuit receives and processes the voltage signals (valves, { transducers, 26v dc and 5c dc regulated power sources and faulfiemory power), the digital memory data and the system status input. The test circuit provides a time delay of all voltage failure indications as required by the logic circuit. The data management in the BITE card microprocessor allows a bkffered flow of signals between the BITE card and all other cards.

--

When the ENABLEIVERIFY switch and the VERIFY switch on the-control unit are pressed, the BITE card starts a system test (BRAKE TEST switch at NORM) or an autobrake test (BRAKE TEST switch at A / B and AUTOBRAKES 'selector switch in any one deceleration position). Failure detected during test is recorded in the memory and shown on the display. Successful tests feed no fiult signal to the BITE card and the system is operable.

Control - Anti-skid System

The system turns on when the switch is in the ON position. The system turns off when the switch is in the OFF position. Anti-skid system off is shown by the illuminations of an amber light on the switch, ANTISKID light on P5 and the ANTISKID OFF advisory message on EICAS display.

When normal (right) hydraulic system power is removed, the alternate (left) hydraulic system activates automatically. The system provides hydraulic power to the alternate anti-skid system.

The front face of the control unit contains BITE switches. These switches are used for checking out transducer, valve and skid circuits on each wheel. A complete anti-skid system test can also be performed with the BITE switches.

moodull l A-1333

Control - Autobrake System

Arm the autobrake system during landing approach:

> J; Place the AUTOBRAKES selector switch to 1, 2, 3, 4 or MAX AUTO

position to turn on system. * Observe that the AUTOBRAKES light goes out and the switch remains

in level selected. This shows the autobrake system is armed. * Depress either or both brake pedals. * Observe that the AUTOBRAKES light comes on indicating the taking

over of autobrakes braking by manual braking. The EICAS display shows the AUTOBRAKE advisory message. The selector switch automatically trips to DISARM.

* Place the selector switch to OFF. * Observe that the AUTOBRAKE light goes out. The EICAS shows no

AUTOBRAKE message. The autobrake system turns off. - - . . - - - .. . .- ---. - - . . - -.

-\\

- --. , - L.

\

I i -.\. ',, , , ', i ! , 3 I \%\ \, ,

' I , ,

i j i I ,,,,,,,,,,,,,,.,,+ \ ' , , I ; I I I * j ~ , ,

j , I

i 8

; 1 , i i

, ,

! ' i , : I : -.' /

,' 1

: I ' - .----‘ i I 7--. '-.- I \, ".; - /'

i 8 ,

1 ! !> I ! '\$ I '1: ;

, . ~ ~ i ! : i , I

1 1 , .

i ! i ! I I

, , i 1 i !

/ i' , , I j ; j , , , / 1 /

I ! !

! L.-/' ! 1. ..-' i I I

,-~, , .

r--' / i '. -

- 83 -

moodull lA-1334

LBP /

Dee 03 - Feb 04 - April 04

Addendums module 11A book LIGHTING SYSTEMS pending amendment action in response to student feedback after taking the CAA examinations.

*** If there is a partial failure of the passenger emergency lighting system the aircraft can be dispatched provided the passenger compliment is reduced to that number that can be carried in that part of the aircraft with a serviceable emergency lighting system. Check the MEL for the specific aircraft. The maximum number of emergency lights that can be out is 25% (www2 .faa. gov/certification/ aircraft).

*** A similar regulation applies to inoperable passenger exits. The passenger compliment is reduced and passengers are not seated near that exit. Again the MEL is consulted.

*** Not more than 25% of floor path lighting may become inoperative in the event of the fuselage splitting in half in a crash landing. Was in AN56.

***** - - . -

', --

NOTE: I t is possible that some of the above statements may not be too meaningful when read out of context, so it is suggested that the appropriate book/subject be read first then'the information above be checked against that

CONTENTS

Page Aircraft lighting 1 Incandescent lamps 2 Halogen lamps 3 The fluorescent tube 4 Light emitting diode 5 Flight deck lighting 7

Integral instrument lights 8 Pillar lamps 8 Ligh tplates 9 Beta lights 10 Floodlighting 10 Utility lamps 13

(: Control of flight deck lighting 13

Passenger cabin and cargo ligh'ting - --

- - - 18

P,assenger area lighting 18 Control of fluorescent tubes 20 Reading lights 2 1 Attendant call 23 Cargo compartment lighting

- -- 25

Other area lighting 2 5 Exterior lighting ' 26

Legal requirements 26 Navigation (position) lights 27 Anti-collision lights - - 30 Xenon flash tubes --- - - - 31 Landing and taxi lights 33 Other external lights 35

Emergency lighting 36 Floor proximity lights 37 Emergency exit signs 39 Slide and door illumination 40 Lighting control 40 Testing 44

Answers to self assessment questions 45

moodull l A-1 336

AIRCRAFT LIGHTING

Aircraft lighting may be divided into the following groups:

Flight Compartment Lights. Flight compartment illumination provides area illumination (flood lighting), control panel lighting and indicator lights - including dimming and testing. The requirements are laid down in EASA CS25.

Passenger Compartment Lights. Passenger compartment lighting provides illumination of the passenger cabin, passenger signs, crew call lights, reading lights, toilet and galley areas etc. A requirement in the A N 0 is that all passenger compartments must be illuminated.

Cargo and Service Compartment Lights. Cargo and service compartment lighting provides illumination of cargo compartments, service areas and cargo door areas for ground operations and maintenance.

Exterior Lights. Exterior lights provide illumination of the ground during landing --- J L--F:---=---c:,,, ,,A ,-I,, ---- r~ - - : - : t - 1 - 1- n:-LC - 1-2-1- A :-- LI- - cu~u LMI U ~ C L ~ L L U ~ L D allu L l L a h r ; ~ l l c a l L ~ l a l L y 1 - 1 ~ 1 ~ 111 I I I ~ I I L , a3 l ~ l u uuwl l ILL LIIL

ANO. provision is also made for thelllumination of tailplane logos and the requirement has to be met for ice inspection spot-lights.

I I 1 I

Emergency Lights. Emergency lights provide interior and exterior illumination of exits and exit paths during emergency-evacuation. Also fitted to emergency

1 I - -

escape chutes. , \ \ _-'

Fig. 1 AIRCRAFT LIGHTING A

F L I G H T COMPARTHENT LIGHTING

Before we look at lighting systems we shall spend a little time on lamps themselves to include:

* The incandescent lamp * The halogen lamp * The fluorescent larnp * LEDs

PASSENGER CONPARTHENT LfGHTlHG

CARGO AND SERVICE CONPIRTHENT L I G H T I N G

i

EXTERIOR LIGHTING

EMERGENCY L16HTING

The Incandescent Lamp

Thomas Alva Edison, (American physicist 1847 - 193 I), invented the incandescent light bulb (amongst other things) in the 1870s. Figure 2 shows a sectional view of a typical incandescent lamp.

The operation of the bulb is essentially very simple. An electric current is passed through a thin tungsten filament (diameter about 0.Olmm or 0.00004"). This gives the electrons a higher energy state and increased motion. This increased motion causes the conductor to heat up.

Electrons only remain at a higher energy level for a short period of time. As they return to a lower energy level, the excess energy is shed in the form of photons of light. This gives the appearance of making the filament glow 'white-hot'. The filament is carried on glass mounts to prevent temperature transfer. The ends of the filament are welded to thicker support wires that hold the filament in place and provide a current path.

~ n e b ~ ;

Fig. 2 INCANDESCENT LAMP

The filament is formed into small coils, mounted onto glass rods and the lot encapsulated in a glass cover. The glass cover is then filled with the gas as required and cemented to the base.

SAQ 1

In the early days of development the filament was placed in air, but would quickly oxidize and burn out. How is this prevented in a modern lamp?

rnoodull l A-I 338

The base supports the wires, has contacts located in the base and provides support for the assembly in the light fitting. Methods of fixing include screw and bayonet connections.

The Halogen Lamp

The incandescent lamp is only about 5% efficient with most of its energy being lost as heat and invisible radiation. A more efficient option is the halogen l a p .

The halogens are a name given to a group of elements that have a valence orbit with one electron missing - fluorine, chlorine, bromine, and iodine. On the periodic table they fall vertically underneath each other, one space to the left of the inert gases. What that means is that in any reaction they want to gain one electron.

Metals are the opposite; they have a valence orbit with one or two electrons and as such want to give u p a n electron in any reaction. When halogens and metals are brought together in the right conditions they will react - one giving and one

', t&ir,g, =d XI ionic S c ~ d is fc-ed. - , -- - . --- . -- -

'>

Lamps of this type have a quartz envelope and are halogen filled. A commonly used fill is iodine. During use the tungsten filament vaporise;~ slowly and under normal circumstances deposits itself on the envelope inner face, thus reducing the light output. The halogen fill keeps theJenvelope clean by chemichl reaction with the deposited tungsten vapour~The reaction actually goes% stage further with the vaporised tungsten being re-deposited on the filament, thus extending its life.

I

The halogen lamp tends to give a harsh bright light. This is excellent for outside work but is undesirable for interior-u&. _ - .

(Some car manufacturers have started using quartz iodine lamps as headlights. The same is true on aircraft with landing lights etc).

Care needs to be taken when replacing this particular type of lamp as the natural oils from the skin will be deposited on the outside of the envelope. This will, during normal operation of the lamp, sink into the envelope and seriously shorten the life of the lamp. Always wear suitable gloves when handling the lamp.

Note. Most lamps get very hot when on. When changing a lamp always switch off some minutes before attempting to complete the task to allow the lamp and housing to cool. If there is no time to allow it to cool then wear protective gloves.

- 3 -

moodull lA-1339

The Fluorescent Tube

Figure 3 shows a schematic of a basic fluorescent tube. The tube is filled with mercury vapour a t a low pressure and sealed at both ends with a cathode and an anode. The inside surface of the tube is coated with phosphorescent metallic salts, such as zinc sulphide or zinc silicate. Note that it will require an alternating current unlike the incandescent lamp which will work with ac and dc.

Electrons are emitted from the incandescent cathode (not bright enough to illuminate anything - in fact some tubes are cold anyway). These electrons have tremendous energy and will aim for the anode at the other end of the tube. Movement will be from the cathode to the anode (dc), but alternating when the supply is ac. During their journey they will collide with the mercury vapour atoms in the tube. These collisions cause radiation to be emitted, which is mostly all in the invisible spectrum (ultra violet).

The visible portion of the radiation is a pale colour in the blue green range. Th- ultra violet (invisible) radiation will strike the phosphorescent coating on the insidyof-t3e tube and cause it t o emitradiation with a longer wavelength and more i 'p~rtant ly , a wavelength tkaPi3in the visible spectrum. During the collisi?n other electrons are emitted and move on towards the anode, colliding many times before getting there. I

I ' ,

1

- .-- I

---- SAQ 21 1 1

I I

b I I

If an ohmmeter was placed across an incandescent lamp and then placed across a fluore$cent filament, what results would you expect to see?

/ 1 \

- -

- - - -

Providing the current is already flowing, and the collisions are taking place electrons will be moving from/to the cathodelanode and as such the tube (once started) provides a very low resistance. So how do we initially get things started7

This is where the choke (sometimes called a starter) and the auxiliary lamp come in. The choke is an inductor which prevents dangerously high voltages; it also acts in the starting cycle.

When current is first switched on the glow lamp lights due to the fact it is in parallel and the bimetallic contacts are a t this moment open. The bimetallic contacts now close and allow the full current to the cathode which makes it incandescent. The bimetallic strip now cools and breaks; this breaks the action of the choke and causes a voltage spike that forces electrons to be emitted from the hot cathode.

The cycle has begun and current flows in the main tube and no current flows to the auxiliary lamp and so the bimetallic contacts remain parted.

Tube "-1 I

f 4

Cathode 1 w

Glow lamp V BI metalic

contact - electrons

Ultra violet radlallon Fluorescent coatlna

Glass envelope

Fig. 3 FL';'BRE8cENT ,+rvf B - GPE;aTiGIi --

- - - -

\ \

Careful selection of the fluorescent material can make the light glow in virtually any colour. The most common is as described and this produces a moderately white light. As the light is caused by radiation from collision ,rather than a n

-- -

energy change (in the case of the ifL%ndescknt lamp) it is very efficjent once started. It produces little heat m d so doesn't add to cabin heai sink.

Due to the alternating current, the lamp can have a moderate stroboscopic effect and b-ecause of this and the fact-that is produces a harsh light, its use on aircraft tends to be limited to the cabin spaces, where they>re.6ed extensively. They require a ballast transformer to raise the voltage sufficiently to provide the ionising effect required to excite the tube. This transformer is also used to control light intensity, as we shall see later.

Fault diagnosis of a faulty tube is difficult. The resistance cannot be measured, as when the tube is not working its resistance is very high. Often, the only course of action is to replace the tube, if that fails a voltage check etc will be required.

Light Emitting Diode (LED)

If you have already studied modules 4 and 5 then this should be mostly revision, if you haven't then this might all be new to you and you will require a bit more study time.

LED'S operate on the principle that under forward bias conditions a p-n junction diode exhibits special properties.

Electrons are easily driven from the n-type material into the p-type material by the electric field applied. The reduced depletion area that exists allows the electrons to travel quickly from one to the other. Upon entering the p-type material, but still close to the depletion area, the electrons meet a plentiful supply of 'holes7 that allows rapid combination. The electron is now returned to a lowered energy level and the excess energy is shed in the form of a photon of light energy.

Careful selection of materials determines where the emitted radiation falls in the spectrum. Gallium arsenic phosphide is a common semiconductor used to provide an orange indication. This colour can be changed with the use of a filter or different material construction.

That is a brief description of an LED and should you require more detailed information you should refer to the EASA module 4 and 5 study books.

Use of LED'S is somewhat restricted at present as they do not provide sufficient light for illumination purposes, but they are used for indication. They are used in seven segment displays and are used for fibre optic transmission.

I - --7 , -

- i They are; useful indicators as theycareefficient, require no additional relay or control devices, and they give off iery little heat. Another advantage is they last a long t ihe , unless a too higher voltage is placed across them.

1 ! I I I

1 - --

I I

LENS , 1 I -

I TRANSLUCENT I PLASTIC CASE 1

_,

- -- FLAT

CATHODE

CATHODE ANODE LEADS

Fig. 4 LIGHT EMITTING DIODE

We shall leave it there for the time being. There are other forms of light devices and we shall be briefly discussing those as they come up.

Try the following exercises. As usual the answers will be found in the text.

1. Explain briefly how and incandescent lamp works. 2. Why is a fluorescent tube more efficient that an incandescent lamp? 3. Where might you find the halogen type of lighting and why? 4. What use might a n aircraft designer have for a LED? 5. Explain briefly the operation of an LED.

- 6 -

rnoodull l A-1342

Suggested Practical Activity (if access and equipment allow):

With the lamp removed from its socket carry out a resistance check of a known good incandescent lamp and note the reading. It should be typically about 5R. Check the resistance of a 'dead' (open circuited) lamp and note the reading.

In a dark room switch on a domestic fluorescent light. Watch for the auxiliary glow lamp illuminating just prior to the main tube flash over.

FLIGHT DECK LIGHTING

The requirements for aircraft lighting are laid down primarily in the A N 0 and is further expanded on in CS25.

Adequate lighting must be provided so the crew can see all the controls and the instruments and the markings near to them. Anytime the aircraft is or is likely to enter Instrument Flight Rules (IFR) the instruments will require illumination. This illumination will need to cover the possibility of operations in dull daylight

\ ~1 L1ll uug;i dusk tu coiiipiete "1- - -- --L-:-- - -- A "---A ~ r r : - - -- --- --- - 1

1 L ~ C L+~LCLLH ~ L L U 111 3~ u111~t; i - i~~ay wib~l tu - --

alter illumination levels. All this needs to be catered for. ' I

The lighting has to be positioned so that it illuminates the instruments or panels clearly but does not produce glareifor the flight crew.'

-

-.

Flight deck lighting includes:

* Floodlights. Used for general lighting of the area. The main instrument panel floodlights are controlled from switches on the lighting panels. These lights may be fluorescent tubes or ordinary

- l amps . Some flood lightslike-glareshield and aisle floodlights are controlled from rheostats.

* Integral Lighting. Lighting for all the instrument and circuit breaker panels is controlled from the lighting panel. The bulbs used for faceplate illumination are usually soldered onto a circuit strip attached to the rear of the instrument faceplate.

~r Miscellaneous Lights. Map lights illuminate the pilot's lap area. Chart lights illuminate the pilot's letdown chart holders and utility lights that are moveable are for miscellaneous use.

* Override Light Switch. The override light switch can be used as a means of turning on all of the instrument panel floodlights from one switch.

* Standby Lighting. With only standby power available certain critical lighting circuits revert to standby power sources to provide emergency lighting.

~r Master Dim and Test. The master dim and test provides night annunciation light dimming and a convenient 'one switch' test of all appropriate lamp annunciators.

We shall spend some time dealing with lighting based mostly on Boeing aircraft but the principals are similar for all large commercial aeroplanes.

Integral Instrument Lights

Each instrument has its own integral lamp with the wires run to the back connector of the instrument casing. A plug will then connect the larnp to the aircraft lighting system. The circuit will include a common switch, fuse or C/B and a dc bus bar.

The filament lamp is so positioned that it shines onto the face of the instrument, yet doesn't show any glare.

A drawback wich chis type of iighting is that a i m p change requires ihe instr&nl& to be removed a d e i t h e r dismantled or replaced. This means additiqn?l functional tests for theinstrument system, and if the bulb failed during alnight flight, the instrument would be difficult to read. It has limited use these days and any new aircr9ft will not normally have this type fitted.

I - -

(cathod4 Ray Tubes [CRTs] or &Ed Crystal Displays (LCDs) require --- - no addition41 illumination, though they do have there own brilliance controls.)

I i I

I , I

Pillar Lamps ,(fiwre 5) - .. - -

An incandescent filament lamp is housed inside a small pillar located near the instrument/s and provided with a power supply. The cap on the pillar prevents any glare and the slot on the side deflects the light down onto the instrument or panel.

Fig. 5 TYPICAL PILLAR LIGHTS

- 8 -

rnoodull l A-1 344

Often two pillar lamps will be fitted over the primary flight instruments, thus providing a degree of redundancy. Replacement of the bulb is simple enough with the cap being held in place with a spring collar. Once this is removed the sub-miniature bulb can be removed. The cap alignment is crucial as this determines the light effect; the cap is keyed to the pillar.

An improvement of the pillar lamp came when two were located together on the same housing. This made for easier maintenance. Sometimes called a bridge lamp.

Trans-illuminated Panels (Lightplates)

This type of illumination is common, particularly on 'glass' cockpit type aircraft. I t relies on the principle of reflection and refraction of light between two polished surfaces.

With reference to figure 6, a light source, typically an incandescent bulb, has the light directed into the end of a glass sheet. This light will travel down the

\ -13-E. by-reflectiaE aff the po!ic,h~,d- %ides. ,AAt 2 ~ x 1 n ~ n c l-trhorl ~ r - ~ c \ - n - \ r r the b A U " V J bay" \" L"""" - bU" 1 "tL "'" reflecting surface the light will be refrairtedout of the plate and-become visible.

I

I I

I Scratch location

Pollshed sqrfa'ce I

I

1

Fig. 6 LIGHT REFLECTION PRINCIPLE ,

The lightplate used on most modeyn airliners is an application of this principle. Referring to figure 7 the incandescent lamp is placed in the middle of the panel. The panel in this case is Perspex (a transparent plastic), with a coating on the outer side which forms the front of the panel. With the 5Vac supply on, the lamp will reflect light down the inside of the Perspex which will be refracted out a t any coating discontinuity. These discontinuities form the light pattern on the control panel.

There is a limit on the distance that a small incandescent lamp can work to, so to provide good panel illumination on larger panels a lamp has to be placed every six to ten inches (15mm to 254mm). The lamps are identical to the small filaments of the pillar type lamps and are similar to replace.

- 9 -

moodull l A-1 345

White coating (allows Plastic laver reflected-and

refracted light) \

(blocks light)

I Rubber Threaded Cover

Shroud

instrument instrument Panel Standard 'P' bulb

light source

Fig. 7 TYTPICAL LIGHT PLATE

With an increase in bulb life, it has become possible to insert the lamps into the panel itself. There are more lamps fitted than is actually necessary to provide the correct level of luminescence and this allows for some failures. In this case if there is a problem with the lightplate, the whole panel is replaced. Again with modern-aircraft this is not so ,demanding on maintenance due-to the BITE that most-systems have, making such a-replacement possible in,a matter of minutes.

I I \

A further development i s to replace the incandescent bulbs with LED's. LED's as you know have an almost infinite! life and 9s such the problems of illumination

I .?

source fylure is largely eliminated. - -1

I I - , Y \ /

I 1

~ a s e o u s l ~ r i t i u m f ~ i ~ h t Sources /(Blacklights) ' I

1 1 I

Not veryfcommon but you may come across,them. Unlike other lights on the flight deck therelare no ON/OFF switches, these lights are on all the time.

- - - - -- - -- - - - - - -

They work on a principle of producing Beta radiation from the tritium and fluoresing, in a way not too dissimilar to the principle a fluorescent tube. They tend to be used (if a t all) in emergency light packs and to show a route to a torch for example. Also used on some avionics equipment.

They need care and attention when replacing. A s a general rule they are not dangerous providing that the glass is intact, but if this should be broken clean u p all glass carefully ensuring that no dust is inhaled. Follow the procedures laid down by your company and if you are cut with the glass (the same as with fluorescent tubes) you should seek medical attention. They are, as their name implies, mildly radioactive.

Floodlighting

Floodlighting is a general term used for the rest of the lights in the flight deck, however these can also be used to illuminate instrument panels.

I t is not uncommon for strategically placed floodlights (under the glare-shield for example) to be used to illuminate the instrument panels. Sometimes lights placed in the ceiling area are directed down to illuminate the centre console. Figure 8 shows a typical floodlighting arrangement and figure 9 shows a typical light controls layout. You should note the different types of lights and their location. All these lights are of the incandescent filament type, although a t least one manufacturer uses fluorescent emergency 'dome' lights.

Sometimes the dome filament is a different lamp to the rest as it has two concentric filaments, one for normal use and the other for emergency lights.

NAP LIGHT \

Fig. 8 TYPICAL FLIGHT DECK LIGHTING LAYOUT

SAQ 3

Why use incandescent lamps instead of the more economical fluorescent tubes in the flight deck area lighting?

Figure 10 shows some of the different types of lamps and their respective holders. For control reasons the flight deck is separated into sections and a control is provided for each section. All floodlights can be controlled independently or collectively. The collective control is often referred to as 'storm lights', and pressing this button will put on all the flight deck floodlights a t full illumination. There are many variations in flight deck lighting so you should refer to the specific AMM ATA chapter 33 prior to any maintenance or replacing any lamps.

- 11 -

moodull l A-1 347

AISLE STAND 'FLOOO LIGHT

-- - 7 1'

\ ----, - I I

Fig. 9 TYPICAL FbIGHT DECK LIGHTING C0,NTROLS ' I I

l i I

1 I I - <.

I

Fig. 10 TYPICAL FLOODLIGHTS

- 12 -

moodull l A-1348

It is a requirement of the A N 0 that the aircraft carries on the flight deck 50% of the total number of lamps used on the flight deck as spares. This affords the crew the chance to replace any lamps that might fail in flight. This occurrence should be notified to the maintenance personnel by the crew, entering the fact in the Tech Log sector reference pages.

Utility (or Wander) Lamp

Most aircraft are required to have, somewhere on the flight deck, a t least one utility light (see figure 11). These are a simple incandescent light connected with a coiled extension lead and a control switch on the back. Adjusting the focus is by movement of the front ring.

Usually powered from the hot battery bus (28V), which effectively turns them into an emergency lamp a t the same time. (For details of dc power distribution, refer to the books in this series on electrical power).

Sometimes referred to as 'wander' lamps

LIGHT- - - CONTROL.

RETAINING

R1NGY3 Fig. 11 UTILITY LAMP

Control of Flight Deck Lighting

This is usually done by altering the voltage to the lamps (the output of incandescent lamps is more or less directly proportional to their input, except for some heat and resistive losses, that are not significant in this application. For a dc supply a variable resistor type of dimmer control is used and for a n ac supply a variable transformer is used. (See basic electrical theory module 3 for details of how these operate).

Flight deck lighting control is more complex than it seems (figures 12 and 13) and usually involves:

* Each panel or instrument cluster or map or chart or floodlight has its own individual control.

* There is a master control that will brighten all the instrument lights together.

* Often an override switch is provided to allow full illumination of all the instruments immediately.

I I I , 'F&. 12 BRIGHTNESS CONTROL - EXAMPLE

I I - - - -- -

/

-- --- -- - --

To cope with these requirements there are several dimmer controls located at various locations around the flight deck. Figure 9 shows an example.

Figure 12 shows a circuit taking 28Vac from the ac transfer busses. Variable auto-transformers provide adjustable brightness control. This output is further reduced to 5 to 0 volts by additional transformers.

Each pilot has a transfer bus which effectively allows him/her to transfer electrical power from the other side of the flight deck should the supply fail on this side. It allows for back-up supplies for essential flight deck instruments.

Figure 13 shows another example of a control circuit. Note the 28Vac supply from the ground service bus and battery bus with dimmer control for the first officer's and captain's dome lights via the dimmer light switch, the dome switch and the flood dome switch. Note the 28Vdc standby supply to both front domes.

18V AC

l4 5101 lL,%3DoME Lls L4I A l l 111 OK DOYI

Z 1 V DC

Is" I' l W O O O l l l 11s

111 4 CIRCUIT B A E A K I R

L I ~ r los OOUI

1111 UIMSI nltnr LIOHTIHG CONIROL IPSI

118 C1PI S OOMC

r i t G n l COMPARIMENI OOUI t l e u i

-- -- 131 - . - -- -. .

\ - - -

Fig. 13 EXAMPLE - FLIGHT DECK DOME LIGHTING '

I

Figure 14 shows a more modern light control circuit. Switches alter the control voltage and sends the modified voltage-to'a\dimmer control unit:-(There may be a large number of these located behind the flight deck consoles - t h C ~ 7 7 7 has 14 for example). The dimmer contxol unit uses this voltage to alter the low voltage output to the instrument panel lighting.

A similar function occurs in the dimmer control unit for the'slightly higher voltage floodlights that are used~fo~instrument and panel illumination. It is not normal to control area floodlights (sometimes referred to as dome lights) in this way, especially as sometimes they are fluorescent tubes.

Pressing the master dimmer control will override all the control voltages and apply a maximum voltage to the transformer windings and bring all the instrument and panel illumination on to the highest level.

Figure 15 shows how Airbus vary the applied voltage to the dome lights without using any direct voltage control. The initial supply is 28Vdc to switch 5LE with DIM, BRIGHT and STORM settings and to switches 5LE and 19LE.

Each dome light has four halogen long life (about 2000 hour) bulbs.

MANAGEMENT PNL TO STBY INSTRUMENT LIGHTS (P2)

PNL LT DIMNER MASTER BRT CTRL & SU CTRL UNIT

m T - L I G H T I N G CONTROL S!!:!lN:?,,!!Tff:!!k":PN:kUb?GYr:! S::SHfJr .nI""I., YI..LI.-.a.ILYI.~L . ,.l.k.. b-..... --.."-- .-

(OUTER KNOB) ARE SIMILAR \

FWD PNL BRT/HTR CTRL (P I31 I

I - - -,, I I I 1

I I

, I

I Fig. 14 TYPICAL MODERN 'FLIGHT DECK LIGHTING I

CONTROL1 CIRCUIT ---- -1 ,/

<. I -- --

\

\ I -- --

I

-

-- --

601PP +2BVDC

403PP +28VDC

L - - - - - - J 5GA1

Fig. 15 DOME LIGHTING - AIRBUS

When switch 5LE is in the BRT position, bus 601PP supplies dome light 453VU with 28Vdc through circuit breaker 1LE2. The essential bus 403PP supplies dome light 452VU.

When the switch is in the DIM position the light decreases. In the BRT position the lamps are supplied in parallel (26V) and in the DIM position they are supplied in series in groups of two. This means there is a reduced voltage drop across each lamp.

When the switch set to STORM all lamps come on a t maximum brightness.

If the main bus is lost only dome light 452VU remains operative via the essential bus.

On the ground, in the acceleration stop configuration, the dome light 452VU is automatically activated what-ever the position of the switches.

Also on the flight deck are lighting strips each with four 28Vdc halogen 11.5W lamps controlled from a simple ON/OFF switch. ', - - - -

- -

Flight Deck Lighting - Automatic Control ,

I I

For some instruments (typically items like audio selector panels and radio panels) the illumination is controlled automatically. Instead 'of using a control voltage from the dimmer switch an amplified signal from a photosensitive cell is - used. I I I

The photo (light) sensitive cell is positioned in such a location that, as the ambient light-increases, so too does theyoltage from the cell, and this through the d h m e r control units, increasesthe voltage and hence illy&bation of the instruments. A s with other flight deck lighting this can be overridden to full illumination by the selection of the dimmer override button.

That covers flight deck lighting. We touched briefly on emergency lighting and shall return to it later. Now try the following exercises. A s always, the answers are to be found in the text.

Describe integral instrument lighting and list its limitations. What is a bridge lamp? How is a lamp changed in the pillar type of lamp? How is the position of the slot in the pillar larnp maintained? Explain briefly the lightplate and how it functions. Explain what floodlights are used for in the flight deck, what sort of lamps are used and how they are controlled. What is a utility lamp? Why are so many controls needed for lighting? Where is the requirement laid down?

Suggested practical activity - if access is available:

Locate the flight deck lighting controls, see how many there are and what effect they have. Check on the flight deck lights and identify the various types used.

Spend some time looking through the maintenance manual chapters that are relevant. In this case ATA33- 1 1.

PASSENGER CABIN & CARGO LIGHTING

SAQ 4

What is the most likely style of lighting to be found in the cabin area?

We have covered most of the area lighting system concepts when dealing with flight deck lighting, and passenger area lighting is virtually the same, though rhere is~onsiderabiy iess cuniroi.

/ -A -- 1

The liihting can be divided intol: , I

I I I

1 Area lighting (day/nikht lights),- attendant controlled. T Spot lighting (reading-lights) passenger controlled. I

T Emergency lights - atitGi%akic] crew controlled. I * I Information lights - automati'c/crew controlled. I

I

I i

Area Lighting' : ---- - - .-- --

Figure 16 shows a general arrangement of a passenger cabin lighting system. The main stay of the system are the fluorescent 'washlights' and ceiling lights. These operate at one of two levels, bright and dim. Also you might expect to see incandescent night lights, doorway entry lights and reading lights.

In the galley areas there are some work-lights (spot lights) and area lights, these being a combination of fluorescent tubes and incandescent lamps.

Control of the cabin lighting is from the area control panels located a t the attendant locations. This may be through software related systems (on newer aircraft) but equally through traditional rotary switches as found on most older aircraft. The control range is often as follows:

* High (sometimes called 'Day'). " Medium. * Low (often called 'Night').

INDIRECT AND NIGH1

OUTBOARD STOWAGE B I N

OUTBOARD PSU

SIDEWALL

Fig. 16 CABIN AREAPLIGHTING - GENERAL-LAYOUT

I

P 3 2 0 GND HDLG/SVCE PWR PANEL

Fig. 17 AREA LIGHTING CONTROL

What lights come on and go off under these switch selections varies from aircraft to aircraft but the basic principle of operation is the same. Generally on high, everything is on bright. On medium the fluorescent tubes dim slightly and when low is selected the tubes switch off altogether leaving just the small incandescent night-lights lit.

Control of the Fluorescent Tube

Figure 17 shows a switching system for the control of side-wall lights, ceiling lights and night-lights. The lights are controlled through a software package inputted from an area selection panel located in the attendant's area. Physical switching of the lights is carried out through transistor action remotely.

The fluorescent tube consists of a tube and ballast, with a power supply. The ballast is a transformer that steps the voltage up to that required by the tube to operate. Normally the ballast outputs 205Vac to the tube for 'dim' lighting and adds an additional 115Vac for 'bright' lighting. It is usually possible to bypass the switching system and provide 'full' power lighting from the ground service bus.rrhis-faciiity aiiows the aii'craf~-m be cieaned and serviced wiihoui ihe rriairl

7 , system Gower having to be on. [ --'

' I I I

; I I i

u PASSENGER CONTROL UNIT

PART

/"REE::Ak:NS LENS

Fig. 18 READING LIGHTS

- 20 -

moodull l A-1 356

SAQ 5

A ballast often has two tubes supplied from it. How can this help in terms of fault diagnosis?

Reading Lights

Most seats are provided with reading lamps. These are usually fitted as a 2, 3 or 4 lamp module in the ceiling panels immediately above each seat group. Some seats have individual stork mounted (or swivel mounted within the side head- rests) reading lamps attached directly to the seat, these are low voltage supplied and are individually switched. Figure 18 shows examples of those fitted above the passenger seats.

TX-

DC GND RX-

I I PASSENGER S E R V I C E UNIT

Fig. 19 ENCODING/DECODING UNIT

The lights are usually of the incandescent type as they are easier to control. Operation of the lights above the seats is initiated by pressing the light control button on the passenger control unit located in the seat arm-rest. This sends an ON/OFF signal (typically through an ARINC 429 data bus routed within the seat tracks in the floor) to a n electronics unit at the front or rear of the cabin. This sends a signal through the control line above the ceiling panels to the individual De-coder/En-coder Units (DEUs) above each seat group. This signal is decoded in each DEU and that switches on the appropriate reading light for that seat.

moodull lA-1357

The Control and Display Unit (CDU) on the flight deck is used to BITE test the system when the aircraft is on the ground. The CDU sends a signal to the electronic control unit which transmits a signal to each DEU. The DEU puts the system to test mode and puts power on to each lamp for a short period of time (the lamp should light for a moment). Normal switching is un-available during this time.

Test data is stored in the BITE memory of the electronic control unit and is displayed on the screen of the CDU. It will indicate if no fault is found, and if a fault is found will give the location of the unit that is not operating.

Attendant Call/Lavatory Assistance Lights

The attendant call system provides a method for the passenger to get the crew's attention should they require assistance.

LAVATORY

Fig. 2 1 ATTENDANT CALL - OVERVIEW

- 23 -

moodull l A-1359

Call buttons are fitted to each passenger seat control unit (in the arm rest), or above each seat, and located in each toilet cubical.

At the attendant locations and often above the emergency exit signs a series of indicator lights show. These display blue for seated passenger requests, amber for toilet requests and pink when the attendant's internal phone is ringing. In addition to these there are chime speakers above the attendant locations.

-- --

- - Fig. 22 SYSTEM GIRCUIT DIAGRAM

Briefly the system operates as follows. The passenger presses the button in the lavatory (or seat location); this applies an earth to the respective overhead electronics unit and the zonal management unit. These earth triggers and latches solid state devices that put the light ON above the toilet door/appropriate seat and the master lights at the attendant's station. Also (not shown in figure 22) the chime will sound.

The attendant goes to the location, addresses the passenger's needs, then presses the reset button which cancels the demand and clears the latches. All lights go out and the chime stops. The reset can also be performed from the main terminal should the need arise.

- - 34 -

moodull lA-1360

Cargo Lighting

Figure 23 shows the general arrangement for cargo bay lighting. It will consist of either recessed incandescent lights or the flush fitting fluorescent type.

CARGO

\ COMPARTMENT

\ LIGHT (TYP)

-- -

Fig. 23 TYPICAL-C~'GO BAY LIGHTING -- /

The lighting for the cargo bay is controlled from one switch located close to the freight door. The power to this switch is usually controlled to prevent the bays from being powered in the air. TherdBi'e several ways of doing this, but a common one is to provide a grGnd handling ac bus that is airground sensitive (ie it loses power with weight off the wheels).

The lighting is usually made up of a large number of lights in the ball transfer mat regions and the area where the cargo comes into the aircraft. The area going down into the aircraft hold tends to be less well lit. In addition a light is provided that shines onto the pallet loading truck. This light will either be positioned outside recessed into the fuselage or on the inside of the cargo door (remember of course the door is up during cargo loading).

Other Aircraft Lighting

It is usual to provide lighting around the aircraft in all the servicing bays, wheel wells, APU bays etc. These are always incandescent lamps with a switch located nearby. A s with the freight lighting, these will extinguish as the aircraft becomes airborne, though of course, they should be switched off before push-back anyway.

That concludes this short chapter on interior lighting. Now try the following exercises:

1. How are varying levels of lighting achieved in the passenger cabin? 2. What does a ballast resister do? 3. How many ballasts does a single fluorescent tube require? 4. Explain how a passenger reading light is signalled on? 5. Where are the cargo loading lights located?

Suggested practical activity - if you have time and access:

Look in the AMM a t the relevant chapters for cabin and cargo lighting. Study the circuit diagrams and make sure you understand how the lighting is controlled, and any tests that can be carried out.

Have a look at the types of lighting in your aircraft - their location and operation.

If you.canassist in any rectification work,then do so. Make sure you keep your persohall'Gg book up to date. I 7~ \

\ -

I -. ~xte r ib r lighting is, in si=iZT-to'the other types of lighting on the aircrag so I shal1,not repeat how it is controlled, but focus more on the types of lights and where and how they ar,e fitted. They include navigation lights, runway take-off iights, logo lights, taxi lights, wing illumination lights etc .

/ I

- -, .- -

- -- - - -- - - -- -

The Regulations

Certain requirements have to be met regarding navigation lights. Essentially it stems from the theory which is: see and be seen.

The A N 0 and the associated EASA CS25 define what must be fitted, how much light must be emitted, in what direction and the colour specification.

CS25 defines mandatory external lights as:

* Position lights. Red, green, and white near the extremities of the aircraft to show its position.

~i Anti collision lights. On or more red or white flashing lights to give adequate coverage.

- 26 -

rnoodull l A-1362

Position light requirements as stated in CS25, in general, should not be less than:

* A steady green light of a t least 5 candela to the side (40 a t dead ahead), showing to the starboard (right) side of the aircraft through an angle of 110" from dead ahead in the horizontal plane. Vertical plane intensities are also specified.

* A steady red light of a t least 5 candela to the side (40 a t dead ahead), showing to the port (left) side of the aircraft through an angle of 110" from dead ahead in the horizontal plane. Vertical plane intensities are also specified.

* A steady white light of a t least 20 candela showing through angles of 70" from dead astern in the horizontal plane to either side of the longitudinal axis.

The colours are Aviation red, Aviation green and Aviation white and their - chromaticity is defined.

- -- - -

--

CS25 regulations for anti-collision lights are (in general): I

* A flashing red or white light of at least 20 candela (407 in the horizontal plane), showing in all directions and 75" above and 75b below the horizontal plane of the aircraft. There may be more t h a n m e light to ensure adequate coverage and the flash frequency must be_not.less than 40 per minute and not more! than 100 per minute. , '

!

A flashing white strobe light is common on large commercial transports and some aircraft use a ?ota'ting lamp system.

-

--- -

Note. The candela (old name candle which is fractionally smaller) (cd) is the SI unit of luminance and is based on the light emitted from molten platinum as it solidifies at 2042K. At this temperature lcm2 will emit 60cd. (1 candela roughly equates to a quartz halogen lamp of 20w supplied by 28Vac).

SAQ 6

What implication will the failure of one navigation light have on the continuing airworthiness of the aircraft? What type of lamps would you expect to find being used for external applications?

Navigation (Position) Lights

Figure 24 shows a general view of all the exterior lights on the B777. Take a moment and study the drawing.

- 27 -

rnoodull l A-1363

ANTI-COLLISION LIGHT (STROBE, 3 )

ANTI-COLLISION P O S I T I O ~ LIGHT (5) LIGHT (BEACON)

Fig. 24 EXTERNAL LIGHTS - B 7 7 7

Note the beacon anti-collision lights on the top and bottom of the fuselage and the wing and tail anti-collision strobe lights.

, - 7 ,- --- Note the wing tip and tail moulte'd pos$iip~ lights.

Note alsd the landing lights (4), the taxi lights (2), the turn off lights (2) and the logo lights (4). -. / I

I /'

I ' I - _ i I

- \ 1

I

--

REAR POSITION LIGHT

Fig. 25 NAVIGATION LIGHTS - B 7 7 7

moodull l A-1364

Figure 25 shows the wing tip fitment details for the position and anti-collision lights.

Navigation/position lights are an essential system and control is through a relay activated switch in the flight deck. Normal power supply is 28Vac from a protected bus such as the essential or standby bus.

The lamps themselves will be a dual filament bulb to provide redundancy, or alternatively you may find two bulbs fitted in parallel with each other.

Another way of providing the back-up supply to the lights is to have a separate supply of 28Vdc from the battery, This is initiated by switching, in the flight deck, from ON to ON BATT. Not common these days, but still around.

Fig. 26 EXTERNAL LIGHTS - EXAMPLE 1

Figure 26 shows the external light arrangement for a B747. Note the angles for the position lights (mandatory) and the angles for the landing and turn off lights which are specific to this aircraft. Their main requirement under CS25 is that they produce a clear light to light the intended area and do not produce glare for the pilot. Note the anti-collision lights and the emergency lights.

Figure 27 shows a n example of another aircraft. Take a moment to study that drawing. Besides looking at the various fields of coverage for the lights, take a moment to study the flight deck switching.

- 29 -

rnoodull lA-1365-

-, '? , . - . . _ ~ . -~

, ,

- -

' I ' i i

i

I Fig. 27 LIGHTS I I - EXAMPLE 2 I

' ! i '

I

~nti-collision Lights i --

Anti-collision lights must be mounted in such a manner as to show all around the aircraft a red or white flashing light. Usually this means one light on the highest part and one on the lowest part of the fuselage and often lights on wing tips and'tail-plane, but there will be deviations from this.

- - - - - , .-

The lights must be on anytime that the aircraft is on the airfield with engines running. In addition the commander must be able to switch off these lights should the need arise due to glare or flash-back. If the aircraft was in fog for example the lights could simply 'flash-back' and dazzle the crew.

Figure 28 shows the simplest form of obtaining a flashing beacon. It shows a fixed lamp shining onto a rotating reflector. The speed of rotation is controlled at about 40-45 rpm, thus giving a flash repetition of about 80-90 flashes per minute.

Power to the rotating beacon is controlled from the flight deck through a relay in the power panel. Normal supply would be 115Vac and as with the navigation lights this supply needs to be protected in some way. Notice from figure 26 that the reflector emits two distinct light patterns. One half of the reflector is straight and gives a narrow beam of very high intensity light; the other is curved slightly to allow a wider less intense light pattern.

- 30 -

moodull l A-1 366

MOTOR AND DRIVE

METAL COVER

ELECTRICAL CONNECTOR

Fig. 28 ROTATING REFLECTOR TYPE ANTI-COLLISION LIGHT

Another type of rotating beacon employs the same principle but the lamp unit itself rotates. The unit has two lamps which rotates at about the same rpm and the reflector type

Xenon Flash Tube

Another method of producing a flashing light is the xenon tube gas discharge lamp.

The flash tube provides a very high intensity light for a short period of time. I t achieves this by applying a high voltage to two electrodes and then introducing a trigger voltage (figure 30).

- 31 - moodull l A-1367

Voltage drop Rectifier resistors

Main electrodes (2)

5 - H i \

Ignition , - - - - - Condenser I

n F c 1 1 ' voltage 1 -Main tube -

condenser

Ignition Step-up 11 8 I

~ynchronisin~ logic

control circuit

Fig. 30 XENON FLASH TUBE PRINCIPLE

The ac supply voltage is, if necessary, transformed to a higher voltage, which passes through the rectifier and charges the main flash condenser (capacitor) A s the-capacitor is charged a large voltage is now available at either end of the flash bube; typically about 5 0 6 ~ but-nothing happens yet. The flash tube itself i s filled yi th xenon gas a t low pessure.

I ' I

\

-

! Q. REFLECTOR- p, LAMP REFLECTOR

! -

RELAMPING H A N D L E 1 @

UPPER BODY BEACON

'-ELECTRICAL CONNECTOR

LIGHT ASSEME

WING ANTI-COLLISION LIGHT (REMOVED)

Fig. 3 1 TYPICAL ANTI-COLLISION LIGHTS

With reference to figure 30, the flash tube has an ignition circuit with a smaller capacitor (condenser) and a transformer. This capacitor is charged from a voltage reduction resistor pair.

- 32 -

rnoodull l A-1 368

The synchronising logic control circuit now allows the ignition capacitor to discharge through the transformer and into one of the electrodes. This induces a high voltage in the other electrode of about 10,000V. This voltage is high enough to ionise the xenon gas and cause current to flow. This produces a short duration high intensity white light, Basically, a streak of lightning has just been generated, albeit for only about 1 / 1000th of a second. The light is bright enough, however, to be visible several miles away.

The usual arrangement is for the flash tube to be separate from the rest of the components. The transformer unit being located close to the light a s from it comes the high voltage necessary to light the flash tube. Figure 3 1 shows a typical installation of such as lamp. It contains a parabolic reflector with the .

flash tube wrapped around the inside. Flash tubes can be made to fit any shape or size.

SAQ 7

What precautions would be appropriate when working on this unit? - - -

L --

Landing and Taxi Lights

As the aircraft approaches the ground it is dbsirable that the flight crew' can see ahead. The aircraft is provided with high intensity forward &nting lights to

-

illuminate the landing area. --\ ',, i

Fig. 32 TYPICAL LANDING & TAXI LIGHTS

The location of these lights is one of design choice, but will normally have a beam angle of about 14". CS25 states that they must not cause glare to the pilot. Landing lights may be fitted on the undercarriage, on the wing inboard end or sometimes they are of the retractable type. Figure 32 shows an example of those fitted to the landing gear and figure 33 shows a retractable type.

3

In the case of the retractable type the retraction mechanism is usually integral with the ON OFF switch. It is often also controlled by the undercarriage mechanism, in that when the undercarriage is selected down the lights will come down - when selected up the lights will retract and go off. Figure 33 shows a typical retractable installation with its associated control circuits.

Fig. 33 RETRACTABLE LANDING LIGHT CIRCUIT

Study the circuit and note the use of limit switches to control the range of movement of the lamp, the two field windings (one up, one down) and the control of the light itself.

Maintenance practices of these lights include checking that the alignment is correct, and that the lens is clean. Care needs to be exercised in the use of these lamps in that they can temporally blind and also they get very hot so, to prevent damage and cracking of the lens, the duty cycle must not be overrun in still air.

7 A - d7 -

moodull l A-1370

Because they are cooled in the airflow when airborne, when on the ground they are limited a s to the amount of time they can be ON. Duty Cycles are normally 5 minutes ON and 5 minutes OFF but check the manual.

A taxi lamp (sometimes termed a runway turn-off light) provides the crew with a reduced intensity light that highlights the edges of the taxiway and any runway turn off points. Usually a halogen type of lamp, like the landing lights, but of a reduced wattage. Again location is a matter of design choice, bu t common locations are the nose leg or the wing leading edge root.

Other External Lights

There are other external lights for a number of reasons, some cosmetic some for flight reasons.

Wing illumination (AN0 schedule 4). Lights are often provided tha t shine down the leading edge of the wings. These provide two functions. In flight they can be

k used to help identify ice formation on the wing, as a back-up to any automatic \ . ice detection system. On the ground they provide illumination-of the-,wing

helping to avoid hitting the wing with ground service vehicles etc: . ' I

Logo lights. These are installed in the upper surface of the hbrizontal tail surface and are designed to shine onto the tail logo. Purely a commercial thing that allows the airlines to essentially advertise a s they taxi around airports.

- - 8 - -

I

I Caution. When working on lights and lighting circuits it is important to isolate the power supplies and tag the system to prevent power being re-applied. Lights can get very hot when on (and just after switch-off) and present a burn hazard. With certain lighting systems --xenon.-tubes for example - high-voltages can be stored within the power circuit so must be left for a period of time after switch- off before any work is carried out. The 'wait' times are stated in the AMM.

There are emergency external lights still to cover, but for now I would recommend that you consolidate your knowledge before moving on. Try the following exercises:

1. Where are the requirements for external lighting laid down? 2. Where are the navigation lights to be located? 3. What colour light shows to the rear and through what angle? 4. Explain how a rotating beacon is controlled. 5. Explain the use of the ignition circuit in the Xenon flash tube. 6. What maintenance precautions must be taken whilst working on high

intensity strobe lights? 7. What is the landing light duty cycle? 8. What safety implication does a failed logo light present?

Suggested practical activity - if possible:

Check your aircraft MEL to find out which lights are 'nogo' items.

Look in ATA chap 33-30 a t external lights, find out how lamps are replaced, and look for dual bulbs and single systems. Study the control schematics to see how the higher voltages are controlled.

EMERGENCY LIGHTS

The purpose of emergency lighting, as its name implies, is to provide illumination in an emergency - if the normal lighting should fail - if electrical power fails - during emergency evacuation of the aircraft etc.

Emergency lighting is provided on escape slides, a t exits and escape hatches, and along escape routes within the aircraft. Emergency lighting must also be provided a t evacuee point of first ground contact. Escape hatches, doors etc must have minimum self illuminating values (in microlarnberts - a unit spec~ini_reflectiveness) or be-illuminated by emergency lighting. For aircraft -.-- seating 19 or more passengers orrarcraft over 5700kg the emergency exit signs size abd colour is specified and they must be internally illuminated with a

I I minimupl value. This is all laididown in CS25. i i I

1

The emelgency lighting must be provided~by an independent battery supply, althodgh the batteries can be chaF@d-by the aircraft electrical system.

I ' I

The ~ k d scale Z requires that dufficient lighting be provided both inside and outside the cabin to facilitate the safe evacuation of the aircraft in an emergency..~kis 'to be operable. in' the event of normal electrical supply failure.

I - - - -- - - - -,

It (and CS25) requires emergency floor path lighting in the passenger cabin sufficient to allow evacuation of the aircraft in the dark or in smoke conditions. These lights may be fixed to the bottom part of the seat structure next to the aisle or fixed to the floor. On some aircraft they are also fitted around escape hatches.

CS25 emergency lighting requirements include:

* The emergency electrical system must be independent of the main aircraft electrical system except that the batteries may be charged by the main system.

A The emergency batteries must be able to supply the emergency system for a t least 10 minutes.

* If the fuselage was to split in two in a crash landing then no more than 25% of the emergency lights should fail.

-k The floor path and emergency exit floor area must be illuminated and illuminated (back light) signs placed at specified locations within the cabin.

- 36 -

moodull l A-1372

Sign letter sizes are specified as is their colour (red) and back ground colour (white). Emergency signs are to indicate locations of emergency exits.

* Floor path lighting must be provided so if all illumination 4ft above floor level is obscured the escape route can still be followed by the passengers

* Switching by the pilot or cabin staff with switch selections ON, OFF and ARM. When selected to ARM the emergency lights will come on automatically if the normal electrical system fails.

* External lighting to include over-wing exits, escape slides and ground areas around the bottom of slides.

The specifications are comprehensive to include minimum lighting values and sizes for emergency signs and minimum lighting values and coverage areas for emergency lights.

Figure 34 shows an example of the emergency lighting on a n aircraft before the days when floor path lighting was mandatory. Take a moment to study the

\ drawing and note the battery packs and the portable exit lights.

- -

I \ - \

7 - - -. ,

Fig. 34 EMERGENCY LIGHTING - GENERAL

Floor Proximity Lighting

In a fire smoke will rise and fill any room space from the top down. This means that crawling out of a smoke filled aircraft gives the best chance of survival and floor proximity lighting is designed to indicate the way out. Figure 35 shows a seat mounted light and figure 36 shows a floor mounted system.

- 37 - moodull l A-1373

Fig. 35 SEAT MOUNTED FLOOR PATH LIGHTING

Fig. 37 FLOOR MOUNTED ESCAPE PATH LIGHTING

Seat mounted lights offer the advantage of being less susceptible to damage, but it does mean more cables running together with the IFE equipment and possible associated interference problems.

The purpose of the lights, regardless of where they are fitted is to enable a crawling person to find a way out of the aircraft. They are essential for flight.

Emergency Exit Signs

Figure 37 shows a typical example. These are situated within the cabin at locations to comply with CS25 and meet the requirements of size, colour and minimum internal illumination.

Fig. 37 TYPICAL EMERGENCY EXIT SIGN

As with most lights there are many variations on a theme, but the basics are the same. Note from figure 34 that the sign has the words EXIT written in red and an arrow pointing in the direction of the door. The cover panel is removable to give access to the two incandescent bulbs that run off the 28Vdc supply system that includes battery packs and charging circuits.

On some, particularly older aircraft, these battery packs were of the throw away type. Care needs to be taken to ensure that these packs are not inadvertently discharged. The lights are sometimes paired with a single battery pack and this poses a problem during maintenance. A paired light sometimes sees the missing lamp as a power failure and switches on the emergency lights.

moodull l A-1375

Slide and Door Illumination

Sometimes mounted in the side of the fuselage, but also possibly in the passenger door, the 28Vdc lamp will illuminate the deployed slide and some of the surrounding area. Lights are also placed to shine on the overwing exit areas.

Lights on either side of the slide illuminate automatically on slide deployment.

LAHP LEN'S

I I

i 1 Fig. 38 SLIDE AREA ILLUMINATION, I 1 I

I I I I I ' /

Lighting I~ontrol -

1 - _. _ _ - i Control of the emergency lighting is-effkcted from the flight deck by a 3 position switch (figure 39).

With the switch in the OFF position all the power is removed from the lighting systems, through relay action in the essential bus breaker box. The emergency lights will not illuminate a t all. This is the position that the switch needs to be in before power is removed from the aircraft. With the switch in this position an advisory message is displayed in front of the crew.

With the switch in the O N position the lights are all lit from the aircraft power supplies that are available. Usually this is from the hot battery bus (the bus that is always live whenever the battery is connected) or from the dedicated battery power packs if this is not available.

To select the switch either on the flight deck or the attendant's panel, the guard must be de-seated. The middle position and the one that the switch assumes when the guard is closed is AUTOMATIC or ARMED.

-

Flight Deck Control LIGHTS LIGHTS TEST

P87 ATTENDANT SWlTCH PANEL

Main Attendant's Locatlon

Fig. 39 LIGHTING CONTROL LOCATION

This is the 'flight' position and where the crew place the switch before take-off. In this position the lights are under the control of relays in-the essential breaker box. Providing aircraft power is-available &d on, the lights-are-off. Should aircraft power fail (or be switched off) the emergency lights all come on automatically. This will drain the qircraft battery power packs within ,about 10 minutes.

1 I .. -

EMERG EXIT OVERWING BATTERY

DOOR SWITCH ASSEM. 2 PLACES OVERWING EMERG

(2 PLACES) LIGHTS - 4 PLACES

SW CLOSED - DOOR IN

SW OPEN . DOOR OUT -pJJ

Fig. 40 POWER SUPPLY & CONTROL

- 41 -

rnoodull l A-1377

The battery packs are on charge when the switch is set to ARMED or OFF with the rest of the electrical power supplies available.

In addition to the flight deck control a control is provided in the cabin area, usually at the main attendant's (purser's) location. Door 1 left or 2 left are typical locations, but there are others. This switch will allow the lights to be switched on independent of the switch position on the flight deck switch.

A test switch may also be provided, this will put the emergency lights on for about two minutes. This is long enough for all the lights to be checked, but not long enough to discharge the battery power packs. There may also be a similar test switch for the exterior lights. Most modern aircraft have BITE systems so the lights can be checked using BITE.

Figure 40 shows a simplified circuit diagram for the power supplies for the exit signs, ovenving lights, escape slide lights, aisle lights, and door lights. It also shows the battery power supply packs (4). Note the switching and dc power supplies.

Fig. 41

- - - - - -- - -

-. - -.

EMERGENCY LIGHTING-LOCATION

0 EMERGENCY POWER SUPPLY UNIT

LXI EXIT SIGNS

@ CEILING LIGHTS

~ ~ P ~ \ ~ P ~ ~ f ~ I : : G W E R G E N C Y ESCAPE

@ ESCAPE SLIDE LIGHTS

A330 EMERGENCY LIGHTING SYSTEM - OVERVIEW

rnoodull l A-1 378

Figures 4 1, 42, 43 and 44 show the power supplies and emergency illumination for the Airbus A330. Study the drawings carefully and check on the component locations and the system operation. There is no need to remember the details, but the philosophy is similar for any modern passenger aircraft so you need to know the general operation.

E X I T S IGN 1 E X I T S I G N 2

FPEEPH 11 5VAC (CROSS-AISLE)

FPEEPM ? ISVAC (MAIN-AISLE)

(WAIN A ISLE) DEU-B

NOTE: @ SWITCH CLOSED WHEN - EMER S L I D E RELEASED

Fig. 42 LAMP POWER SUPPLIES

The electrical system is controlled and monitored by the Emergency Power Supply Units (EPSUs) - a total of 8 installed in the ceiling a t intervals along the cabin length. These are supplied with dc and ac (which is converted for use in the system). With failure of the dc essential power the EPSUs will continue to supply the system for a period of time from their internal batteries.

- 43 -

moodull l A-1 379

NORM BUS

E3 E X I T S I G N S

@ C E I L I N G L I G H T S

- 28VDC GND WHEN E X E S S I V E ALTITUDE-OR ; Nor.; a < :!M NO SMOKING S I G N ON

O F F

I

l:EbEPRFHnffEK:E!Rf:!k:s 0 ESCAPE S L I D E LIGHTS

I Fig. 43 A330 EMERGENCY ,POWER SUPPLY UNITS ; I - ,

I I ; --. \

' 1 --

The floor path lighting Electro qukinescent (EL) (figure 42) flexible light strips are installed along the aisles in the carpets and also on the non-textile floor coverings of the' galleys and cross aisle areas.

1 / ,'

~hotoluminescent strips are used on some floor aisle areas for emergency path lighting (figure 44). These absorb light energy from the normal cabin lighting, or from daylight, and will emit this back out as light when in the dark - ie when there is no daylight and other forms of lighting are OFF. These strips will give out light for about 8 hours under normal operation.

Testing

The BITE test will test the system and the capacity of the batteries. The system test, which has a limit of 20 seconds to prevent battery pack discharge, includes:

* Testing the dc and ac supplies. * Checking for short circuits on each output. * Checking the voltage output of the battery packs. * Testing the battery heating devices. * Checking the EPSU logic and switching circuits. * Checking system loads.

rnoodull l A-1380

The battery capacity BITE test checks that each EPSU battery has sufficient capacity to run its emergency lights for a t least 10 minutes and does not have a drop of more than 0.40 AH between two consecutive tests. The test will take up to 3 hours to complete.

NON T E X T I L E FLOOR COVERING

NON T E X T I L E FLOOR COVERING

Fig. 44 PHOTOLUMINESCENT -- LIGHT STRIPS

That concludes this section and this module on aircraft lighting. Try the following exercises, the answers as usual are in the text. '

1. What is the reason for embrg&nii lighting? I - ,

2. List the two types of floor proximi$li~hting. - -- ,

3. How is emergency lighting controlled? I I

4. How are exit signs powered?

- - - -

Suggested practical activity - if you can: I _- _

Check out the AMM and study the wiring circuits from the power supplies to the lamps. Check the location of the components/lights on the aircraft.


SAQ 1. The oxygen in the air has to be removed somehow. Initially and until fairly recently this was done by evacuating the bulb. This had limited success and it is much more common these days to use a gas with no oxygen or possibly an inert gas. The gas used will change the colour of the light emitting from the bulb. Common choices include Argon or a mixture of Argon and Nitrogen. These two both provide a white/yellow illumination. You may also see Neon used which provides a more orange illumination.

SAQ 2. Across the incandescent lamp there would be a small resistance. Across the fluorescent tube there would be a complete open circuit and an infinite resistance.

SAQ 3. The first reason is one of choice. Consultations with aircrew and development over a number of years has meant that the incandescent light provides a 'more pleasing' illumination than the fluorescent tube. Sounds a bit like crew pampering, but it is important the flight deck affords them the best options. Another reason is one of control. To change the illumination of a fluorescent tube the voltage has to be changed. Usually there is only two intensities available, dim or bright. The A N 0 stipulates that flight deck lighting needs to be controllable to accommodate all flying conditions. Incandescent lamps offer this facility, as the luminescence is proportional to the voltage applied.

SAQ 4. In the passenger cabin good general illumination is required. Bright in the daytime and virtually dark at night. Cost is a consideration here and the most effective light source is the fluorescent tube. They are used extensively.

SAQ 5. This has ideal fault finding possibilities. It is the use of the 'half split' method of looking a t the last known good place. If one lamp is on and the other isn't it is highly likely to be the tube. If both lamps are off, both tunes could have gone, but more likely it is the ballast that has broken.

-- ---- -- - - -

SAQ-&; ~z~ mandatory piece of eqtiiment must be fitted and operational before a flight cbn commence. The A N 0 gives the r'pquirements for this and is translated into other documents such as 'the Minimum Equipment List (MEL).

I I

All external lights are tungsten \filament type, and most are of the halogen fill, quartz envelope style. This is b d c a u ~ - t h e luminescent intensity is got sufficient for any &her type of bulb to be uked. (1 will introduce during the text one notable addition to the list of bulbs in use - The Xenon flash tube, more of that later.) 1

/ i- -

~~QG.-Several-t 'hou~hts shouldcome to-mind here. 1. We need to be aware of the dangers to ourselves (or others around us). The voltages concerned are large enough to kill. The unit contains capacitors and sensible precautions before working on any such system would be to allow several minutes to elapse to allow any charged capacitors to dissipate the charge. 2. when the system is tested, care must be exercised as the light is so intense that it can damage the eyes. 3. Strobe lights can c:ause stroboscopic effects and these can appear to show rotating machinery as stationary (propellers and rotors etc). So we need to take care on and around aircraft. White strobes are to be switched off during night time activities around aircraft and red strobes should be on for a minimum time. 4. Due to the heat that is generated during operation we need to ensure that oils don't remain on the glass surface of the tube. This would cause localised hot spots and subsequen.t cracking.

CONTENTS

Page

General High pressure systems High pressure system - services Low pressure systems Air supply Flow, pressure and temperature control Enane-ble,ed system - ---I

Preslsure/control - -- \ ~ e r n ~ d r a t u r e control I --\ \ \\

~istrihujion network I , '

Leak detection i ; t

Other /a& sources I , / /

,/' ~ervicks, - engine and pylon / vacuJm1 systems I \\\\\\

1 , ~ n s w d r s to SAQs, i 1

'\ ', 1 I , ' /

I I / i : I I

GENERAL

Pneumatic systems can be divided into two main areas:

Low Pressure (LP) Systems. These are typically medium to high volume, low pressure systems that are used for engine starting, wing anti-ice, engine anti-ice, cabin air conditioning, pressurisation and other aircraft systems.

High Pressure (HP) Systems. These tend to be lower volume, but higher pressure to operate such systems as flaps, landing gear etc. HP air systems several disadvantages but are used on some smaller and older aircraft for the operation of services, and even then their use is diminishing as hydraulic systems have a better power/weight ratio and are generally easier to maintain.

Advantages of pneumatic systems:

I- --1 -- - - - -

* L-Sofne weight saving in !that theT<'is no return pipeworkrequired - as in

I I---- ' \\, '\ \ hyaraulic systems. 1 1

\

1 I \

' 1 I ' / 1

* There is an abundant supply of the medium. Air is available (free) all the

/

I ti&e, so we don't need to tarry a rese/rvoir of fluid. , / L - - <, I 1 - 1

* ~uldst components do not :quire liX$cation. I

-1

I 1 I I\

I I \

* Pdour freel,/air has no toxins in it - thbugh it often becokes contaminated yi;h aircraft oils etc and qmells. , 1 I ; '-/, /

L- -1 ' / 7 1 -

, -- Disadvantages of pneumatic systems:

* Leaks are difficult to trace.

* Not suitable for large components due to the rapid drop in pressure when selection is made.

* Air carries moisture which condenses and posses a corrosion, contamination and freezing threat to the internal workings of the components and the pipework.

* High pressure air is "explosive" in nature. Should failure occur to any container filled with high pressure air (or any gas for that matter) then it will burst with explosive force. Hydraulic fluid in a container under pressure will split the container if it fails but will not erupt, as fluid (up to about 3,000 or 4,000psi) is more or less incompressible.

- 1 -

moodull l A-1 384

* As the systems generally require high volumes of air to move the actuators, the compressors and the pipework can be large - for low pressure systems anyway.

Designers tend to favour hydraulics over pneumatics or electrics because of the high powerlweight ratio, but LP air systems will be with us for many years.

HIGH PRESSURE PNEUMATIC SUPPLIES

Not many aircraft these days have this type of system, although they will be found in some aircraft.

Figure 1 shows a basic system.

Compressed air is generated from-engine driven compressors ,-these-c.an produce up to-3,500psi (24.1 MPa) , but more commonly that figure is-aropnd-.1,500psi (10.3MPa) - with some systems yorking &:600psi (4.1MPa). %h< relief valves ielieve excess pressure should the normal pressure rehlators fdil to ensure 'that the system is not ovkr-pressuri'sed. 1

I I -- /' I

-- -. -. , -

SAQ 1 Why is the pressure kelief va&e p'laced near the cornpress&?

The air passes through the Presdure Regulator (if this facility is bot fitted directly to the compre3sor) then to the Oil b d Water drain (sometimes cal!ed an oil and water t rap)~Oil is introduced during-the-lubrication of the compressor and water is always present in air. Most of the oil and water is removed by the trap.

The air is usually passed through a filter, which may be placed further down stream and repeated in front of any debris sensitive component. There may also be fitted a dryer filled with silica gel that removes the remaining moisture from the air. An anti freeze bottle may also be fitted.

SAQ 2 How do we know if the dryer is working or if the crystals have become saturated?

There are a series of check valves (non-return valves) in the system and these will prevent back feeding of air from one compressor to the other (also loss of pressure from a bottle should a leak occur up-stream of the bottle).

The air now passes to a number of air storage bottles, again through a series of check valves. The system shows a primary bottle (though not all systems have them).

moodull l A-1385

This provides for damping of the compressor output and storage for assistance in major system operation eg, flaps or landing gear. (Similar to the accumulator in the hydraulic system).

In the system shows all three bottles are the same and are interchangeable, but their functions are different.

PRESSURE RELIEF VALVE

LEFT HAND

PRESSURE

Fig. 1 HP PNEUMATIC SYSTEM

The brake bottle provides pressure to the wheel brakes, which cannot be used by other systems due to the presence of the check valve. In an emergency, should system pressure be lost then the brake pressure is available from this dedicated bottle.

rnoodull l A-1 386

SAQ 3 Given the high pressures, what are the likely sizes of the pipelines and what materials are they made from?

SAQ 4 What is the purpose of the ground charging point?

Lookng a t each component in turn.

Compressors

These are either single piston (two stage compression), or double piston (four stage compression) operated.

They are usually engine driven though some may be electrically powered. The two stage compressors produce pressures up to about 600psi whilst the four stage compressors will give pressures u p to 350Opsi. -

- - I

They are iubricated with oil from the engine oil system. Any t i l u r e indication is sent to tkde flight deck. I

1 I

The single cylinder double compression compressor first com$resses the air in the top, side of the cylinder, then: compresses it in the bottom side (the lower part of the cylinder and piston arrkgkment being a smaller swept volume).'

, I

I

The four 'stage compressor has1 two pistons with each compression sequence similar t o the above - so the air is compressed four times.

-

Because the compression is adiabatic the air gets hot so cooling fins are fitted to the compressors to help in heat dissipation.

Normal pressure control is by the use of a Pressure Regulator Valve (unless pressure control is fitted within the compressor design).

The failure indication works by taking a reading of the pressure downstream of the compressor. If this falls bellow a certain level, an earth is made and this allows a lamp to be illuminated on the flight deck.

Pressure Relief Valve

The purpose of this valve is to relieve the pressure should failure occur in the compressor control or Pressure Regulator Valve for whatever reason. Its operation is simple (usually using a ball and spring). Figure 2 shows a pressure relief valve.

The valve is held down on its seat by the spring. Pressure in the system tries to de-seat the valve. Under normal pressure the spring is strong enough to hold the valve on its seat. If the system pressure approaches the preset level (overpressure) this overcomes the spring force and the valve is de-seated, allowing system pressure to be ported to atmosphere.

The spring force can be adjusted and in doing so the pressure at which the valve de-seats can be altered within a certain range - any adjustment of which must be carried out in a properly equipped maintenance bay.

\ : 1 ' I

if , Fig. /

/

ADJUSTER SCREW

$ , ,

1 HIGH P ~ E S S ~ R E A l R

\ \ 1 I

PI-essure Regulating Valve

This may be fitted close to the compressor or actually on the air bottle itself (if a single bottle system). The whole idea is to off-load the compressor when normal system pressure is reached. It works similar to an Automatic Cut-out Valve in a hydraulic system.

When normal maximum system pressure is reached pressure acts on the underside of the bellows which opens the outlet valve to port the air to atmosphere, a t the same time a check valve within the unit closes to prevent air escaping from the system. The system pressure keeps the outlet valve open via the bellows.

When system pressure falls to a preset value the outlet valve spring pushes the outlet valve closed and compressor air then goes through the check valve and normal charging continues.

PORT TO OUTLET OVERBOARD VALVE

CHECK VALVE

AIR FROM COMPRESSOR I

___)

I

Fig. 3 PRESSURE REGULATOR VALVE

- - -

SAQ '5 - These valves are expensive compared to an ordinary-Pressure Relief ' Valve - and less reliable. So why not fit a ~ r e s w r e Relief Valve

instead? 1 I I

-- I

Oil and Water Trap \ I I - ,

---

I /

I \

Figure 4 shows a simple Oil and Water rap. >The air comes i'n a t the side and is caused to impinge'on the baffle plate. The heavier droplets of water and oil stick to the plate and run down to collect at the bottom of the trap. The cleaner air is directed out of the top of the unit into the, system. The trap has to be drained periodically as laid down in the maintenance schedule.

AIR IMPINGES ON HALF TUBE BAFFLE

.;piTzu~

COARSE STRAINER

DRAIN w Fig. 4 OIL & WATER TRAP

Not all of the airborne debris is removed by the trap so the 'clean' air is passed through a filter (and sometimes also a dehydrator) to remove the remaining particles of water and oil. These extra items require periodic cleaning/ changing/ emptying.

Air Storage Bottles

Normally made from steel, with some wire wound on the outside, and then sealed, before being painted (usually grey). This identifies it as an air bottle and it will also have a tag with the date of manufacture, safe working pressure and the date of the last hydrostatic test stamped on. Usually installed in the upright position with the connection at the bottom. They sometimes have a stackpipe standing up within the neck of the bottle preventing moisture from entering the system that has collected in the bottle.

Periodically they will require, draining-and purging with-dry compressed air. Bottles,stored off the aircraft should be-done so in a verticdrac,k. At i ts lifed date the bottle will have to be remove$ for hydros'tatic and other tests.

I I I I

I I I i I I i I

Anti FrLeie Bottle I ,-- -- ,/' ,, ' 1 I

I - \

I ' _ I - -

Fitted doivn stream of the Oil & Water ~ ; a ~ and any dehydratqrs aridfilled with antifreeze (ethylene glycol or similar - but kheck the AMM). Thfs is picked up by the air as/ it pas?& through the bottle. It pt'events any rnoisthre that might be in the air /from freeqing. It require+ fhe rnaidtghance task of tbpbing up the anti- /' freeze-at kgular intervals. I i - - 1 ,

7 / 1

I

Pressure Reducing Valve

i Figure 5 shows a schematic drawing of a pressure reducing valve. This valve is included in the system, as some units require lower pressures than normal system pressure. Brakes require the highest pressure usually, with flaps and gear retraction requiring lower values.

With reference to figure 5. It can be seen that there is a spring and bellows mechanism. The spring balances the air pressure above it. The spring has an adjuster screw and this is used to set the operating pressure of the valve.

When air supply commences the valve is below its operating value; the spring is fully up and all the HP air passes through to the air out port. As the pressure rises to just above the operating value the HP air on top of the bellows starts to compress the spring; this in turn moves the valve inlet plunger to the left, via the bell crank level. As this happens the inlet to the valve is choked and the pressure felt at the outlet is reduced.

rnoodull l A-1 390

VALVE INLEX' PLUNCgR

m A I R IN LP A I R OUT

BFTL (IRAN)(

FIXED PrVOT

COHPRESSION

ADJUSTER SCREx

Fig. 5 PRESSURE REDUCING VALVE

r - ----

When air outlet pressure reaches maximum the valve shuts completely. I 1 1 1 In real time, provided the maximum outlet pressure is not reached, the valve is

constantly moving in and out to keep the lower outlet pressure korrect. I

- --

Check Valve , i

These may be designed with a ball and spring; a half ball and spring; a plate and spring, or a flapper valve and spring. Figure 6 illustrates a flapper valve type and a ball-and spring type.

-

HINGE FLAPPER VALVE

SPRING

- __t

\ ARROW (marked on body la show now d8recbon)

%ALL SPplNG

\ ARROW (mark& on body I0 shwr flowdlrecboo)

Fig. 6 CHECK VALVES

rnoodull l A-1391

The spring is reasonably weak and as such requires little effort to overcome its force providing the air is moving in the free-flow direction (in the same direction as the arrow shown on the outer casing of the valve). If the airflow is reversed, the spring will close the valve assisted by the reverse air-flow.

Pressure Indication

A bourdon tube type pressure gauge is fitted in the system for maintenance purposes where-ever there is a need to monitor the pressure. If pressure readings are required in the flight-deck then a pressure transducer is fitted in the appropriate part of the system and a dc of ac supply is sent to a moving coil or ratiometer type instrument.

Figure 1 does not show these transducer positions but they will be fitted where- ever the designer feels there is a requirement, and they could be fitted at the

i pressure , gauge positions anv-ay.- , ,

~ i r c r d t Services I

The adr sl I edge &d

similar in

1 I

lpplies will be fed to aiicraft - _-A services (eg, landing/ &r, flaps - trailing leading edge, windscreen_wipers etc) via selector valves. These are principle to the manLally o&$ed valves fitted to hydraulic systems,

the main difference is that the2e ,;Ke no return lines from thelvdves - returning air bein2 ported,tA outside the yalpe. Pipeline connections toeach valve are: 1 - presTure, 2 - to one side of the actuattr (jack), 3 - to the other side of the actuator.-If the actuator is returnkd by-the action of a spcng then there are only two pipeline-connections to the-selector-vave - pressure and exhaust.

The HP system can be used to supply air to a variety of systems to include:

* Brake systems. * Flap systems. * Gear retraction systems.

Aircraft with HP pneumatic systems tend to be older and often use a cabin blower for air conditioning.

Let's first of all take a look at actuators (sometimes called jacks). These fall into one of two main types, single or double acting.

SAQ 6 Consider why they are called this and, therefore, briefly explain the difference.

- 9 -

rnoodull l A-1392

Single Acting Actuator

Figure 7 shows a single acting actuator. Pneumatic pressure is applied from the selector valve to one side of the piston. This overcomes the spring pressure and moves the actuator (in this case in the extend direction). Removing the pneumatic pressure via the selector valve allows the spring to return the actuator to the retracted position.

AIR FROM SELECTOR VALVE RETURN SPRING 1

Fig. 7 SINGLE ACTING ACTUATOR --

- - - -

Carbon impregnated rubber seals are used to prevent leaks aro;und the ram and piston. Single acting pistons are limited in their applicationlbut are used in brake systems and sometime in gear retraction systems. i I

I i -- L

SAQ 7 Apply your underst&ding of flyink controls and dneumatics together to dec'ide why we don't use single acting actuators on flap drive systems? I

-- -

- - - -

Double (or dual) Acting Actuator

Figure 8 shows a double acting actuator; notice it is essentially similar to the single acting one. The difference is that pneumatic pressure is used on either side of the piston depending up selection.

RRM DAMPER PISTON FIXED S L I D I N G

\ PISTON ROD PISTON HEAD

/ / ,CYLINDER HEAD

AIR- IN/OUT FLUID PORT 01; P ~ L L E D

Fig. 8 DOUBLE ACTING ACTUATOR

rnoodull l A-1 393

The selector valve is similar to the two way hydraulic selector valve (at least in principle), in that if one side of the jack is pressurized and the other is ported to atmosphere it will move. The picture shows an inner damper piston which is not always fitteda41t prevents rapid movement and hence 'piston slap':

R Basic Landing Gear System

Figure 9 shows a complete supply system similar to figure 1 except that it has 'attached to it' a simple jack system to operate a retractable landing gear.

The selector valve is shown in figure 10. Mechanical up and down locks will be fitted (but not shown in the drawing).

M F L A P S 3 EXHAUST

Fig. 9 HP SUPPLY & ONE SERVICE SYSTEM

A n 'up' selection is received from the landing gear select lever in the flight deck. This sends a 28vdc signal to the selector valve (via the closed contacts of the uplock sequence switch). This supply energises the up solenoid (as shown). This in turn moves the pilot valve and allows compressed air through to the right hand chamber, which moves the control piston down, and through mechanical linkage moves the left control piston up.

The right hand chamber allows the compressed air through to the up line and this acts on the 'up' side of the main actuator. Also at the same time the left chamber has allowed the 'down' side of the main actuator to port to atmosphere.

- 11 -

rnoodull l A-1394

CDKPRESSED AIR SUPPLY

Fig. 10 ELECTRICALLY OPERATED SELECTOR VALVE

--

The gear moves, and is locked in \he up position by other rnkans-(spring loaded plunger lock for example). As this, happens, the sequence s+tch on the uplock actuator will open, removing the 28vdc from the selector solehoid. The pilot valve now returns under spring pressure and through differential {orces on the piston surfaces the control pistons return to neutral.

- -

If the solenoids were "removed" and control cables or push/pull rods connected instead direct to the linkage, then the valve would be mechanically selected.

Braking System

On large aircraft it is standard to use hydraulic braking systems because of their better powerlweight ratios. However, lighter general aviation aircraft still use pneumatic braking systems, not to mention the old 'dinosaur' aircraft that are still with us today.

Pneumatic brake systems can be divided into two categories:

* Hand operated. * Pedal operated.

Pedal Operated Systems. Figure 11 shows the location of the 'foot motors' and figure 12 show a cut-away of a typical foot motor.

'Tipping' the toe pedals forward operates the system. This has the effect of moving down the brake control valves. As this happens the valve sleeve moves into the motor body and closes off the exhaust valve. Pressing the pedals further down allows the valve sleeve to push the valve stem up and starts to un-seat the valve head. This allows some pressure through to the brakes. Pressing still harder opens the valve head still further and increases the pressure to the brakes. This provides for 'progressive braking'.

The harder the push on the toes the harder the brakes go on - exactly the same as in a vehicle.

TOES PUSH FORWARD ---. RUDDER P W A L S

- - 1 -

, ir J BRAKE CONTROL VALYES -

' I (FOOT MOTORS) I

I I I

I ! I I RUDDER PEDAL I '

1

TORQUE TUBES 1 1

l i / I - ,

I I I /

' 1 /

I I\ I I 1 I

/ r i I / i -1 Fig. 11 BAKE i -- PEDAL'FOOT MOTORS

- / - 1 - - - - - 2 - - I

ATTACHED TO AIRCRAFT STRUCTURE

IXLm VALVE

AIRIHLET VALVE STm

BRAWS

SPRING

Fig. 12 DETAILS OF FOOT MOTOR

Hand Operated Systems. The heart of this system is a Dual Relay Valve (DRV). This is fitted in the flight deck close to the rudder pedals. To apply the brake the pilot will pull on a control cable (a Bowden control cable; a Teleflex system; or even a n ordinary control cable), by moving the input lever, which is usually mounted on the control wheel. This puts equal pressure down to each brake (via the DRV) and thus gives straight line braking.

If the pilot applies rudder in one direction at the same time as pulling the brake lever the DRV will apply air pressure to one main landing gear whilst pressure to the other side is reduced. This produces differential braking and allows the aircraft to be steered (on many older aircraft this is the only way to steer the aircraft).

Both systems have the ability through pawls and valves to hold the brakes on - parking brake. So much for the pilot's side of things, what happens down a t the business end?

- - -- -

Brake ~ r u & Assembly. Figure 1.3 shows a typical brake d d q assembly. I I \ I

On application of brake pressu;e, air is forced into the reinforced rubber brake bag. This inflates and forces the' brake shoes outwards into contact with the bimetallic keel drum. The brake is,c?nnected to the wh'eel and is therefore rotates with it. The brake shoes are fixed to the hub and axle. On b r d e release the retbrn springs pull the brakk shoes ahay from the drum. Vanes-on the top of the drum are to increase heat dissipation f;om the drum. I

I

I BVAKE 8- OR LINING RrVBT BRAKE , e x ~ m r o l r n ~ CLIP 1

-- -

D W I H G FRW CAP 561

Fig. 13 TYPICAL DRUM TYPE BRAKE

Control and Indication

'There are few controls and indications for these systems.

Control over the compressor is limited as normally the compressor is a fixed stroke compressor and is designed to give the correct level of output.

Anything above the required pressure is exhausted overboard by the pressure regulating valve or an in-built compressor system. This is all mechanically controlled and the pilots have no indication of its serviceability.

There could be a pressure transducer (on older aircraft there would be a pressure line to the flight deck with a bourdon tube type pressure gauge in the instrument panel), which could be a variable resistor type or an E&I bar type or a synchro type. Ratiometers could also be used of course. See the LBP books in

(' this series on Instrumentation for details.

dc BUS

I:

r -.--, - - - - - -

Some-prcraft have an indication d an over'-pressure (or under pres-sure) situation I develops! One manufacturer uses a bellows arrangement conhected to a micro-

I 1 switch1 &d at the pre-determined level thd micro switch is made and puts a warniig light on. There could bk a micro-,switch attached to tfle valv2 seat of the over prqssure relief valve (or ah -- under A pressure transdlcer). As the valve operates the switch is made apqlyingg e'&h to the warning light-via a 'hold on re-lay' (figure 14). The hold-on 1 relay keeps;the warning illuminated until

I / L- - '

vL VALVE SPIN

the pilot

LATCHING RELAY

u n l a t c h ~ s the relay. I I I /

EARTH 2

/ i 1 I I

1 I I

1 I /

I I

DLE

f j r

P EARTH 2

-MICRO SWITCH

PUSH TO RESET

Fig. 14 HIGH PRESSURE WARNING CIRCUIT

SAQ 8 What colour do you think the warning light would be on the flight deck and where should it be placed?

SAQ 9 Why is does the switch 'earth' the lamps / warning light and not operate the supply? (It is normal practice to close a circuit by using a micro-switch to provide an earth to a. normally live bulb).

That concludes this section on high pressure systems. As stated earlier they are - -

not common these days, but you may still see them and the subject is in the J A R 6 6 syllabus.

Perhaps now is a good time to consolidate your knowledge of HP systems before moving on to LP systems, which are common to all commercial pressurised aircraft.

Exercises

1. List the advantages and disadvantages of an HP pneumatic system over a hydraulic system, --

I

I 2. Assuming the compressors are running correctly, what protects the

system from over pressure? And how does it work9 /

I -

3. The bellows of the p,ressure reducing valve are cracked, what effect I would there be on the flight deck? I

I l 1

1 1 4. Explain the operation of the components on the HP system.

5. Look in your AMM at chapter ATTA 36, and study the systems just - described. This may not be possible if you are not-working on older aircraft, but have a look anyway.

6 . How and when storage bottles purged?

7. How is the action of some single acting actuators damped?

8. If the system failed to build up pressure what might be the causes? Try and list at least 10.

LOW PRESSURE SYSTEMS

High volume^ low pressure (hot) air is supplied normally through large diameter ducting, and this pneumatic supply can be used to deliver air to any of the following systems (in some cases suitable cooled):

* Wing, fin, tailplane, slat and windscreen anti ice (see the book in this series on Anti Icing).

* Cargo compartment heating. * Engine Anti Ice (EAI). * Air conditioning packs (air cycle systems) for cabin air conditioning. * Air driven hydraulic pumps (auxiliary pumps). * Hydraulic reservoir pressurisation . * Engine starting. * Thrust reverser(s) . * Potable water (drinking water) pressurisation. * Smoke detector aspiration.

1 - - - -- . - -

*I ,prpde heating and aspiration. -,

* 'Sdme systems operatioq (the ? 4 ~ uses it for LE flap deployment for k&nple but this is not common). I I

I I 1 1 I I I

I I I /

/ / I

- /' 1

AIR SUPPLY . I -- I I r - - '

The a& can come from a numb& of source's and these include: 1 17 I 1 I I I 1 I

1 ~ s i & a turbo cornprdssor as fitted to some of thh qlder aircraft. _,* L- -& &ngine driven cornpressbr'or blower - such as the Godfrey

/

1 - - - compressors as fitted to the -VC 10. -

- -

J; Using the exhaust gases from a turbo prop engine. * A tapping (engine bleed) from the compressor side of a jet engine. * The APU (Auxiliary Power Unit). * External connections from ground carts etc.

1

The air has been compressed adiabatically and is hot. It is ideal for anti-icing but has to be cooled for such purposes as cabin conditioning and pressurisation.

Figure 15 shows a schematic of the pneumatic system as fitted to the early B747 and figure 16 shows the general arrangement of the system as fitted to the A320. Study these drawings carefully and note the ways the various supplies are connected, note also the services that use the air supplied.

The Turbo Compressor

Originally designed for the piston engined aircraft where engine driven compressors where fitted for the supply of air.

rnoodull l A-1400

- --- \

- 1 ,

Fig. 15 B747-,100 PNEUMATIC SYSTEM I

1 !

To wing anti-

To air conditioning packs fh ',* ventilation

Wing anti-ice valves

/ \ \ Engine HP/IP bleed

Crossfeed valve Ground Fan air bleed connector

Fig. 1 6 AIRBUS PNEUMATIC SYSTEM

Pure jet engines would, at certain times, generate insufficient bleed air from the engine. During these times of high load the crew would open the shut-off valve and use some of the bleed air to drive a turbine driven compressor. This would draw in external air from the engine intake and after compression add it to the manifold. The manifold being the central pneumatic system for supplying all the services.

This sort of system is limited to older pure jet aircraft - the B707 for example.

SAQ 10 When is the lack of air likely to occur? And when would the crew run the turbo compressor?

Fig. 17 TURBO COMPRESSOR

The Compressor/Blower System

Not too unlike, in principle at least, the HP systems described a t the begmning of the book. The compressor is driven by the engine and supplies air to all services such as de-icing, cabin conditioning, pressurization etc.

The compressor (on the VClO at least) is a screw type with two intermeshing (large) screw threads driven from the auxiliary gear box of the jet engine. It is fitted with a slide valve which automatically regulates the air output.

Two compressors feed the air to services such as pressurisation etc.

Again, fitted to some older aircraft.

Engine Exhaust System

Not common but used with some turbo-prop engines. Some of the jet efflux is ducted away from the jet pipe to a heat exchanger. Here it exchanges its heat to clean ram air before being ducted back into the jet pipe. The hot ram air is used for de-icing/ anti-icing/ heating purposes.

HOT RAM AIR, TO SYSTEMS zh

JET EFFLUX

v I

I JET PIPE I

I I I

Fig. 18 JET EFFLUX SURFACE HEAT EXCHANGER ~ Y S T E M - -

- - . - --

SAQ 11 Why is the jet efflux system sometimes fitted to turbo-prop engines and never to pure jet engines and why is the air put through a heat exchanger and not used 'as is'?

The most popular system on most commercial aircraft is to use air taken from the compressor of the jet engine.

Engine Compressor Bleed System

This consists of a tapping (or tappings) from a stage (or stages) of the compressor of the jet engine. When more than one tapping is used it is usual to have a low pressure tapping and a high pressure tapping with modulation provided for the high pressure tapping.

The tappings provide an abundance of clean air - provided the air entering the engine is clean in the first place and there are no engine malfunctions such as front bearing oil leaks. The air is hot - about 150°C - but check the actual 'aircraft manual.

With reference to figure 19. It shows a typical arrangement of the tappings on a fan engine. This has two LP tappings (about the 8 th stage) and a H P tapping (about the 15 stage).

At low engine power settings air is taken from all the tappings with non-return valves in the system to prevent the air entering the LP side.

A s engine power is increased, enough air can be provided from the LP side and at a pre-determined pressure the HP valve shuts-off automatically. This air coming from the engine is called Charge Air and is at about 50 to 60psi pressure - but again check the manual of the actual engine.

(- \

- 1 - - - --

Charge +r normally needs coolingand,?his is usually carried out using surface heat e qhangers (similar to radiators in a road vehicle) and iusing ram air as the P cooling medium. To assist this1 skstem A? Cycle systems a re used such as the ~ o o t system. I I I

1 ' I I / , 1 _ - _ A '

In some aircraft a Vapour Cycle Cooling System (refrigerant s$stem)'is used this system is not common. ( ~ d r mor&inf&mation on cooling of charge air

\ \ to the book in this series entitdd b b i n Cnnditinninrr 2nd pr&ssurisatinn .l

, HP s m OFF VAL,?

----

S P A R T W

V3U.n NRV

I KFAT EXCHANGER BYPASS VALVE

ENGINE BLEW

OVERHEAT VALVE

PROTEITION

t.- PNEUXATIC OVERHEAT PROTECTION

., but refer

Fig. 19 TYPICAL HIGH BYPASS ENGINE PNEUMATIC SYSTEM

rnoodull lA-1404

Ozone Converter

Fitted to some aircraft in the supply ducting from the jet engine - on the BAe 146 for example, (those fitted with this option), It is between the engine tapping and the isolation/pressure reducing valve. It removes contarninants/odours from the air supply by a reaction process between the metal plates and the hot air.

The converter consists of special metal plates fitt.ed within a metal container covered in a fire proof material and connected into the ducting using "V" clamps.

C A T A L Y T I C

I / I

- --

I

I Fig. 20 CATALYTIC OXONE CONVERTER - - ' I

I ' I I I I

FLOW, PRESSURE AND T E M P E ~ U R E CONTROL I

The charge air must be at the correct flow rate to meet the demands of the user systems and must be at the correct pressure and temperature.

Because of the adiabatic compression of the air it is always hot so the biggest problem is that the air may get too hot. In this case cooling is provided usually in the form of surface heat exchangers using ram air. The exchangers are placed in the airflow and often placed within the engine fan airflow on fan engines.

ENGINE DRIVEN COMPRESSOR SPILL VALVE

---m TO SYSTEM

FLOW CONTROL VALVE N.R.V.

Fig. 21 SIMPLIFIED SCHEMATIC - SPILL VALVE SYSTEM FOR ENGINE DRIVEN COMPRESSOR

13ressure and flow may be controlled by the same valve as they are related, but the way it is controlled depends on how the air has been compressed.

For air that comes from an engine driven compressor or cabin blower then, if the flow rate is too high (at high engine rpm), the excess air is dumped over-board (figure 2 1). If the air comes from a tapping on the compressor casing of the jet engine then a restrictive type valve system is used (figure 22).

PRESSURE NON-RETURN REDUCING

VALVE VALVE

_I) 1 C__C

ENGINE SHUT-OFF VALM FLOW COHTROL TO SYSTEM

COMPRESSOR STAGE

Fig. 22 SIMPLIFIED SCHEMATIC FLOW CONTROL FOR AN --- ENGINE BLEED SYSTEM - -

- r 7 , - - -

1 I \

\ \ 1 I l 1

SAQ 12 1 Why is the excess &r !dumped over-board when it Comes from an I 1 engine driven corndressor and y h y is it simply restricted when it

/ ) comes from the en&&? -, / I I I

Spill System ,,> I I /' !'

1 I 1

~ e s i ~ d e d I i_ to __- spill unwanted air bvbrboard (hat is being deliveted from an engine L- --

driven compress'or/ blower. -i L - -- ,' - -

With reference to figures 21 and 23. The charge air flows through a venturi within the duct where density and velocity are monitored. The static connection (pressure) supply is fed to one side of a diaphragm whilst the venturi connection (suction) is supplied to the other.

The diaphragm will move in response to these pressures and move contact Y to close onto Z or X depending on direction of movement of Y. This will send a signal, via the phase reversing contactor, to either close or open the spill valve butterfly.

The eIectrical supply is to contact Y and altitude compensation is allowed for in the provision of the absolute capsule pack.

actuator c l o s e s the sp i l l va lve the actuator opens the spill va lve

-

T,he duct s t a t i c union s e n s e s a The Venturl tube provides a negative pressure relative to the DENSITY pressure (suction) relative to thy of the supply air. VELOCITY 01 the alr flow.

I

Fig. 23 SPILL VALVE SYSTEM I I

Engine Bleed Systems - -

-- -

The amount of air bled from the engine will depend on the demand from the services. The amount of air coming from the engine will depend mostly on engine rpm. At high rpm there is usually too much air being delivered so air that is not required is stopped from leaving the engine by a restrictive type valve system.

SPRING INLET

OUTL

' A L E

PISTON ORIFICE PLATE

Fig. 24 VARIABLE ORIFICE TYPE VALVE

moodull lA-I407

On the simplest system this is achieved using a Variable Orifice Valve (figure 24).

The air enters the valve and leaves via the orifice plate. When flow rates are high the pressure acts on the piston which is caused to move to the left against a spring. This closes the acorn valve a little, restricting the air leaving the unit

The above system may be used on some smaller/older type aircraft. With larger more modern systems the flow-rate is controlled by a bleed system.

A MODERN ENGINE CONTROL SYSTEM

A modern engine compressor bleed system consists of a number of sub- systems:

* Engine air supply. I - - 7-- - -

control and indication. - - \

* 1 Qistribution network. 1 1 i \

Conti-ol/indication/warning {auld include: 1 I I I I

-- - - /

High stage p ~ s s u f e control valve. 1

- \

Low pressure waning. \ / \ '1 I I - - - High temperaturd protectioh -. thermal shut-qff.1

I ' I 1 I /

~ h e s k &sterns 'usually have :bkick-up ,$sterns and appl$ to primary and I

interrnediad air as appropriate. - --l , I - - _ - , r '

- - I

The HP Valve

This valve can be controlled in one of two ways, pneumatically or electrically. Figure 25 shows a typical pneumatic control system.

Pneumatic Control

With reference to figure 25. The system is made up of two parts, the High Stage Valve itself and the Valve Controller.

At idle power we would expect this valve to be open but initially it is closed by spring action. High stage air is felt in chamber A and this keeps the valve closed. Opposite the chamber A orifice there is another off-take going initially to a test port and then onto the controller. The Shuttle Valve moves over and the air pressure is allowed to the switcher solenoid (this is energised to close the valve so is normally de-energised as shown).

- 25 -

rnoodull 1 A-1408

",r.i+5,uS m.tl@ ,,I sub7

I _ - - -

I

I

O Y € t . l O ~

_I- L_------- I I - / ,

Fig. 25 PNEUMATIC PRESSYRE CONTROL SdSTEM ,

! I

I i l

Air passes through the Over Pressure Control and (for now) through the Low Pressure Control, it now h a s a free run to chamber B of the High Stage ~ a v e d i a p h r a g m . Differentidpareas now force the vdve to open and allow 1 5th stage (HP) air into the pneumatic system.

At a predetermined pressure the HP valve will shut, allowing IP (low pressure) air into the manifold. For this we need to focus on the Low Pressure Shut-off control in the high stage controller.

Initially as the pressure is fairly low this is fully open, as the air felt on top of the diaphragm is not sufficient to move it against the spring. As the engine speed increases and the HP pressure increases the pressure on top of this diaphragm and it starts to exert a downward force on the control aperture. This starts to close and in doing so reduces the air getting to chamber B of the HP valve, causing that to start to close.

Further increase in flow and pressure closes the aperture still further and eventually the HP valve is completely closed. This actually happens when the HP bleed pressure is about 120psi.

There are other items in figure 25 that we shall return to when we consider indication and protection. But for now lets move on.

- 26 -

rnoodull l A-1409

Electrical/Electronic Control

The system is similar to the one previously described, it has an HP valve sprung loaded closed and opened by pneumatic pressure. There is still a controller and a valve. With reference to figures 26 and 27 showing details of a digital system and the inner worhngs of the electrical controller.

, ,' , 1 - /' / 1 1 1

' / I

i __- -- Fig. 26 E L E C T R O ~ C PRESSURE CONTROL- - -

A I R

I I+/ ) AMBIENT VENl

SUPPLY A I R (FROM HP A I R SOURCE)

Fig. 27 FLOW CONTROLLER

- 27 -

rnoodull lA-1410

The valve (as before) is sprung loaded closed and bleed air keeps it shut. 15th stage air enters the controller where it passes through two regulators, these bring the pressure down to 55-65psi. An electronic unit measures the pressure down stream from the high stage valve. This will initially be very low/zero. The unit sends an electrical signal to the torque motor to move in the direction to cover the vent. This allows the full 65psi down to chamber B on the valve to open it.

A s the engine speed rises the pressure downstream will rise. The electronic unit senses this and sends a signal to the torque motor to open the vent and start to cover the inlet. This reduces the pressure to chamber B and starts to close the valve.

The electronic unit in figure 26 is called the ASCPC, it is an Air Supply Controller.

SAQ 1 3 What would happ-en if the air supply controlkr failed? I I

--- --

SAQ 14 What would happen if the pipe from the contrqller to the valve was leaking (through a loose connection or b e ~ a u s e of damage)? I I I

I

1

I - -

So far we have considered just the HP valve. The purposL of that; just to recap, is to control when we take HP air from the engine and when we allow 'unregulated' IP air from the engine. Now that might pose some problems. Surely we can't have unregulated air in the pneumatic system? True - read on.

- - /

- - -

PRESSURE CONTROL

Another valve is added to the system, not too dissimilar from the HP valve, down stream and this is used for pressure control. Referring back to figure 19, look for the Engine Bleed Valve. (You may know this valve by a different name on your aircraft, such as Pressure Control Valve etc.)

This engine bleed valve will typically have the following functions:

* Pressure control of the pneumatic system. * Non Return Valve, preventing a stronger engine from cross bleeding

to a weaker one. * Shut off valve - through flight deck switch action. * Reverse flow to allow cross bleed engine starting. * Temperature control (backup mode).

- 28 -

rnoodull lAp1411

Pressure Control Mode

The valve behaves in a similar fashion to the HP valve and functions in the same way. Sprung loaded closed and pneumatically opened, it can be controlled pneumatically or electrically. It is usually controlled to a slightly higher pressure than the HP valve. It is this valve that sets the pressure in the p-neumatic manifold and ducting, which is typically about 55psi.

Non-Return Valve Mode

This closes the valve when the pressure sensed by the downstream port is higher than the upstream port. This is done by the servo unit in the controller and vents the sense line to atmosphere, allowing the spring to close the valve.

Shut-off Mode '..

- 1 I -

- (-

L

If the flight crew are unhappy withthe engine or the airflo,wTthe air to or from it can be stopped by enkrgising a so,lenoid. It allows the sense ,line to ' I vent t q close the valve. In addition to the switch action ofi the flight crew this bolenoid will also be ener'gised if tlie~fire handle is pulleld.

I

~eveksk Flow Mode 1 ' \ I

I / I i I ' If the 4ngine i s 'not started it kslnot pro,dubing any air. A s the start valve is downstread of the engine bleed--~alv< (oh start selection), due to the non

1- returnmode the air will ndt-be able-to reach the starter71n this case the engine bleed valve is put into reverse flow mode and forced open. On some bleed valves this function is removed as designers fit the start valve off-take downstream of the engine bleed valve.

Over Temperature Mode (Back-up)

Primary temperature control mode has yet to be covered, but if the primary mode should fail the Air Supply Controller (the electronic unit) will signal the engine bleed valve to start to close. This reduces the mass flow and hence the temperature. There will be a reduction in flow (which could cause a problem), but not such a big a problem as over heating the ducts.

OVER PRESSURE CONTROL

I,P systems can suffer from over pressure which is normally the result of the HP valve remaining locked open at high engine speed. Simple spring loaded flap operated pressure relief valves are fitted in the system to prevent over-pressure.


TEMPERATURE CONTROL

We have already touched on a back-up method but lets now look at the primary method of temperature control. Figure 28 shows a typical method of cooling the air.

This shows charge air being cooled in a precooler cooled by fan air (ram air from the fan of the engine). The amount of fan air allowed through the precooler is controlled (modulated) by a modulating valve taking its signals from the fan air temperature sensor.

On some systems, particularly engine driven compressor systems, the air may not always be hot enough. In this case a Choke Valve is placed in the supply duct and when operated chokes the air in the ducting and causes the air supply become further compressed and heated.

-- FAN A I R

/ OUTLET --

\ 4 I

I I

BLEED AIR I I BLEED AIR DUCT (FROM -- DUCT (TO

FAN AIR -

INLET

Fig. 28 CHARGE AIR COOLING

ENGINE)

-+ - I

I - I

Charge air is un-modulated passing through the pre-cooler. To control the temperature we allow more (or less) of the cooling ram air through. This is achieved by altering the position of the modulating valve. This valve may have other names such as Temperature Control Valve, Fan Air Modulating Valve, Pre-cooler Valve, etc.

FAN AIR; I -

TEHPERATURE SENSOR /

Control of this valve is done in one of three ways:

I (PRSOV)

\ ,' I

BLEED A I R PRECOOLER

* Bi-metallic switch/valve operation. * Pneumatic operation. * Electrical operation.

FAN A I R MODULATING VALVE

The last method is the most common, but the first two are still in use on older and lighter aircraft.

I - 1 ' ----------

PNEUMATIC SENSE TUBE

- 30 -

rnoodull l A-1413

PRESSURE REGULATING

AHD SHUTOFF V A L V E

AIRPLANE)

l 1 -+

Bimetallic Control

Figure 29 shows a bimetallic control valve. It is placed downstream of the pre-cooler and as such will feel the warmed fan (or ram) air after it has passed the pre-cooler. If the temperature of the air is high the metal will expand and open the valve; as the temperature falls the valve will close. The valve never quite fully closes.

0 1 - H F T A L L I C TEXPERATURE

, S M S I N G S P R I N G

.ow FROM EXCHANGER

MODULATING

' I VALVE OPEN j I VALF I CLOSED

I , / I -- ' 1

I Fig. 29 BI-METALLIC FAN AIR ' I I ~

\

'\ \ /?

\

I I I pneumatic Control I I i I / 1 1 j 1 1 i (-1 1 I This achieved by placing a probe in-the airflow from thepneumatic system

I (figure 3 0). -- - - -- -

As the temperature of the air rises, the sleeve of the probe expands quicker than the inside and de-seats the ball. This allows pneumatic pressure in the sense line to the temperature control valve to bleed off. It is a sprung loaded open, pneumatically closed valve, and as such will open more. This has the effect of allowing more cooling air across the pre-cooler.

NOTE. This type of probe has also been used connected direct to the HP valve, operating on the same principle but closing the valve this time. This is because it is usually the HP air that requires cooling only as the IP (low pressure air (even at its hottest) is still below the over temperature value of .the pneumatic system.

/'- BLEED M ATMOSPHERE

BW AND SEAT

HIGH EXPANSION - HATEXIAT..

LOW EXPANSION

KATERIAL

Fig. 30 ;BI-METALLIC PROBE I

Electrical Control - /

I I - - I -- , I

Figure 25 is actually the Boeing 777 system, and figure 31 shows the Airbus A319 system so that we can consider both temperature control systems together. They are similar to each other and similar to other aircraft, the only real differences being the termi~ology and component

- details. - - -

Study figure 31 and follow the air supply from each engine, the APU and the ground supply to the various services:

* Air conditioning packs 1 and 2. * Hydraulic reservoir pressurisation. * Drinking water tanks pressurisation.

The main parts of this system are:

* The Air Supply Controller (ASC) (Electronic). * The temperature sensor. * The fan air modulating valve.

The temperature sensor sends the signal of the value to the ASC. The ASC then sends a signal to the Fan Air Valve to either open or close depending on whether the system is too hot or too cold. The Fan Air Valve has a rotary variable displacement transformer (more of that later) to provide feedback to the ASC of its position. This is a closed loop negative feedback control system.

rnoodull lA-1415

-

I _ - \

I I 1 -

Fig. 3 1 SYSTEM CONTROL - A 3 19 ,- /

1

1 1 I r\ 1 I I I

\ i / i '

NOW :to look j~ t some of the c o ~ ~ o n e n t s ~ , I , i / I - 2

I - /

ri - - - , I -

The Temperature Sensor. Refer to-fig-~ire 32. If a conductor is heated its resistance value changes - for most materials the hotter the material the greater the resistance (some have a negative coefficient of resistance and the hotter they get the lower the resistance).

Fig. 32 TEMPERATURE SENSOR

Heat is supplied to the temperature sensor to compensate for the cooling effect of the airflow. As the airflow temperature changes so the temperature of the sensor changes and so does its resistance.

It forms one leg of a Wheatstone bridge, the output of which is sent to the ASC for processing.

The Fan Air Valve. It is sprung loaded to fail open. This ensures that the maximum cooling effect is available. (Incidentally, as far as "Fail Safe" is concerned, the problem of what position a valve should fail in is one of common sense. Think about what is least desirable if a component can fail in one of two modes and in the event of failure the valve should go the opposite way.) The valve has a manual position indicator so that the engineer can see the exact position if necessary.

Also it will have some form of feedback pick off. This may be a digital pick- off, or a potentiometer, or a variable displacement transformer. As a guide, Airbus tynd to use the former whilst Boeing tend to use the latter. Whatever is fitted the principle is-the s&e. -

I

I \ l 1 The Air8 Supply Controller. The ASC will take signals from various systems and the, pneumatic sensors for processing. This informatioh can be:

I / * Analogue, voltages are either ac or, dc from a variable displacement ' transformer, or a potentiometer (vadable resister). 1 -

* Digital, possibly from a $igital pick-off, but equally could be a digital I I Binary Coded Decimal (BCD) word from other systems.

* Discretes. This type of signal is an open/closed typei signal. On or off. -

- - - -

From this information the unit calculates how far each of the valves has to open. It then sends the signal to the valve and monitors for a response. It takes into consideration how much air is required by the aircraft. (No point bleeding air off the engine if you don't need to.) For example as the pilot switches on an air conditioning pack, a BCD word is sent from the pack controller to the ASC, the ASC opens the engine bleed valve more to compensate for the loss.

Modern aircraft are automated and there are many signals moving from unit to unit. This can make fault finding difficult and we need to take notice of all the fault codes. You will learn more of this in the LBP book covering JAR module 1 1.18.

There are often two (or more) ASC7s to allow for a degree of redundancy.

SAQ 15 The temperature control valve of a pre-cooler style system is defective and is sticking and slow to operate. What 'symptoms7 might you expect to see in the technical report raised by the flight crew?

- 34 -

rnoodull lA-1417

Over Temperature Control

The temperature sensor sends a signal either to the shut-off solenoid (via a relay) in the pneumatically controlled system or to the ASC in the digital system. This has the effect of closing the engine bleed valve and hence stopping the pneumatic bleed. The aircrew (or ASC) will now open isolation valves to allow cross bleeding from the good enginejs).

That concludes this chapter on control, indication and protection. It has been somewhat of a marathon, but the JAR 66 syllabus requires you to have a sound knowledge of systems operation.

The details of the control systems described need not be committed to memory but you should understand them.

THE DISTRIBUTION NETWORK AND MANIFOLD

the

APU I

Fig. 33 TYPICAL PNEUMATIC SYSTEM MANIFOLD

- 35 -

moodull lA-I418

Study the system noting all the components and their location. The system comprises:

* Ducting and joints. * Isolation valve(s). * Pressure transmitter(s). * Off-takes to user systems. * Non-return valves * Ground cart connection * APU connection

Ducting

The ducting used is usually titanium or stainless steel. Stainless steel has a slightly higher tolerance to temperature changes and a better coefficient of expansion; as such it is used nearer the engines. Often, by the time the a i r reaches the engine bleed valve is sufficiently 'cool' t o use the lighter titanium ducting. . \ ,

I

The ducting is supported throughout its length with clamps and stays the aircraft structure. An example is shown in figure 34.

Fig. 34 TYPICAL DUCTING SUPPORT

It is important that the supports are not over stressed and that the ducting is aligned correctly. Care needs to be taken to ensure that the tie rods are adjusted properly. Shims may be used to provide for better alignment.

When the ducting is in use it gets hot. Even the 'cooled' air leaving the pre- cooler will be in the region of 120-150°C. The ducting needs to be able to expand without buckling, and still be leak proof at the start of operations (cold) and when it is hot. To achieve this, flexible joints are used and there are several types available. Figures 35 and 36 show two examples.

rnoodull lA-1419

DUCT

EXPANS ION & GAP

(IROSS SECTION OF

ASSKIBLED WCT

Fig. 35 "V" CLAMP CONNECTION

In figure 35 the expansion is taken up by allowing the sections of ducting to expand into a gap that is closed by the clamp. That in itself causes a couple of problems. Firstly the clamp becomes lose and can slip. If this is c in -- a critical location, next to a cable run for example, the clamp will r e q ~ i r e ~ s o m e form of anti-roiation device and this is often i n t h e form of

i r- two springs. Also this type of clarnp'rekpires the ducting 70 be slightly stresseld (in tension only - ndt bending ever). This means! that imaginative solutiops have to be found to klbsing th& last clamp in the pipe-run.

I t I 1 / I / '

Fig. 36 SIMPLE PIPE COUPLING

Figure 37 shows a different approach to the same problem. Here the ducting is allowed to slide in and out of the flange. This type of seal allows for a certain amount of flexing caused by aircraft loads in flight. It does not allow for ease of fitting.

Bosses are welded on the ducting at 120" separation on the circumference. A seal is slid over the joint to prevent leakage. Three cables are connected and tensioned to keep the ducting together. Not a common type of fitting, and a variation on this uses the seal 'sleeve' idea which replaces the cables with a clamp that goes around both ducts and the seal.

- 37 -

rnoodull l A-1420

MEASDRED GAP

m SWAGED

SEAZ, DUCTS

Fig. 37 EXPANSION JOINT COUPLING -

\ I -

I 1 I

These types of joints tend to leak around the seals in the flange. When putting any pneumatic duct together a small addition of silicone grease assists in the assembly and doebn2 add dlanger to the sydtekn - but check your AMM as to any lubricants used. -

- -

~ /

I

SAQ 16 What rnaintenanck activitiesi might be requirc!d o n pneumatic , ducting and what precautions should be carried out?

I -

--

Duct Maintenance

Treatment of minor damage to ducting includes (but check your AMMISRM):

* Smooth dents are normally permitted providing they do not substantially restrict the airflow.

* Shallow scratches and gouges are generally acceptable providing the bottom of the scratch is smooth and it is not deeper than 10% of the wall thickness.

* No defects are permitted within 1/4 inch (6.5mm) of any fusion weld.

Titanium is susceptible to hydrogen embrittlement and as such needs to be kept away from acidic chemicals, such as trichloroethalene.

Fire resistant hydraulic fluid (Skydrol) breaks down at high temperatures and will cause embrittlement. The normal operating temperature of the ducting is high enough to cause this and as such, great care needs to be taken to ensure Skydrol does not come into contact with the pneumatic ducting.

rnoodull lA-1421

Leak Detection

Leak detection falls into two categories - checks carried out after installation and checks carried during service due to duct and joint failures.

After any installation is carried out the system is pressurised and checked for leaks. This is normally carried out using a ground cart, but even so, any leaking air will be hot. The following is a general procedure (always check your AMM):

* Visually inspect the ducting and the joints for signs of discolouration. Escaping hot air, particularly over long periods, may leave fine black deposits.

* Pressurise the manifold and listen for leaks. Even quite small leaks can be audible.

* Attach tape or thin cotton to a stick and pass it over the general area of the suspected leak. Any movement or fluttering will establish

- - - -

[ - Fx'actly where the leak is. - \\ [ - r - * : eemove the pressure andlrem&? the joint, then try again.

I I I

I

~eaJsb that develop in flight Fan damage1 local structure; babies land any / equi$ment that is near. If yo$ consider,,the normal operating terriperature .- -

rangk of a pneumatic systerh 1s cl&e..to the ternperaturk for,/annealing r-- -, aludidium alloy and above the temperature for precipikation--tr&atment, \

darn&e can soojn occur to the 4tructure if\subjected to the leaking air for a 1 perid,d bf time. qeaks need to be detected quickly and isolated/repaired.

/ I I I , /

1 Some &rcraft have a 'firewire' style-of l e k detection systdm that detects

, anyoverh-eat situations. '- - - - - - - -

The firewire is run the length of the ducting and is usually split into sections so that the system can work out exactly where the hot air leak is. Firewire is covered in detail in the LBP book on Fire Detection Systems but is general they may be of the Resistive Type; the Capacitive Type; or the Systron Donner type. In all cases a local overheat is detected and a signal sent to the flight deck.

On older aircraft this would illuminate a warning on the flight deck and the crew would take the appropriate action. On more modern aircraft the signal is fed to a card file or processing unit (computer). This registers the fault and sends signals in BCD form to the Air Supply Controller. The ASC will automatically close off the section of the duct that is leaking by using the isolation valve(s) and/or the engine bleed valves.

The crew would do this manually on older aircraft.

Unit detectors or thermal switches can dlso be used for leak detection.

A s the switch heats u p contacts come together and this will put the indication light on the flight deck. As with the fire wire system the switch can equally signal a computer to operate a n automatic process. (Again described more fully in the book Fire Detection Systems).

Isolation Valves

Figure 38 shows an isolation valve. It is a motorised valve that can be selected to one of two positions. Normally in the open position, but can be closed for system failures, duct leaks etc. In the basic system of figure 33 there is one valve fitted, but normally there would be two (or more).

WING ISOLATION VALVES

SHAFT

I POSITION I I INDI CATOI.

WING ISOLATION VALVE' (TYP)

- 0 Fig. 38 ISOLATION VALVE

- -

The valve is moved using an electrical actuator supplied with 1 l5vac. Figure 39 shows the circuit for control of the various valves in the system - four engines and an APU. Each motor incorporates two micro switches (limit switches) that stop the motor when it reaches the end of its travel. The valve can either be selected by the aircrew, or more commonly these days is automatically operated. In this case the valve will either be closed or on automatic.

Figure 40 shows the flight deck control panel associated with figure 39. Study both figures so that you get a good overall picture of how the system works.

In digitally controlled systems an OFF selection will turn the unit off, but an ON selection will put the unit under the command of a computer. Some manufacturers have renamed them to 'request' switches rather than the traditional on/off switches.

I I I I

I I

I

1

I

1 1 '- I I I

V I I C I

1 I I I

I I

V 1 L V 1 I I I

U I I I I

I . I I

V

. . - -.

I I

! -- -

Control the position of the wing isolation valves.

DUCT PRESSURE INDICATOR Has two pointers marked L and R indicating left and right hand ducts.

Fig. 40 FLIGHT DECK CONTROL PANEL

Pressure Indication

When we considered indication in the previous chapter, pressure indication was deliberately left out. On newer aircraft pressure is measured a t the engine off-take, but this is for monitoring reasons and is not for system pressure reasons.

Pneumatic system pressure is taken downstream in the region of ten feet (3m) or so from the engine. This allows any pressure fluctuations to have flattened out before the reading is taken. Figure 41 shows a typical pressure sensor electrical diagram.

- - - - - - - - - - - - - , - - - - I FLIGHT ENGINEER'S PANEL

Fig. 41 PRESSURE INDICATING CIRCUIT

This works on the synchro transmitter principal. A 28vac is supplied to the excitation coils of the transmitter and the indicator. As the pressure changes the excitation coil moves relative to a toriodal winding. This induces a signal in the winding, and this signal is then passed along two wires to the flight deck indicator, where the reverse happens and the toriodal winding moves the flight deck instrument pointer.

SAQ 16 Notice both excitation coils get the same excitation voltage, can you explain why?

Mass Flow Indication

Figure 42 shows an early system for measuring mass flow. The transducer consists of a small turbine arrangement placed in the duct that is free to rotate.

Turbine rpm is proportional to flow rate so by providing the turbine with excitation and a similar excitation coil in the synchro, it becomes a fairly simple job to obtain a frequency signal dependent on the air mass flow rate. This frequency signal is passed into a signal conditioner where it is converted to a voltage for moving coil instrument displays.

I I - r-- I '

\ ' ,\ \ , ,

PS CIRCUIT BREAKER -- AC '1 PANEL ! I

1 ,

j ,'

, .

PACK 2 ,,

PACK NO 3 AIRFLOW SENSmA

I PACK NlU 3 AIRFLOW SEN30A

Fig. 42 MASS FLOW RATE INDICATION

Figure 43 shows a more modern method of measuring mass flow. The sensor is a platinum resistance device that when its temperature changes its resistance changes. A s the air flows passed, the temperature of the resistor drops and this is measured, usually, directly by the ASC.

HOUSING , ELECTRICAL

R E S I S T I V E TEMPERATURE DETECTOR

Fig. 43 RESISTIVE TEMPERATURE INDICATOR

--- --

The-ASC h a s held, in d a t a b a n k ~ , ~ the values of res is tanc~ expected for various airflows. The signal is drocessed and sent as a BCID wdrd to the symbol generators for the flight deck instruments.

I

OTHER AIR SOURCES - -

So far this book has covered air coming from the engine$ only, but other sources are used and figure 33 shows 2 of them.

I

I

The Auxiliary Power Unit

The Auxiliary Power Unit (APU) is usually located at the back of the aircraft in the tail cone section. It has two functions, these are to provide 400Hz 30 2001 115vac to the electrical systems and pneumatic pressure to the manifold.

A check valve is installed so that when the APU is not running air from the pneumatic manifold cannot back feed through the APU.

Also there is a bleed control valve and some form of dump valve. These two work in opposition to each other. When one is open the other is closed. The bleed control valve is similar to the engine bleed valve but in this case it is a simple shut-off valve (it doesn't regulate the pressure like the engine bleed valve did).

VERTICAL SUPPORT TUBE

THERMOCOUPLE PROBE AND HARNESS ASSEMBLY

FUEL I --

D R A I N S L I N E , -

AND VENTS - -, \

' I r ' -

\

' I

1 ~ i ~ l 44 APU'~DETAILS I

I

I

I I I I

I I ' // /: I -- ' /

Air p'rebsure from the APU is contro!l6d,by altering the 1hlkt Guide Vanes to the toad compressor. orei ion this i"n the Jet Engine bdoks in this series entitled APU. 1-1 I ~ \

I I 1

I I I I 1 ; I

An APU cannot stop produdirig air when it is running, even when the J syst&n doesn't require it. The dump-ddve or surge bleed~valve opens when

the~151eFdpTdiG is closed to cillow-th-isqexcess air' to bleed (spill) overboard. Figure 44 shows a typical APU installation; note the position of the pneumatic components.

Ground Rig

For normal maintenance activity it is not desirable to run the engines or APU every time pneumatic pressure is needed. Ground connections are provided for this. Figure 45 shows a typical installation located beneath a panel. On the panel there is normally operating instructions and maximum pressures. Make sure you are aware of this data. Also there is a non-return-valve (check valve) to prevent the normal system pressure from going out to atmosphere through the ground connections.

The rig is connected to the system after checking the AMM and making sure the aircraft is configured for the test. The engine is started on the rig and pressures and flow rates adjusted to the figures given on the panel/AMM.

Great care needs to be taken when connecting pneumatic hoses to these points, a s they can break under the sudden loading. This can leave large volumes of high temperature air exiting the broken nozzle. Prior to applying pneumatic pressure you must carry out safety checks on these hoses as well as flight deck safety checks.

SERVICES

Various services are provided from the pneumatic supply including cabin conditioning, anti-icing etc. These are covered in other books in this series so here we will look at engine start and various components around the engine and pylon.

Figure 45 shows a typical layout of the engine and pylon components together with details of the air starter (based on the early B747s).

LOW STAGE \ \ CHECK VALVE u

START OPEN LIGHT VALVE PRESSURE

SWITCH \ / START TOR

Fig. 45 ENGINE & PYLON COMPONENTS

Air Start Motor

The start valve will open if the solenoid is energised and air pressure is available. The solenoid is energised if Engine Ignition Switch is at GND START.

The Pressure Switch illuminates the START VALVE OPEN LIGHT in the flight deck if there is pressure down stream of the start valve.

Fig. 46 PYLON VALVE

Pylon Valve

Situated within the pylon and is electrically operated but can be manually operated for maintenance purposes.

I t regulates the pressure to 45psi but will reduce this if temperature exceeds 200°C. At 230°C the valve closes.

It acts as a check valve preventing air from the ducting going to the engine from another engine.

Warning lights come on - when valve closes. - when duct temperature exceeds 255°C.

The close solenoid is energised to prevent the valve opening if:

(a) Engine bleed air is at OFF.

(b) Fire handle is pulled.

The reverse flow solenoid is energised to cancel the check valve function and so allow air from the duct to engine start if engine ignition switch is at GND START.

Note. The engine bleed switch must be at ON if the valve is to open.

Figure 47- shows the pneumatic system for engine and pylon.-Take time to study the drawing and make syre you can understand how-the system

8 .

works. ~ I ) \

1

1 ' ISOLATION VALVE - - - I -

\ FROM APU O R , GROUND CARTS

, 1 : I I PNEUMATIC

MANlFOL D L- r

HYDRAULIC TANK

PYLON VALVE

TEMPERATURE STRUT

w

Fig. 47 ENGINE & PYLON SCHEMATIC

- 48 -

rnoodull lA-1431

That concludes this section on low pressure pneumatic supplies. Now have a go at the following exercises. The answers, a s usual, are in the text.

1. Look again a t the basic engine bleed schematic diagram (figure 24). Explain where we get air from and a t what times certain valves close.

2 . What is the purpose of the pre-cooler?

3. How is the air cooled going through the pre-cooler?

4. What is the purpose of the IP check valve?

5. What type of valve is the High Stage (HP) valve? How is it operated?

6. How is temperature controlled normally and how is controlled it if that method should fail?

i 7. If there I is an over-temperature, - - . what is the most likely -- - cause? I -

\ /' -

-- - -'\

8. Cook in ATA chapter 38 of your\rn&ntenance rnandal.-F&s on page ljlock 1 - 100 and look ho,w your prcraft controls pneumatics.

I I I / I

I I

9. Gist the various types of duct c1,aknd and explain how they &e sealed. I -- - / /

/ - 1 L - 10. what is the correct p+cedse, fpr leak checkrng a--pneumatic

siystem? , I \ 1

I 1 I I I i

I ' , I I ' 1 1. ; What sy'stems are employed to measure temperatur~? I ,

L- - // /

- / -1

VACUUM SYSTEMS

A vacuum system produces a negative pressure relative to ambient pressure which allows the pressure difference to drive components such as de-icer boots, inflatable seals, gyros etc.

De-icer boots and inflatable seals use vacuum to ensure complete deflation when selected to that configuration.

Although called a vacuum system it is not an absolute vacuum that is produced, but a pressure below that of ambient. This negative pressure can be obtained in two ways:

* By using a venturi tube system.

* By the use of an engine drive pump.

The Vedturi Tube

This is placed in the airflow close' to the fuselage, and fitted t b some small aircraft. I / I

I 1 - -

I '

P VACUUM GAUGE

I I - -

f ITUDE INDICATOR HEADING INDICATOR TURN 8 BANK INDICATOR

Fig. 48 VENTURI VACUUM TUBE

It has the advantage of low cost and is simple to install and operate. A single engined light aircraft can be equipped with a two-inch venturi, which will give 2in Hg (mercury) vacuum capacity - to operate a turn needle for example. With a n 8in venturi there would be sufficient power available for the attitude and heading indicators. (all gyro operated instruments).

The supply line to the vacuum system is connected to the throat of the venturi. The drop in pressure as the airflow passes through will cause a suction a t the throat and a suction in the vacuum line.

Throughout the normal operating airspeed range the velocity of the air through the venturi creates sufficient suction, but when taxiing, or at the start of the take-off run, or a t the end of the landing-run the negative pressure is insufficient for the system. There is a wide range of airspeed that the venturi throat will experience, so the vacuum pressure to the system will vary.

r -- - -- - For gyro-operated instruments the - gjFro-rotor - does not reach itsnormal operating speed until after take-off sopre-flight checks ofkhle instrument cannot be made for operationkl servicedbility. Due to this, limitation, the systeP1is only used for light Arkraft inskTrnent training y d limited %lying undd? knstrument weather cohd\tions,.'~ii-craft which fly through a wider speed range, altitude and weatherTi5nditions 1 r - - require a more-effective

I sourle lof supply which is independent of airspeed and less affected by atmospheric conditions. 1 1 1 I

! I I

I ,' / I I 1

I3nginei~rivgn Vacuum urn^' -

/

1 / I I J , I

1 - --- I - -

The vane type of engine driven pump shown in figure 49 is the most common source of vacuum for systems that are fitted to general aviation light aircraft. Can be fitted to the accessory drive shaft of the engine or

I belt driven on piston engines. On a piston engine aircraft it is connected to the engine lubrication system, the oil providing a seal, cooling and lubrication.

Another type of engine driven pump is the dry vacuum pump. This pump has no lubrication, and the installation requires no lines to the engine oil supply and therefore no air-oil separator or check-valve. Its principle of operation is the same as the oil lubricated pump.

The disadvantage of pump and venturi systems is that with altitude air density decreases and vacuum pressures are more difficult to achieve. With pump systems routine maintenance is required for filters, pumps, valves, pipework etc.

AIR IN AIR OUT

4 VANES t

Fig. 49 CUTAWAY OF A VANE TYPE VACUUM PUMP

Typical Pump Driven Vacuum System - - -

-

A typical vacuum system with a pump capacity of approximately 10 inches Hg at an engine speed of 1000 rpm is shown in figure 50. ~ I

SEPARATOR

HEADING INDICATOR

ATTITUDE INDICATOR

\

FILTER

TURN 8 BANK

RELIEF RELIEF VALVE VALVE

RESTRICTOR VALVE

Fig. 50 TYPICAL PUMP DRNEN VACUUM SYSTEM

The pump size and capacity may vary from aircraft to aircraft and depend on the number of gyros/equipment to be operated.


The air is drawn into the system via a filter and, via pipelines, will be directed to equipment/instruments. It will impinge onto the turbine wheels (that are part of the gyro on gyro systems) and cause them to rotate, or give vacuum to services that require it.

Various components are fitted to the system for system protection etc.

Air-Oil Separator

The oil and air from the vacuum pump are exhausted from the system through the separator in order to separate the oil from the air and return it to the engine lubrication system. The air is then vented to atmosphere.

Suction Relief Valve i

1- - - I - -

I- --

A s the pump capacity is more than is,needed for the normalroperation of the i~s~ruments/equipment, hI adjustable relief valve is included in the system. It is set at the requireh pressurk, so that any exceps negative

e is prevented by the slpring-loaded valve, allowingladr in inrdm the atmosphere and preventing the pumgf;om being overloadbd. A typical

/' valve1 is shown in figure 5 1. 1 -- - , \ \ 1 I - --

AIR \FILTER SCREEN

Fig. 5 1 SUCTION RELIEF VALVE

Pressure Relief Valve

If situation occurred where the system was subjected to a positive pressure from the pump, both the check-valve and the suction relief valve would close. This could result in the system pressurising, rupturing pipelines, damaging instruments, and causing systems to work that shouldn't (retracted de-icer overshoes would inflate for example). Therefore a pressure relief valve is fitted to vent any positive pressure to atmosphere.

- 53 -

moodull l A-1436

The purpose of the check-valve is to prevent possible damage to the instruments and operation of services in the event a reverse flow of air.

Fig. 52 CHECK-VALVE

Selector Valve

In a twin engined aircraft which has vacuum pumps driven by both engines, there is provision to select either pump due to the possibility of either an engine or pump failure. The selector valve also lincorporates a

-, check-valve to seal off the failed system. \

I

I Restrictor Valve

Fitted in front of some instrurnentSsuchas the turn and slip Edicator which operates on less vacuum &an that rkquired for the rest ofthe

1 1 system. 11t reduces the vacuum of the main line supply. 1 1 I I $

This valve is either a needle valve adjusted to reduce the vakuum from the mainline-by approximately a -b l f , ;a spring loaded reelating valve which maintains a constant vacuum for the turn indicator, unless the main line vacuum falls below a minimum value.

Air Filter

This filter removes debris from the air flowing Lo all instruments, which also have individual filters. If the filter became blocked, the result would be a reduced airflow and a lower reading on the system's suction gauge.

Most filters have a bypass valve fitted that opens in the event of a blocked filter element.

Suction Gauge

The suction gauge is a pressure gauge indicating the difference in inches of mercury between the pressure inside the system and atmospheric or cockpit pressure.

If the vacuum for the attitude and heading indicator is Sin Hg, and the minimum is 4.6in Hg, then a reading on the gauge below the minimum will indicate that the airflow is not spinning the gyro fast enough and the gauge indication is therefore suspect.

In many aircraft there is provision for the pilot to check the vacuum system a t more than one point.

Suction

The suction pressures are minus or negative pressures. As a n example, if sea level pressure equals 14.7psi then lin Hg or lpsi vacuum is equal to - 1 psi negative pressure, or 13.7 psi positive absolute pressure. If there is 3 inches Hg which is -3 psi, then as a positive pressure there would be +11.7 psi.

Typical&ystem Operation I- -. ', \

I ' 1 --\, ',, \

~ i ~ u $ e 53 shows a typical vacuum systemfor a twin engined aircrqt. This vacubin system consists of thk hllowing $omponents: 1 1 I

I 1 I - - / / ' I *, Two engine driven pumps,'., ' I 1 L - - /

I I , 1 * Two vacuum relief valves. '\ - --

' *I I Two check-valves I 1

*' / ' I ' I

I A ,vacuum manifoldi i

' I I *I d ~ a c u u m restricdoq for e a ~ h turn and bank indicator. , *i - 1'

An engine four way selecttl'on/valve. -- -- pp

/

, * I One vacuum gauge. -

* A turn and bank selector valve.

The left and right engine driven vacuum pumps and their associated lines and components are isolated from each other and act as two independent vacuum systems. The vacuum lines are routed from each vacuum pump through a vacuum relief valve and through a check-valve to the vacuum fbur-way selector valve.

From the engine four-way selector valve, which permits operation of the left or right engine vacuum system, the lines are routed to a vacuum manifold. From the manifold, pipelines connect the vacuum operated instruments into the system.

From the instruments, lines routed to the vacuum gauge pass through a turn and bank selector valve. This valve has three positions; main, left turn and bank, and right turn and bank. In the main position the vacuum gauge indicates the vacuum in the lines of the artificial horizon and directional gyros. In other positions the lower value of vacuum for the turn and bank indictors can be read.

TURN & BANK SELECTOR VALVE \ VACUUMGAUGE

PILOT'S T & B

CO-PILOT'S TURN 8, BANK INDICATO

CO-PILOTS ART HORIZON

-

LEFT VACUUM PUMP

Rig, 53 VACUUM SYSTEM FOR A MULTI ENGINED AIRCRAFT

-- - --

Air is allowed into the system via each instrument - suitably filtered of course.

A separate system not too unlike the one described above can be used to operate the vacuum side for de-icer boots. The vacuum pumps may be electrically driven and there may be a vacuum reservoir with the supply lines connected from the reservoir to a cyclic valve on the de-icer system.

Testing The System

Like all testing, the best way is to consult the manual and carry out a functional test. This would require an engine run for engine driven pumps and a vacuum/pressure supply for the system.

Leak tests are carried out much like a Pitot/static system. Instruments may be blanked off (but check the AMM) and vacuum applied slowly until the required test gauge reading is obtained, this should be maintained over a period of time.

If the system cannot keep the vacuum for the specified period, the leak (in) must be traced and rectified.

Remember, for leak testing do not use any leak testing solution such as Snoop. Any leaks in the system would suck the solution in and this may cause serious problems with instruments/components.


Answers to Self Assessment Questions

SAQ 1. Placed near the compressor so it is less likely to freeze - so if all else fails it should still be working.

SAQ 2. Blue crystals indicate dry silica gel; pink crystals indicate that they are saturated and require replacement (they can be re-activated by heating gently in an oven). If the filter becomes saturated frequently, it might be indicative that the oillair separator is full or the compressor is loosing too much oil. It requires maintenance action.

SAQ 3. Typically these pipes are in the region of 1 to l/z inch in diameter (25.4mm to 12.7mm). In most cases the pipes are fabricated from steel, usually stainless.

SAQ 4. To allow the bottles to be pre-charged if required and to allow the ground rig to be used for flap / gear retractions during maintenance.

I -

SAG 5. To keep a pressure relief valve open the compressor must continuously compress the air so it is on-load all the tiime - this takes power hhich comes from the engines. 'when a pressur{ regulator valve ports t h e air to atmosphere the compressor is working offbload - pumping air, yes, but at little or no pressure. In general, its power c~onsump,tion can be as low as 10% off-load compared to its on-load valu4 - frorh, say 10 horse power down to 1). I

I I -- 1 I I SAQ 6. The single and double acting implies the way actuator is

returned to neutral. In the double acting actuator it is retLrned to neutral by using the reverse selection. -

- - -- I

SAQ 7. A single acting actuator could be prone to 'blow back' from the flight effects (on a simple hinged flap system), meaning that it would have to be kept pressurised to keep the flaps deployed - though on some aircraft this might be used.

SAQ 8. It would be either amber or red depending on the level of importance the system has. It would located within the pilot's line of vision or possibly on a master fault panel.

SAQ 9. Two reasons really, firstly switching an earth is less likely to cause a spark so the contacts of the micro-switch are less prone to erosion and radio interference is reduced. Secondly all lamps on the flight deck have got a common supply which means that they can all be dimmed together from one dimmer switch.

SAQ 10. Lowest bleed air occurs when the engine is running at idle. The crew would select the turbo compressors on during taxi. They would leave them running until the gear is retracted, and would start them again on the approach to landing.

SAQ 11. A jet engine relies for its thrust on the high speed air leaving the rear of the engine. A turbo-prop engine uses most of its energy (about 90%) to drive the propeller so only a residual thrust of 10% is obtained from the jet efflux. Affecting this efflux in any way (scooping some of this away for anti-icing for example) has very little effect on the overall efficiency of the engine. Besides the problems of the high velocities and temperatures in the efflux of a non prop jet engine any disturbance of this airflow would have a more marked effect on its performance.

Gasses leaving any jet engine are toxic and corrosive so they are best ducted out of the aircraft as soon as possible. The ram air leaving the heat exchanger {unless there is a leak) is clean (and hot).

SAQ 12. If a restrictive type valve system was placed down-stream in an engine driven compressor system then as the restriction started it would load the compressor and cause it to consume more power from the engine.

I So if demand air is less than supply then the excess is simply allowed to - - spill-to-atmosphere. [-- - - , I - - --

r2 I \ - - _ _ - .

'\

~ a ~ ~ i n k s from the jet engink a re different. If demand l dr is less than supptyi then any excess air Ldn be used by the engine so a rystrictive

l 1 systeinlis used - air not wantdd stays id the engine. I

I I 1 I

/ I 1 - ' /' 1 '

SAQ 13. Initially it could be considered that the system wb?ld fail,,but this i is not the case. The torque motor wi8 Btop moving, but [ the i r is still 1 1 coming off thefingine. The redlators qave already broGght the pressure

down tp 55-6Spsi and this will 'back off tide bleed signal arid close the HP

\__ 1 -/' 1 valve1 at a slikhtly higher pre$spre than $e torque motdr bould do. This

tyge of redundancy is common irelectronic systems. - , - L -- - -- - -- - I

SAQ 14. Consider the valve itself, it is sprung loaded closed, pneumatically opened - that pneumatic pressure is coming down the sense line (that is leaking). There will not be as much pressure in chamber B so the valve will

i stay further closed than before. The effect of this is to reduce the system pressure in the pneumatic control due to the fact that the apertures are set. In the digital control, the Air Supply Unit will see the reduced pressure and apply more sense pressure to open the valve. It will sense the failure (and report it.

SAQ 15. After start the temperature would rise above the normal operating temperature, possibly getting close to the over temperature limit. But it would eventually stabilise. As the engine is throttled forward (on take-off), and the HP valve closes the temperature would drop well below the normal operating temperature. During cruise it would appear fine but might re- occur after landing on the taxi in, as the HP valve opened again.

SAQ 16. Ducting is usually made of thin gauge metal and is susceptible to damage. Care with handling and installation is required. Consult your AMM but as a general guide:

* Always replace seals when ever ducting is changed or joints disturbed.

* Do not over (or under) torque the clamps. * Ensure attachment tie rods are adjusted correctly on installation. * If shims are fitted these need to be fitted and adjusted correctly. * If ducting clamps are near electrical cable runs, great care needs to

be taken not to trap the cable or allow the clamp to come into contact. (Cables should be checked to see that they do not come into close proximity with hot ducting - consult the AMM on correct cleating/ cable supports etc) .

SAQ 17. Having the same excitation voltage means that any fluctuation in that voltage is felt at both coils and will not effect the reading on the instrument. -

CONTENTS

Page

Ilomestic water supply systems - general The domestic water supply Toilet Systems

The removable type Re-useable liquid flush type The vacuum flush type system

Toilet cubicles C~or~osioncontrol I - -

)Adti corrosive measures-, - -

~ a l l e $ s 1 ~evisi'on exercises 1 I I

~ n s w k r s to SAQs I - /

- \

I - , ' \

1 I I \

I ' 1

1

, ' I

I 1

' I I ! , I j ,

I - /

DOMESTIC WATER SUPPLY SYSTEMS - GENERAL

Have to meet the requirements of CS25 (was called JAR25, then EASA25, now called CS25) (CS = Certification Specification) in that they must not constitute a hazard to the aircraft and all replenishment connections should be so designed that misconnection of water services to any other systems is not possible.

Water is used on aircraft for many reasons, and as such can have different levels of water quality. Water is used for:

* Drinking. Has to have a high standard of cleanliness and quality control.

* Washing - using wash-hand basins (sinks).

* Galleys. Water for galley sinks, soft drink dispensers and makmg hot beverages (tea, coffee etc) .

*ILTojJets. The flush system-fEthe vacuum type toi1ets.r -

\ r , - --

I I I ' 1 1 1 * IHhrnidifiers. For h u m i d q n g the air for air conditioning.

I I

I 1 I * cQoling heat exchangers. '&,r conditioning heat exchanger units may use

water for additional cooling.-This h$y not come from the @mestic water ,sdpply but taken straigh 1 frlomthe water extractor. 1 /

I 3 /

\

I ' 1 1 , * Engines -/?ater/rnethan&l ibjection for'jet engines. ~ h d ? system may have

kts own tanks and supply/system. / I

/

i- - I \ - //

Theifast three are not part of tj.$sbookas-they are covered elsewhere in the LBP books but have been include to give an overall picture.

THE DOMESTIC WATER SYSTEM

The drinking water system (sometimes called Potable Water) consists of:

* Storage tanks. Sufficient water has to be stored for the flight duration. The tanks and system must protected from icing, if the is a risk. A system has to be provided to allow draining, flushing and refilling of the tank(s). A contents system is provided for the crew.

* Heating. The water has to be heated to the right temperature for hot water taps (galley and toilet sinks) and has to be hot enough to make drinks with (galleys). Note that a tap is called a faucet in American aircraft manuals.

* Cooling. Water for drinking fountains require the water to be cooled.

rnoodull I A-1 445

* Distribution. There needs to be enough pressure to force the water to all the taps, heaters, beverage makers etc. This requires the tanks to be pressurised and a system of pipe-work to carry the water to the required cabin areas.

* Contents indication. This needs to be available to the engineer when filing the system and also the cabin crew need to be able to see the contents of the tanks during flight.

* Anti-frost system. To prevent freezing of the water in the supply pipes electric heaters are placed around some pipes.

* Removal of waste water (grey water). The waste water must be removed from sinks, drinking fountains etc to outside the aircraft.

Fig. 1 TYPICAL WATER SUPPLY SYSTEM

Figure 1 shows a schematic of a typical system. Study the drawing carefully and note the following: the location of the tanks; the provision for filling and draining; the supply to the toilet wash-hand basin, galleys and drinking fountain; the tank pressurisation system to include APU and engine bleed via an air compressor and the various indicators.

- 2 -

moodull l A-1446

Storage Tanks

Figure 2 shows a typical water tank installation. May be made from stainless steel or glass re-enforced plastic. Some aircraft may have just one tank, others will have more than one. The tank/s may be positioned vertically or horizontally.

Tanks are usually installed under the passenger floor in the freight bay regon. In some aircraft they may be positioned elsewhere such as in the space above the cabin ceiling. If in the freight bay CS25 stipulates that they must be separated from the freight bay potential fire region in some way and this is normally achieved through use of fire blanking material such as glass fibre.

Fig. 2 TYPICAL WATER TANK INSTALLATION

The tank shown has re-enforcing bands to strengthen the tank and provide strong points from which the tank can be secured to the airframe.

In cases where the tank is suspended from the floor beam structure, the fmtures may well be some form of structure fuse connection, that allows the tank to break free at times of high deceleration (forced landing, for example). This is to prevent the weight of the tank collapsing the floor structure and pulling passengers into the freight hold. Care needs to be taken to ensure the correct bolts are used.

Tanks will also have connections for distribution, fill (and overfill), draining and air pressurisation. Notice also that it has an attachment pad and some form of quantity transmitter, providing the necessary information to cabin crew and ground crew.

rnoodull l A-1447

Tanks normally come as a pre-inspected unit from overhaul, the panels at the end are for inspection purposes.

SAQ 1

You have to carry out a detailed visual inspection of the structure behind the tank. Would you consider it necessary to remove the tank and if so list the procedure for removal and refitting?

The Distribution System

The distribution of potable water is accomplished by use of pipework (some flexible, some rigid - check your aircraft) that comes from the bottom of the tank. The tank pressurisation system forces the water up riser pipes and into the cabin area system. --- - -- ,

- _ 1

In case of a burst pipe the flexible pipeline, may be fitted into anj duminium shroud. When considering leak checks - they need to be dond b~fore the shroud is fitted. 1 I I

1 1 I I

The outside of the flexible pipe is usually fabric reinforced to brovide additional strength and give a small measure1 of protectipn against freezing. ~ h e _ r e - - freezing is a problem, ribbon heaters are fitted. I

1 t,

I 1 I

THERMOSTAT

l T 5 V AC

Fig. 3 RIBBON HEATERS

Figure 3 shows a typical ribbon heater and its electrical supply. Note the single phase ac supply via a C/B and thermostat. Note that the four heaters are arranged in parallel.

The freight hold of an aircraft is usually heated, but if, in the area of the water tanks, it is not, electric heaters are fitted to prevent freezing of the tank and/or pipes. Ribbon heaters provide a constant, low power heat source. There is no switching and as aircraft power is applied the heaters come on. They are sometimes ganged with the drain mast heaters for power supply purposes.

To prevent overheating a thermostat is fitted.

Removal/refitting of these pipes is the same as any other system pipe-work and is covered in module 7 of the EASA part 66 syllabus.

Often, a t the system connection between the system and the galley/toilet modular units, a quick release (self-sealing) fitting is used to allow easy removal of the toilet/ FCey unit. Sometimes the distribution system has anisolation valve in the

f - cabin and the Cabin Services ~ i r ~ c t b c c q \ i f required, shut2off the water to the entire aikcraft. (Boeing 707/720 and the early 747 had this facility).

I I I I

Fig. 4 TYPICAL HOT TAP WATER HEATER

Water Heating

Each wash hand basin has its own water heater in the line to the hot tap. Figure 4 shows a typical water heater installation housed in the area under the sink cabinet.

Each water heater has a capacity of about 3 pints (1.71) and heats the water in a cyclic action. The water tank has the following equipment fitted:

* A O N warning light. * A ON/ OFF control switch. * A n overheat reset switch. * A pressure relief valve. * Supply and feed connections.

OFF-ON LIGHT

- - *.. r - - 1 I 115VAC I

OVERHEAT OFF SWlTCH

I I &,

I 420 WATT H ~ A T CYCLE

HEATER SWIT'H I

L - - - A CIRCUIT HEATCONTROL -- - . -

BREAKER PAN EL

SWITCH I

I

I

Fig. 5 WATER HEATER C~NTROL CIRCUIT

Figure 5 shows the water heater control circuit. Study it fora moment. When the heater is switched on (1 15v single phase ac) the light will illuminate and the heater starts to heat. The current flows through the cyclic switch (a thermostat) at the bottom of the tank which is closed and so power is allowed through the heater elements.

As the water temperature rises to about 50°C (125°F) the cyclic switch will open and stop the power to the heater elements. Notice, however, that the light will remain illuminated. If the cyclic switch should malfunction, the temperature of the water will continue to rise, as the heater elements will remain powered. To prevent an overheat condition and possible fire another switch is installed (by regulation), called the overheat switch. At about 87°C (190°F) the overheat switch will open and latch. This will stop the power to the heating elements and the light will go out.

The heater will require resetting - by pressing the reset button. The reset button i s located on the tank and is usually covered with a rubber bubble. The water heaters are powered all the time power is available on the aircraft.

rnoodull lA-1450

If the system has been drained and the heater C/B's have not been isolated, the overheat switches will operate and will require resetting - after filling the system of course.

The possibility of fire is small, but the water heater is usually protected with a 'wax stat' fire detector and fire bottle and care needs to be taken during maintenance on the water heater that this nozzle is not dislodged. (More information on the fire extinguisher is available in the book in this series entitled Fire Protection.)

The on/off switch allows the heater to be locked out of operation for whatever reason.

The over pressure relief valve must never be touched. It is a core plug type of valve and will, should the pressure rise to about 140psi (965kPa), blow to relieve the pressure in the water heater. This will cause a water leak, which would need isolating at the toilet shut-off cock. If the water heater pressure relief valve has operatedzthe heater requires replacement,, - and you will need to investigate the

\ -- - cause! I

I \ \

I 1 Removd, of the water heater, and Kor that matter any component in the toilet area, doesn't require a full system dr- as the:tqilet has its own slo~)cock, usually locate? in the sink cabinet areal ~aving'isolated the water heater by/&ipping and tagging the appropriate C/B, the h-tcr-isremoved in much the -e/way as any system component. The basics of this covered in SAQ 1. I 1

Fig. 6 WATER COOLER CONTROL CIRCUIT

rnoodull lA-1451

Drinking Water Cooling

Water for drinking fountains will normally require cooling. Figure 6 shows the circuit diagram of a cooler.

200v three phase ac supply is used to drive a motor which when the water temperature rises above 10°C (49°F) will drive a fan to blow air across the thermoelectric units, which uses energy change entropy to reduce the temperature of the water. The motor is controlled via a control thermal switch similar to the heater circuit already described. The transformer rectifier (the 2 delta windings in the drawing) is used to change the 11 5v ac into 28v dc for use by the thermoelectric units.

A s the system is removing heat from the cooler, so this heat energy is dissipated by the motor fan. Should the control switch fail the unit might overheat so an additional switch is added that breaks the circuit at a higher temperature.

- - -

1 I , I

Contents Indication I

I The contents of the water tanks need to be relayed to both the cabin crew for use during the flight and the ground ctew for recharging purposes. l ~ r o u n d crew seldom dse the contents gauge at the fill pdint now as reguladions regarding drinking water dictate that the system rnustibe full for departure. -

1 I \ I

Fig. 7 WATER CONTENTS INDICATOR

This leads to the practice of filling to overflow and effectively negates the need for the contents gauge - but it is still there.) Figure 7 shows the location and style of an indicator.

Older aircraft (and sometimes as a back-up on modern aircraft) have a simple sight tube on the side of the tank to provide contents indication.

Figure 7 shows an indicator and figure 8 shows what is now a fairly common style of contents indicator system. It consists of a float with a magnet riding up the outside of a sealed tube containing several reed switches. A s the magnet passes a reed switch it closes and this signal is sent to the signal conditioner for onwards transmission to the cabin crew.

In the case of the latest aircraft where the cabin systems are provided centrally through a digital Cabin System Control Panel, the water contents information is passed to the cabin file server and converted into digital signals for display on the CRT/flat screen display on request.

I 7 - - - - -

L T - ' I - , I -

l 1 I I I

REED SWITCHES

REED PUSH BUTTON SWITCHES INDICATOR

(IF FITTED)

Fig. 8 WATER CONTENTS MEASUREMENT SYSTEM

rnoodull l A-1453

System Servicing (figures 9 and 10)

Replenishment of the water is carried out from the water service point usually located under the aircraft fuselage, near the tank location and away from any toilet servicing panels. To fill the system the water is connected to the fill connection and the fill valve is pulled to turn the valve to the service position. Water will flow under pressure from the tank maintenance rig to the tanks and they will start to fill.

A s the tanks reach full the water is forced up the stack pipe and through the overflow and will start to flow from the overflow port on or near the service panel.

At this point the filling operation is stopped.

Fig. 9 POTABLE WATER SERVICE PANEL TYPICAL LOCATION

WATER DRAIN

SERVICE PANEL

]REPLENISHING POSITION^ [NORMAL FLIGHT MODE - - --

I --

1-

I / -

- -

I ' Fig. 10 TANK FILLREFILL SYSTEM S C H E M A ~ C I

I

' I I 1

I I ' I

1 1 I

l 1 I I

~ r a i n i h ~ is also carried out from the wafer Service panel. ~ i r s t the filljhandle is pulled; which rotates the fill/vent valve and vents the tank ( e n s u r e m a t is un- pressuribed). The drain valve is t h p t a t ~ d , which, a s it is lbcated a t the bottom

I of the sistern, allows the water to drain b$\ g?avity. This wateb Should not be re- ' ,

I I used. I 1 I I

i I j ; I ~ On lmLd\aireraft, there may well hie-tw-o,or more drain points under the fuselage,

r' eg one-in front-of the wing section-anddone, behind. This all&s for the fact that sometimes the pipe-work has to rise to get over the wing centre structure.

The water system is susceptible to bacterial contamination. We can help to prevent this by having a high standard of cleanliness during servicing operations. We need to be aware that this is drinking water and we must apply high hygiene standards. In addition, if the system has been disturbed during maintenance, it has to be sterilized.

The sterilizing solution (typical a chlorine like Puregene@) is added to the flow of water during a fill operation. The tanks are then pressurised and all taps and toilets are operated to ensure that the sterilised water flows in every part of the system.

You can tell that the solution has passed through the tap when the flow of water from it turns a milky colour. It also has a distinctly chlorinated smell. This solution is left in the system for some time, then fully flushed out.

Note. When filling always ensure that the water has come from a known drinking water source and that the replenishing cart is exclusively used for drinking water.

Tank Pressurisation

Compressed air is fed into the tanks to force the water up the risers to the toilets and galleys. This supply of air will either be from the aircraft pneumatic supply, or a purpose built compressor. Sometimes on large aircraft both are used, with the compressor providing backup for the main system (figure 11).

- FILTER INLET

AIR - AIR

COMPRESSOR

piiJ=@ AIR - CHECK

- FILTER VALVE

I I

AIR CHECK FILTER VALVE PRESSURE

PNEUMATIC SYSTEM / RELIEF CROSSOVERDUCT - -- I - ,VALVE

I \

I , I

I

I PRESSURE SWITCH

- -

Fig. 11 WATER SYSTEM PRESSURISATION

The water pressurisation system consists of:

* A filter, normally a paper throwaway type, to prevent any airborne particles entering the water tanks.

* A pressure regulator to control the air pressure to an acceptable pressure (about 30psi - 260kPa).

* A pressure relief valve to relieve the pressure should it get too high.

* Possibly a pressure switch that switches the compressor on when the system pressure drops below a pre-set trigger level.

* A compressor interlock switch (not illustrated) fitted to the ground service panel door, that will turn off the compressor when the servicing door is opened. Usually a proximity type switch.

Notice also in figure 11 that there is a number of Non Return Valves. Also take note of the riser loop, this is fitted to prevent water from entering the compressor system from the tanks when the compressor is off.

Water Disposal (Grey water only)

Drain water from the sinks (galley sinks and wash hand basins) is dumped overboard through a series of pipes and heated drain masts. Figures 12 and 13 shows a typical sink waste disposal system.

- 1 --- _ - I r -

As thk sihk plug is lifted - u s u ~ ~ l y , byasmall lever, the water flows down the pipe to the ldqain mast. The water is then allowed',to flow overboard. This presents no environmental issues, as the wdter is moderately clean and soon disperses in the

I atmosphere. (Toilet waste water, [Blue water]' is never dispose'd of overboard, but kept on-board in tanks to be emptied-when on the ground.) 1 Lp - 1

I -- -

When thle plug is lifted the line is open to ytyosphere. A s thk a i r ~ a f t is usually pressuri$ed, the,$ressure differchtial will ass~s t in the removal of the water.

I I

FOUNTAIN-

SINK DRAIN LINE

DRAIN NOISE

GREY WATER

Fig. 12 SINK WASTE DISPOSAL SYSTEM

rnoodull l A-1457

This system presents a couple of problems in that the noise generated is high and the loss of air from the cabin causes inefficiencies. A poorly fitting plug, for example, will cause a noise all the time and lead to fuel penalties (more air being bled from the engines to make up for the leak).

The noise problem is largely overcome by putting a muffler (very similar to a car exhaust silencer) in the outflow line. We must also ensure that the plugs fit correctly and that the rubber bung is not perished.

The loss of air problem is harder to overcome and one recent solution is to provide a Gray Water* drain valve (figure 13).

* The term Blue Water is often used to describe liquid from the toilets and Gray Water, liquid from the sinks. Blue water waste is kept on-board, gray water waste is dumped overboard. Gray water is a Boeing term and hence the spelling.

On the ground the valve opens completely. When the aircraft becomes airborne, the gray water valve closes andprevents all water from exiting thepGrcraft.

- 1 ' ! I i

I \

, ISOLATlONl ISOLATION VALVE

I

FWD FWD

FORWARD SYSTEM DRAIN LINE

t

GRAY WATER

DRAlN MAST (2) -.)

Fig. 13 B777 SINK DRAIN SYSTEM WITH GRAY WATER VALVE

A pressure switch in the drain line from the sinks will establish when there is 1.8m head of water on top of the valve which then opens and allows the water overboard. 1.8m puts the water just below the plug hole, and means that several sinks can be emptied before the valve is opened to allow the water overboard.

Drain Masts

The sink waste passes overboard through a drain mast. Figure 14 shows a typical example. The drain mast ensures that the water is removed from the aircraft and not allowed to blow back onto the fuselage where it would freeze on the airframe structure and become a hazard.

The ambient low temperature would also freeze the water inside the drain mast which would cause it to become blocked.

Fig. 14 DRAIN MAST

To prevent this the drain mast is electrically heated. Two temperature settings are involved, one on the ground (low) and one in the air (high) (operated by the weight switch). The lines down to the mast are also heated - by ribbon heaters - also to prevent freezing and blockage of the system.

CAUTION. These masts can get very hot and are a burn hazard. Do not touch.

rnoodull l A-1459

Drain masts are susceptible to being damaged and put out of alignment both on the ground and in the air. It is very important that they are inspected for damage and correct alignment on every walk-round inspection.

A second issue that can arise from drain masts, is that if the heating circuit fails, the water can freeze and cause damage to the composite fairings that they are often mounted on.

All structure around and behind drain masts should be inspected for debris damage, corrosion (metal), water damage (composite) and security. If external frozen water is suspected then all structure and engines etc aft of the mast should be inspected for impact damage.

TOILET SYSTEMS

Unlike the grey water waste system toilet waste (or blue water) must never be deposited overboard. Besides beingunpleasant it presents a health hazard to all who are below the aircraft's flight path; So much so that se;ei-e fines are being imposed on operators whose aircraft have had blue ice' found on them, or blue ice deposits having fallen from them. (The term 'blue' comes fro+ the cblour of the chemical used in the toilet system!)

/ I

- Toilet systems can be classed into three main groups. These &re:

-

1 I I

* Removable toilets - often idled 'Elson dtyle'. I I

* Re-usable liquid flush type. * Clean water flush type, or vacuum flush type.

- -- -- -

THE REMOVABLE TYPE TOILET

Sometimes referred to as the Elson type and consists of a container (stainless steel or composite material) within a toilet unit on top of which is a toilet seat. Inside the container there is a small amount of a chemical mixture (such as Raquasan) that acts as a disinfectant and helps with the smell problem.

Servicing this type of toilet is simple, and involves removing the toilet container (sometimes via a servicing panel from outside the aircraft, sometimes from inside by removing the decor paneling around the toilet), emptying it (into the normal airfield toilet system), cleaning, recharging with a small quantity of fresh Raquasan, replacing and securing. The container itself is clamped to the floor to prevent movement and spillage in times of turbulence.

The area within the decor paneling, the floor and the structure must be kept clean and care must be taken not to spill any of the liquid on the aircraft. Urine (uric acid) is very corrosive.

rnoodull l A-1460

Fitted to smaller executive jets, often within a plush decor paneled unit within a lavatory unit. The container may have a lightly sprung loaded (closed) stainless steel flap on the top which must be aligned with the toilet seat.

RE-USEABLE LIQUID FLUSH TYPE

This has a flushing system which uses the existing liquid (disinfectant and urine) within the container under the toilet. A typical toilet unit is shown in figure 15.

Most large commercial aircraft manufactured up to about the late 1980's have this type of toilet system.

The waste tank is located directly beneath the toilet pan in much the same way as the removable toilet was. The tank is large and, typically, can hold u p to about 40 gallons (1801) of waste (on long flights this can become full which causes toilet cubical carpets to become wet with possible structure contamination). Sometimes the tank is a double tank anCzill-hTve-two toilet pans in two separate cubicles. The tank% usually fabricated frofiglZiis ,re:enforce ldastic$t [can be stainless steel, q d has metal tie down points. 1 ,

I 1

I , I I

' , I I THREEPHASE

PUMP MOTOR

FLUSH LlNE

TOILET FLUSH LlNE

Fig. 15 TYPICAL TOILET UNIT

- 17-

rnoodull l A-1461

Local repairs to this tank are often not possible due to the build up of waste over time, and as such satisfactory repairs can only be carried out at an overhaul facility.

Figure 16 shows a schematic view of the toilet system. Note that the tank has a soil pipe connection underneath to connect the tank to the waste servicing point. A dump (or drain) valve is located on the bottom of the tank and is operated by Bowden type cable from the toilet-servicing panel on the underside of the fuselage.

Fig. 16 RE-USEABLE LIQUID FLUSH TYPE TOILET - SCHEMATIC

- 18-

rnoodull l A-1462

The tank is filled with a 'pre-charge' which is a small quantity of Raquasan or similar solution. (The actual solution and quantity is specified in the AMM chapter 20). This is added to the tank from a connection on the servicing panel, through pipe-work and a non-return valve in the ground flush line.

The toilet bowl(s) is/are bolted onto the top of the tank usually with captive nuts. The unit has a separator (a flap) that keeps the waste in the tank from being visible. Both the separator and the bowl itself are fabricated from stainless steel.

There is a flush ring in the bowl connected to the flush motor, which allows flushing of the toilet bowl.

The waste tank has a fume extractor system connected to the lav7 and galley vent system to outside.

Flushing is achieved using a fixed speed three-phase motor, fitted outside the tank and connected to the pump by a shaft. It alternates rotation direction on each-flu* cycle to help prevent blockXge.~~ound the pump thereis a filter and separitqr mechanism that separates thk\solids before allowing the motor to pump reasonably debris free liquid to thk flush ;ink. I

I

I

I

1

FLUSH LINE

WASTE DRAIN TUBE DRAIN VALVE

Fig. 17 TYPICAL TWIN BOWL WASTE TANK ARRANGMENT

The flush cycle is initiated by pressing the flush switch (lever, mounted on the cubical wall). This sends a signal to the timer control, which runs the 3phase ac motor for about 10 seconds. A s the motor turns the filter screen is rotated against the stationary wiper blade, thus removing any large particles of solid material and keeping the filter screen clean.

FLUSH MOTOR

\

PUMP PORT

CONNEC

PLATE

WIPER BLA

PERFORATED FILTER BASKET

/TOILET FLUSH SYSTEM I -

- - - -

\

FLUSH MOTOR ELECTSUPPLY 3 PHASE FLUSH MOTOR

1 I - - ,

- , - -

SOLIDSTATE -

IIF- CIRCUIT -

L- ---------.------

TIMER

m J

TOILET FLUSH CIRCUIT DIAGRAM^ HANDLE

Fig. 18 FLUSH MOTOR & CONTROL SYSTEM

The pump impeller at the same time draws fluid through the rotating filter and into the flush line. This fluid is moderately debris free (particle size about 300 microns) and the chemical content in the toilet is high.

At the end of the flush the separator (flap) at the bottom of the bowl returns to the 'up' position by light spring pressure.

The logic timer unit will reverse the direction of the motor on alternate flush cycles; this helps keep the filter screen clean. (The relay operating the ganged switches of the two phases in figure 18 is the reversing mechanism).

CAUTION. Before operating the flush motor for maintenance purposes, make sure that there is sufficient pre-charge (liquid) in the tank to prevent the motor from running too fast and overheating.

SAQ 2

After a flush cycle the motor continues to run. List your actions to rectlfy the fault and list the possible causes.

Servicing

Mayrbesplit into two main acfitrities, that which is carried out outside the aircraft j -.

and chat which is carried out qsidetfie aircraft. 1 \ i

Drainin;, flushing and refilling is carried but outside the airdraft from beneath the fusela'ge and toilet cubical maintenance is carried out from Aside the toilet cubic& itself. We shall deal wit4 the drdi.n?hg/flushing/reflldng of th6 system now and the cubical maintenance will lbedealt'with after all toiled systems have been covered.\ \ '\ I -

I I ' I I I

I I I 1

I FLUSH PORTS /

Fig. 19 TYPICAL TOILET SERVICING PANEL

For operators of large aircraft the inside and outside maintenance is normally carried out by contractors, but never-the-less, when the aircraft flies, it does so on the signature of a licensed engineer. So whether you actually carry out the servicing yourself or not you will need to know how it is performed.

Draining/ flushing/ refilling is carried out from a panel at the bottom of the fuselage. Figure 19 shows a typical toilet-servicing panel. There will be one of these located beneath each tank or one fitted to service several adjacent tanks. Figure 19 shows one for a two tank system. Note the two flush ports; the two drain valve handles (connected to the drain valves) and the single drain connected securely to the drain line of the ground servicing trolley.

Normally the tank is drained and re-charged after every flight. To empty the tank proceed as follows:

1. Always wear protective clothing and ensure anti-tetanus jabs etc are

- - up-the-date. Check the operators/local airfield instructions for the anti-infection injections tKat are required. Make sureyou have no open wounds or scratchesT

I

2. Refer to the AMM. I

I

3. Open the toilet se9cing panelusing the quick release buttons (check that there is no blue waterqice). The drain covdr is removed from the waste tank connector (check again that there' is: not blue liquid). On some aircraft a donut plug (safety plug) is removed from the pipe and the drain line from the waste tank servicing trolley i s connected to the waste outlet. If at any stage during this operation fluid leakage is found then the cause must be investigated and the fault rectified.

--

4. With the waste trolley drain pipe connected the secondary valve can be opened (on some aircraft the donut plug is operated by an external lever). Sometimes this takes the shape of a donut plate, but more often it is a separate valve in the outlet line between the waste tank main drain valve and the outlet. The appropriate tank drain valve handle is now pulled and this (via the Bowden cable to the top of the tank) lifts the drain valve and allows the waste material down the outlet and into the tank of the servicing trolley. Some service vehicles have a suction facility to aid in the removal of the waste.

5. Once the tank is empty and before the main drain valve is closed the tank is flushed by applying water pressure to the fill connection of the appropriate tank. This water travels up the pipes and to the ground fill connector on the tank top. From here it is passed into a perforated pipe that sprays water around the inside to clean the tank.

6. The main valve is now closed, the drain pipe disconnected and a pre- charge (a mixture of water and a strong disinfectant) is added (typical quantities - 7 gallons 13111). Check the actual amount in the AMM.

7. Check for any leaks.

8. Close all valves, donut seals etc.

9. Record all work and sign.

The toilet system is now 'serviced' and ready for the next flight. Prior to this operation it is a good idea to have the toilet bowl cleaned first by the cleaners when they "service7' the toilet cubical.

SERVlCE PANEL

1 * (

Fig. 20 THREE TOILET SYSTEM TO ONE SERVICE PANEL

Periodic maintenance includes filling the tanks completely with a strong cleaning solution, checking for leaks, operating the flush motors and draining. Any leaks must be rectified. Any blockages must be cleared and all motors etc function tested. Unserviceable components are changed.

f I

rnoodull l A-1467

I 1 t

L J I

I

I I

TOILET TOILET : [ lAssy '

I- > I

\ ' / I I I .-. I

1 L- , - --

\ I

I

I -

I 1

' I

I (

\

COMMON VENT FOR -

-- \

,

I

I , ,' 1

GALLEYS AND--- LAVATORIES

- - . I '

1

I ! I I

One of the main problems with this sort of toilet was that the waste line connections and plug tend to become perished with being in contact with the chemicals etc). This can cause the contents to leak, the situation being made worse by the pressurisation effects in flight.

The whole area must be carefully inspected from the waste tank down to ensure that there are no leaks. The chemicals are strong and are very corrosive particularly to aluminium and its alloys.

In recent years the number of plugs and secondary sealing devices have been increased and you may well find yourself performing a vacuum test on the waste line to check its ability to seal.

A simple test to check that the secondary sealing valve is leaking (or not) (the one between the waste tank drain valve and the drain line outlet) is to add a quantity of clean water to the tank. Close off the secondary sealing valve and open the main tank drain valve - with all other caps and seals open/removed from the drain pipe. !The secondary sealingvalve - should not leak. If it doecit must be replaced.

' I I \

I Afterwards close the main tank drain valve, qpen the second& sealing valve, allow the water in the pipe to drain and check that the main tank drain valve does not leak. -- /

- ' I

Of course, this will all be specifidd in the AMM. -

' I 1

Figure 21 shows a diagram of the waste connection point. I t is designed to be Murphy Proof in that if the safety plug (donut seal) is not in place correctly the interface levers will not allow the Drain Cap to close. If the drain cap will not close the panel will not close either. The open panel on the underside of the aircraft would be easily seen on any walk-round inspection.

A problem with these toilets is the fact that they tend to spill/overflow during flight (particularly a long one) and the fluid works its way down onto the toilet underfloor structure. This area requires special attention and anti;corrosive treatments and often requires repair or replacement. You should take extra care whilst inspecting these areas. The same corrosion problem applies to galley areas - though to a lesser extent.

- 24 -

moodull l A-1468

RFERENCE LEVERS

WASTE PIPE

SAFER PLUG

I , 1 1 I - - /

Fig. 2 1 TOILET' ;WASTEX+PS & SAFETY PLUG ,

I 1 '\ \ I I - '

Most l ~ g e ~6Wmercial aircraft-built-sin-@'the late 80's uselthistype of toilet.

The system uses vacuum to drain the toilet bowl assisted with a little (clean) water. The waste from the toilet is not stored beneath the unit, but remotely in a separate large tank. This tank (or tanks) is typically located in the cargo hold.

Figure 22 shows a schematic of a typical system. Note the vacuum supply to the tank obtained from a vacuum blower, which automatically cuts out at altitude (typically 16,000ft) and differential pressure takes over through the one-way valve. Note also the tank level sensors, pipe-work to the toilet bowls and the water supply from the potable water system. The anti-syphon valves prevent water syphoning back to the water system.

Suction to the tank is via a separator to prevent liquid etc from being sucked out. This applies a vacuum to the tank/s and pipes up to the flush valves. When the valve is opened (electrically) the vacuum, assisted by a small quantity of fresh water causes the waste from the toilet to pass down the pipe to the composite tank.

- 25 -

rnoodull l A-1469

FLUSH FLUSH SWITCH ANTISIPHONE CONTROL UNIT

FLUSH VALVE

FLUSHING - WATER VACUUM BLOWER SUPPLY SHUT-OFF VALVE

VACUUM BLOWER

ALTITUDE PRESSURE SWITCH

LAV INOP LIGHT ATTENDANTS' PANEL

- - - WASTE SERWCE PANEL

- _ I I -

Fig. 22 VACUUM TOILET WASTE SYSTEM (~320)

FLUSHING - WATER VACUUM BLOWER SUPPLY SHUT-OFF VALVE

VACUUM BLOWER

ALTITUDE PRESSURE SWITCH

LAV INOP LIGHT ATTENDANTS' PANEL

I

TANKS

STRUCTURE

Fig. 23 TANK INSTALLATION - EXAMPLE

Note the rinse and drain pipework, the function of which is similar to that of the re-useable liquid flush type system.

- 26 -

rnoodull l A-1470

Special large diameter stainless steel pipes are connected to the tank bringing the waste from the individual toilet units. There are usually several tanks and the toilets they serve are split u p around the aircraft so that a full tank doesn't lock out an entire section of the aircraft. The level sensors give a "tank full" indication.

On some aircraft there are two vacuum connections to each tank. One is connected to a vacuum blower that provides vacuum on demand. The other is connected to atmosphere and serves two purposes. Firstly it allows the blower to vent somewhere and secondly above 16,000ft the cabin differential pressure is used to provide the vacuum instead of the blower.

Figure 23 shows an example of tank location below the aircraft floor and figure 24 shows a typical toilet bowl unit. It is made of Teflon lined stainless steel. Note the anti-syphon valve, flush valve, rinse water valve and flush control unit. The drawing does not show the stainless ring that fits just above the toilet lip. This is to prevent the complete sealing of the unit with a person sitting on the toilet -

, designed as a safety measure during the flushing cycle, when the vacuum

I I

/

SUPPLY PLUG

Fig. 24 TYPICAL TOILET BOWL UNIT

Operation

The operation of the system is initiated by operating the flush handle which closes a switch. (The handle, incidentally must be located out of reach of the passenger whilst they are sitting on the toilet). Figure 25 shows the run cycle of the system.

- 27 -

rnoodull l A-1471

FLUSHING HANDLE

VACUUM BLOWER

WATER VALVE

FLUSH VALVE

NEXT CYCLE DELAY

SECONDS

Fig. 25 FLUSH CYCLE

- - - -

I ,

- -

The flush handle is pressed (momentarily). The water valve opens and sends a squirt of high pressure potable dater via the anti-siphon tubk around the flush ring to assist pan cleaning. The anti-siphofi tube prevents a& possibiliG of the water getting back into the drinking water system. Now the flush valve opens and the contents of the pan are sucked out and down to the tank. After 2 seconds the flush valve shuts, but the blower continues tp run for a short time. A repeat flush is inhibited all the time this is going on.

I

SAQ 3 - --

An aircraft returns to the stand and the crew report that the toilets are not operating. List your actions to rectify the fault.

The toilet unit has three valves, one anti-syphon, one to control the water suppl, and one to control the vacuum flush (this has a handle to allow manual closure should it fail in the open position.

A s the tanks fill, the liquid eventually covers the level sensors. These send signals to the logic units of each toilet unit that feed into the specific tank. This signal will prevent the toilets flushing and the crew would get a warning. The flight crew would close the toilet off.

Level sensors tend to become soiled and this can lead to false readings, to help prevent this the rinse jets are directed towards them. This means every time the tanks are serviced, the sensors get washed.

/

- -- . -

\ ' /' Fig. 26 ELECTRIC& CIRCUIT FOR THE VACUUM FLUSH SYSTEM I I ,'

, (BOEING) 1 1 1 ,

I I

i /A'

I ~ Take a -- moment - - to study figure !26.-~ote&e'flush switch in the lavatory control unit. Note that its power supplycomes Born the ground serfice p'anel or the flush control system flight via an AND logic gate if the sensors are serviceable and the waste tanks is not full. Note the three-phase supply (one line shown) to the vacuum blower.

Maintenance

The vacuum flush system, unlike the re-usable liquid flush system, has little disinfectant fluid added. During servicing only a small pre-charge is used and this means that there is a slightly higher health risk when servicing this type of system because of the higher concentration of effluent to disinfectant.

The precautions that follow should be observed when-ever servicing any toilet system irrespective of what type it is.

Precautions

* Ensure that your injections are up to date. The policy regarding injections i s up to the company and the risk assessment that has been carried out. The injections could include anti tetanus, hepatitis A and B, polio, typhoid etc. The risk assessment may have concluded that these are not necessary provided all proper precautions are carried out, but would probably have to be given after bodily contact with contaminated fluids.

* Wear bio protective clothing to include disposable gloves, overalls, face-wear, etc.

* Use special dedicated "toilet" tools (not your own tool-box tools).

* Plastic bag any removed items (articles of clothing, needles etc that may be found in the toilet effluent and clearly label as 'Bio Hazard'. Inform the authoritiesif -- any - banned substances (drugs etc) are found.

* Return all tools to stores in labeled plastic bags. I

* Remove protective clothing and dispose of as biological waste. -

* If splashed in the eyes immediately irrigate copiobslfwith eyewash or clean water. Seek medical attehtion. I 1

x If cut or grazed when working on these systems immediately wash with soap and water and seek medical attention.

* Wash thoroughly after working on these systems.

Note. For most operators the ramp servicing of the toilets, from cleaning the toilet cubicles to draining/recharging the systems, is sub-contracted to service companies. This means that specialist teams (cleaners for the toilet cubicles and separate toilet draining teams), will arrive at your aircraft after landing suitably prepared with all the correct equipment etc to carry out the task/s.

However, it is your aircraft and it will fly on your CRS so you must know what is required.

The most common tasks that are performed on vacuum flush systems would include removing and cleaning the single point sensors. Also flushing through the pipes with acetic acid and crushed ice to clear blockages is another fairly common task. (It is general considered that the pipes are not big enough as blockages are frequent.)

TOILET CUBICLES

Daily servicing is usually carried out by the cleaners, but like the toilet systems themselves they are ultimately your responsibility.

The cubicles may come in various forms, but for most large commercial aircraft the unit is fairly standard - this applies whether installed in the main cabin of the aircraft or business class or first class.

Regulations do not permit smoking in the toilet cubical. Also fires have occurred in the towel bin due to people putting cigarettes in the bin.

Therefore smoke detectors are fitted (normally in the ceiling), giving aural and/or visual warning to the cabin staff or flight crew (a regulatory requirement). The smoke detector is operated continuously on the ground from ground power and in the air, when the toilet and galley air extraction system is on.

r- - - -- I ', r -

\ -- EMERGENCY, OXYGEN , I

I DETECTOR

GRILL CABIN FLOOR

Fig. 27 TOILET CUBICAL - EXAMPLE

rnoodull l A-1475

In general the cubical will contain most of the following equipment - the maintenance of which you are expected to know:

* A sink - usually made of stainless steel with hot and cold taps, soap liquid dispenser, plug operating knob (between the taps - usually).

* Toilet bowl - stainless steel (Teflon coated for the vacuum type). On the Elson type and vacuum type the bowl may not be paneled in, though it is usually.

* Lights. A low power one that is on all the time aircraft power is on, and a brighter one that is microswitch operated when the door lock mechanism is engaged.

* Mirror.

x Towel dispensers - for drying hands etc. May contain paper towels - - -

linen towels. - , I

* Toilet roll/ tissue dispenser. ', I

I

,

* Used towel bin - stalinless steel &th a spring loac!iet$ lid. Because of the potential fire hazard mntai'ns a self operating fire bottle, - check the book in this series onFire Detection Systems. ,

I I

* Ash tray -just in case someone walks in the toilet with a cigarette on. Remember, smoking is not allowed in the toilets br &hen walking in the aisles - though on m y fligfits smoking is b,mned altogether.

, I - - - - L -

* A lockable door which can be unlocked from the outside by the crew and when locked puts the cubical lights on and illuminates an external occupancy indicator.

* Toiletries draw or draws. These may contain:

Soap bars, shaving creams (all sealed) etc. Tooth brushes (sealed). Safety razors (sealed). Perfumes/aftershave etc. Ladies sanitary wear.

* Baby changing table (folding).

* Flush handle (except Elson type).

* Attendant call button.

* Fresh air inlet (usually a small grill at the bottom of the door from the cabin). Exhaust vents - connected to the lav' and galley exhaust vent system.

* Various notices and placards (No Smoking, Place Used Towels Here etc).

* Speaker - for inflight entertainment (music) and to relay crew announcements.

* Oxygen supply.

-k Single phase ac power socket.

* Microswitches.

- -

I r C_ORKOSION, -- \ CONTROL \ --- I i 1

All toilets (and galleys for that matter) haGe keta17s worst e d d y - wdter, with the added problem of the chemical additives. Co~rosion is always a problem i n these areas which is even more so in bld or geriatric aircraft. 1

1 --- /

For a kofrosion cell to form therk must b:mpisture present ad> -- potential -

differenoe (pd). The structure uhdbr a toile; is the perfect place for corrosion to start, d d due toits lack of insqection accesd it can develop buickly. There is a need tp provide/ektra protection for these &{as.

I ' I , c / / - - /

l 1 ' / / - -

- - - - Under Floor Structure

The aircraft floor structure (including under galleys and toilets) may be made from aluminium alloy 'clad' in some cases with a thin film of aluminium (al-clad), composites or sometimes titanium might be used.

SAQ 4

What considerations would you take into account when handling equipment/cargo/working around this structure - or any structure for that matter?

Irrespective of the actual structural corrosion control measures taken, structure under the toilet floors will probably start to corrode anyway. This is most noticeable under the re-usable liquid flush type where there are strong chemicals.

rnoodull l A-1477

STA 2f33.5

FLOOR BEAMS

- . Fig. 28 TOILET UNDER 'FLOOR STRUCTURE - EXAMPLE BOEING

1 \

I 1

i I Even if composites are used the seat tracks are still made of aluminium alloy for strength and wear considerations.

Some floors under the galleys and toilets are titanium because it is almost non- corrodable. It will corrode but a t such a slow rate that for all practical applications it is deemed to be totally corrosion resistant.

ANTI CORROSIVE MEASURES

Besides the normal anti corrosive measures applied to metals (eg anodizing for aluminium alloys and cadmium plating for ferrous metals) additional measures for under toilet floor areas will include:

* Corrosion Preventative Fluids (CPFs) applied to the structure. * Floor panel clips taped to prevent scratching. * Floor panels cushioned to prevent scratching and damage. * Floor panels sealed to prevent fluid getting into the structure. * Moisture barriers applied over the entire 'wet' area. * Leaks in the toiIet system traced and cured at an early stage.

moodull lA-1478

Corrosion Preventative Fluid (CPF)

CPF is a water displacing fluid that adheres to the surface of clean (or primed) aluminium and it's alloys. It is a slightly waxy substance that can either be sprayed by aerosol can or by a 5 gallon (221) dispenser onto the structure. Care needs to be taken during this operation, a s the fluid spray will damage the lungs if breathed in. Use only in well ventilated areas and wear face masks. A thickened version of the fluid is available for brushing.

Consult the AMM prior to application. The biggest draw-back with CPF is that it acts as a debris collector. It will collect dust and debris as it is very sticky. During routine maintenance and certainly prior to any structural inspection this needs to be removed. Always record any work done (re-application etc) and clear all work with a CRS signature.

- -- I I

PANEL AlTACHMENT --

A 1 I

I

I ~ I

I

I I i I i /

I COSMOLEY 1060 /

2" VINYL TAPE I

I FLOOR PANEL /

/

I 1 I

I

/ L 1

i I

L p p

i INTERCOSTAL ,, '

SUPPORT -

COSMOLENE 1060

SEAL BONDED TO PANEL SEAT TRACK

Fig. 29 FLOOR PANEL FIXING

Floor Protection

Figure 29 shows the protective measures that might well be taken during the fitting of a floor panel.

Note the tape which is usually some form of Vinyl, this protects the area around the clip and prevents chafing and subsequent scratching. Foam is applied over the top of the clips for two reasons; firstly it helps to prevent the floor panel from chaffing the structure and it also allows some 'give' in the floor panel which is mostly a comfort thing (for walking on).

Different manufacturers use different approaches so it is important that you check the appropriate section in the AMM before installation.

Floor Sealing

Floor panels are sealed in a variety of ways using various materials depending on where they are located and how often it is anticipated they will require to be removed. For example, the floor panel at the entrance to a toilet may require lifting fairly frequently to track down leaks from under the toilet tanks. Other floor panels elsewhere in the cabin may not be lifted for many years.

- - -

7- -

Two types of sealant used are: , I I

I

* Wax based - at room temperature it is solid, when heated it becomes liquid. It is poured into the gap around the floor panels and left to cool. It forms a watertight seal, but is susceptible to/being removed by the coq-osive chemicals from the toilet. It is easy to'remove and does not damage the floor panels or structure in anFay . This means that it is usedlon panels that require lifting moderately frequently. I

I

I 1 * Poly-sulphide based t ~ o - ~ a r t mix solutions. A chemical reaction between

the sealant and the activatorcauses the sealant to cure. They are strong when set and as such provide a resilient dam to all chemicals and water. They are difficult to remove and when removed the process often results in damage to the structure and the panel.

The problem of removal can be minimised if parachute cord is placed in th, bottom of the gap before applying the sealant. This allows the sealant to be cut on removal by simply pulling the cord with pliers. Needless to say this sort of sealant is used on all permanently fured panels. One trade name for this compound is Flexme@ that dries to an almost cement like finish, it is very difficult to remove and care needs to be taken when doing so. Check your AMM for available compounds.

A n example method of sealing floor panels is shown in figure 30. The actual method varies from area to area and between aircraft.

rnoodull I A-1480

Other Corrosion Preventative Measures

After all of the above has been carried out, the area is ready for the toilet or galley complex to be installed. Prior to this a Mylar type moisture barrier is applied over the entire 'wet' area to extend it to 3ft (lm) beyond the area in all directions. This is totally water-resistant and prevents any spillage seeping through but may mean that the liquid moves to the edge of the water proof area. So the trouble may be moved elsewhere.

2'0.5 INCHES

0.1+0.2-nil INCHES

1 - -- J

I

I

I

I I 1 ~ ' I /'

1 '. i

I / ' I

1 ' I STARTITHE CORD AT THE AFT _, ' , OUTBOARD CORNER OF THE I

I ---

- - i ' PANEL. INSTALL IT AS TIGHT TO ' -- ---THE PANEL AS POSSIBLE. WIND -

THE CHORD CLOCKWISE ONE FULL TURN PLUS THE OVERLAP AS SHOWN

2 MIL TEFLON TAPE 2.0 INCH WlDE

\ PERMACEL P-306 TAPE 4.0 INCHES WlDE

1 1 \ ' RUBATEX TAPE

FLOOR BEAM /I.J Fig. 30 TYPICAL FLOOR PANEL SEALING TECHNIQUE

Galleys require additional precautions because they have drains and sills around the outside periphery. The dams collect the water and other fluids and direct them towards the drains, which are ported overboard instead of allowing fluid to collect on the floor structure.

The most effective means of corrosion control is to prevent the liquid getting to the floor in the first place. The vacuum flush system is a great improvement over the re-usable water system, and the under-floor structure on these aircraft is in much better condition.

On all toilet systems it is important that leaks are detected and rectified early. Often the first sign that there is trouble under the toilets is water coming from the lower bilge drains on the underside of the fuselage. Any liquid from these needs immediate investigation. Remember, liquid coming from the bilge drain could be (a) clean water (b) liquid from some cargo on board, or (c) toilet water. The clean water could be rain water, condensation water or potable water. In all cases the source will have to be found and the fault rectified.

GALLEYS

Galleys are used for the storage,: heating arid the preparation :of food. Drinks are also prepared and kitchen rubbish stored. Galleys may be fitded d the back of the aircraft or, on larger aircraft at centre locations. Figures 3 1 ahd 32 show typical galleys including service trolleys. (Trolleys are considered as p a t of the aircraft for certification purposes and are your responsibility - as are the ghleys and all electrical appliances there-in) .

Electrical appliances and t r o l ~ ~ ~ s are considered as controlled items under the ANO. Galley equipment can include:

Storage locations for trolleys and carts. A means of chilling cold cupboards - usually provided by a vapour cycle chiller unit located near to the galley. A means of heating the food - ovens (conventional or microwave). Beverage makers. Ice makers. Water supply - hot and cold. Water heaters. Trash compactors (or simply rubbish bins). Lighting. Interphone system. Fume extraction system. Water waste system (drains to the outside). Toasters etc. Screen curtains.

- 38 -

moodull l A-1 482

ACCESS TO SWITCH PANEL BACK CONTROL PANEL

WATER TANK CONTAINERS

GALLEY WASTE

CONTAINERS

WASTE BAG ACCESS PANEL

- -

RETAINING BAR

Fig. 32 GALLEY 2 - A320 REAR GALLEY

Note that figure 32 shows a full width galley located a t the back of the aircraft. The galley is made up of two half width modules that may be located separately on some aircraft. Note the positive locking of the service trolleys - by regulation.

Galleys are located onto and bolted down to the seat tracks, in a manner not dissimilar to seats. Figure 33 shows an example of the attachment. In addition, tie rods are attached between the top of the galley and the aircraft fuselage structure.

Galleys use a considerable amount of electric power, and have their own feeder circuits that will be load-shed automatically in times of high current demand. Galley power may come direct from the secondary aircraft busses or be provided through Electric Load Control Units, ELCU7s. These units measure the current being drawn by the galleys and prevent overload. They also act as a switch to remove the power to the whole galley complex should aircraft power consumption become high.

Galleys usually have an electric distribution panel accessible from the front of tl galley that allows for individual isolation -- - of;ovens etc by C / B actiK. This panel often has a light to indicate that power to the galley is on.

1

\

I

Fig. 33 GALLEY SEAT TRACK FITTING

All systems in a particular galley have control switches and these are normally located close to the galley. Often there will be a master isolation switch for both electric supply and water.

Galley Chiller Unit

The lower cupboards in the galleys are used for storing food prior to it being heated. They are usually cooled by a chiller unit. Figure 34 shows the location of a typical under galley chiller unit and its associated ducting. Note that the ducting circulates the air from and to the cold cupboards (chiller cabinets) in the galley. Note also that the drawing shows one VCCP removed.

The chiller is basically a small Vapour Cycle Cooling Pack (VCCP). Figure 35 shows the layout of a typical VCCP. Power supplies include a 28v dc supply for control and lighting and a 115v ac 3 0 400Hz supply for the compressor and fan motors.

The fan motor and compressor motor are both switched on from the galley control panel. The fan motor drives two fans, one draws air from the cold cupboards, and after cooling, passes it back to the cold cupboards in the galley. The other draws in ambient air to act as a heat sink for the condenser (cools the condenser), it is

I - - thenpdiscljarged overboard. -

I I I /

l 1 ~ I '

1 I

L

UPPER CHILLER

FILTER 8 CONDENSER

MOISTURE DRAIN LINES

Fig. 34 CHILLER UNIT TYPICAL LOCATION

- 41 -

rnoodull l A-1485

Chiller air is drawn across the evaporator and cooled by heat transfer to the refrigerant (Freon or similar) and passed back through the ducting to the cold cupboards.

The VCCP relies on the principle that as the refrigerant changes state from a liquid to a gas it absorbs heat (from the chiller air). The refrigerant is pumped around the system (as shown) by an electrically driven compressor and enters the condenser as a super heated gas. In this condition it is ready to give u p its heat to the cooler ambient air. The liquid Freon leaves the condenser a s a liquid and is passed to the evaporator via the expansion valve (sometimes called a temperature transfer valve). This valve is controlled automatically by sensing bulbs down stream of the evaporator.

At the expansion valve the refrigerant is allowed to expand and is made ready to absorb heat from any warmer medium. At the evaporator it absorbs heat from the air and becomes heated and turns into a gas. The Freon gas now passes through the temperature control sensor and back to the compressor, this compresses thr gas increasing its heat still further and turning it into a superheated gas. From

- -

here the refrigerant passes to the condenser and the cycle continues.

The compressor serves two functions: To pump the refrigerant around the system and to compress the refrigerant raising its temperature and incieasing the efficiency of the condenser. 1 -

1 . -

\

The expansion valve is a type of variable ristrictor valve contrplled by the temperature sensing bulb in the temperature sensing valve. The line between the bulb and the valve is a sealed line. If the refrigerant comes out hf the evaporator too hot (there is not enough refrigerant flow) then pressure in the sensing line will cause the expansion valve to open more and allow more refrigewt through. It gets its name because it allows the refrigerant to expansion as it passes through.

A programmable timer operates the VCCP in 10 minute cycles. If at the end of the cycle the temperature of the refrigerant is less than 26°F (-3°C) then automatic defrost is initiated - this involves the fan being shut down and the hot gas by-pa, valve being opened. Defrost stops when the temperature reaches 4°C.

Automatic shut-down is initiated if any operating parameters are exceeded eg:

* Refrigerant temperature higher than 54°C. * Refrigerant pressure higher than 300psi. * Refrigerant pressure lower than 5psi. * Compressor over temps.

Some chillers may be shut down when the cargo fire extinguisher system is armed.

i E - - D \CHILLED RETURN AIR FROM CHILLER

CABINETS $ $ , , ~ ~ ~ CABINETS

Fig. 35 A VAPOUR CYCLE COOLING PACK

To help reduce the noise levels from chillers they are usually mounted on anti- vibration mounts.

Maintenance of the chiller unit is confined to checking for security, leaks, damage and correct contents level - also functional checks.

- 43 -

rnoodull l A-1487

If recharging is required it is important that no air or liquid other than refrigerant is introduced into the system. Often units are 'sealed for life'. The ducting requires care during maintenance to ensure it doesn't become contaminated.

Other Galley Equipment

Galley ovens may be single or double ovens and include a fan (usually), a control panel arid a door fitted with a seal. The seal can get damaged and care needs to be taken during inspection to ensure that it is serviceable. Ovens are periodically tested using a thermometer to check that the selected temperature is achieved.

Care needs to be exercised during this function to ensure the electrical supply system is not overloaded - also beware of the burn risk on hot surfaces.

The oven control panel is usually a separate unit with an ON/OFF switch, lamp indication and a three position selector - fan only - high temp setting (350°F) an low temperature setting (300°F), and - - atimer - up to 60 minutes.-

I

An oven is a modular unit and maintenanLe includes checking for security, cleanliness, correct operation of oven and door closing mech+ism. If it fails to operate then the fault finding procedure common to any inoperative electrical component should be followed. -

- -

/

Some geleys have a Trash Compactor. This is a large press tdat squashes the trash to reduce its volume. They need regular:cleaning. If this i$ not fitted the galley will have a simple bin arrangement that has a hinged top access panel.

Beverage makers are installed to brew hot drinks. These are electrically powered, with a supply of potable water. There are lights on the front that indicate low water and system ON and in use. These units are slid in from the front and have rear connectors. This makes changing easy but can lead to water leakage from the aft water push-in connection, particularly when the water pressure is low.

On twin deck aircraft a galley lift may be fitted to allow the carts (service trolleys) to be loaded on the lower deck and transported to the upper deck galley. Currently the B747 is the only twin deck aircraft, but the system could be fitted to aircraft coming into service such the Airbus A380.

On the B747 the lift is driven by two motors that drive a vertical screw jack. The lift is attached to this by way of a ball nut assembly that allows the lift to be lowered and raised between floors. Micro-switches prevent the lift from moving unless the doors are closed. Mechanical locks will prevent the access door to the lift shaft from being opened unless the lift is in front of it.

SAGINAW BALL SCREW ASSEMBLY ROLLER

CARRIAGE ASSEMBLY

I ' 1 1

Fig. 3d I ~YPICP;L~/~ALLEY _ - LIFT 1 1'

I / I --

- \ \

The micioswitch interlocks can be overridded for rnaintenanke purposes. preca*tions wfuld you take befbre-entering the lift shaft to carry out an

I-' - L , / inspec-tion?---' --

What

That concludes this book on domestic systems. Have a go at the following exercises. The answers are to be found in the text.

EXERCISES

1. What can drinking water tanks be made from and where are they located?

2. What is significant about the fixtures of a drinking water tank? 3. What does the light on the sink water heater go out for? 4. How is drinking water cooled? 5. How are the contents of the potable water tank indicated? 6 . Where is the contents information displayed?

When are drain masts heated? What is a gray line water valve - what does it try to overcome? Try spending some time looking at the water system of your aircraft (if it has one, and if you have access). Locate the tanks and note the fixtures. Locate the rest of the components and check on their positions in the system and attachments to the structure. Spend a little time looking through ATA chapter 38. Describe how to ramp service the re-usable liquid flush type toilet. How is the flush cycle controlled during the flush cycle of a clean water type toilet? ~esc r ibd a basic toilet vacuum system. What are the main advantages of this system? What is the anti-syphon valve for? What valves are located on the toilet bowl assembly? Spend some time loohng at toilets and toilet systems (if you can). Use the AMM to check the location of the various components and check on the maintenance that is carried out. What material is normally used for floors under tGilets? How are under toilet floor-phels fitted? I How does CPF work and where is it applied? I

Read the AMM for your aircraft and look into the method used to install and seal floor panels. What equipment might be fitted to a galley?

I -

Explain the principle of a VCCP. Take a close look atlthe galley, it's equipment and the chiller unit (if you have access). Check past entries in the Tech Log looking for the history of reported defects to the galley and its co~mponents/systems.

ANSWERS TO SAQs

SAQ 1. A detailed inspection requires good access, with good lighting and the possible use of other equipment such as a magnifying glass etc. You would almo-+ certainly require the tank to be removed. (Read AN 3 if you are unsure of your responsibilities). Removal and fitment of the tank should include the following:

Refer to the AMM. Isolate the compressor, pressurisation mechanism and water heaters. Release any pressure and drain the system and check it is empty. Remove all the connections and carefully remove the tank. Ensure all parts are labeled and attaching parts are inspected and salvaged if appropriate. Fit blanks to the tank and place on a clean protective covering. Remove any insulation/equipment to gain access. Inspection of the structure, rectifying any defects found. Ensure all insulation/equipment is refitted and inspect the tank mounting lugs for security, damage or corrosion.

Check that the correct part number bolts etc are used. Refer to the IPC. Inspect the tank. Install the tank with care, considering that it is easily damaged and it is also very bulky. Remove blanks and inspect the interior as far as possible for damage and contamination. If serviceable fit the tank and make all the connections. Leak check the tank and the connections using an appropriate quantity of fresh water. Check the contents indication system. Sterilise the system. Test any systems disturbed. Fill tank ready for departure. Complete the appropriate paperwork.

It is important that you study the style, content, and the logical pattern of the answer. The CAA would e x p e c t j ~ u tqgive -. this sort of detail in -- any - essay answer.

C _ r d _ \ - 7 - -

SAQ 21. Isolate the motor by tripping the appropriate C/B (arid tagging). Check the motor and the wiring for any sidn? of overheating - change wiring and/or motor if found. Ifi the motor has overheated then tlje thermal protectionl device that is desided to prevent this (a mandatory fitrheht - ANs) has failkd. If it is not part of the mbtor then it must be changed also add a report made td theaircraft

r -- l 1 \ I manufacturer. -- ,

', I I 'i ' I

Check t i e troub,le shooting section of the AMM. The most logical fault is likely to be the timer logic' unit which shbdld be changed and the systeA tested. Record all the wdrk\don6"qd sign the CR-S. - - ' / / -

--- -- - I- -. _-, !

SAQ 3 . Check if the tanks are full. If they are the system will not allow the toilets to work. They should not be full as the aircraft has just departed and someone has not done their job right - but drain the tanks and check the system.

If the job looks like being a big one the passengers will have to de-plane.

If the tanks are empty check the fault finding section of the AMM. Check to see if any of the toilets are working, if non are it may well be a power supply problem. The rest of the aircraft systems should be working otherwise the crew would have reported them as a fault also, so the main bus bar supply should be OK.

If no electric power to the toilet system fault check the electrical supply system - visual first - check C/Bs or fuses - continuity check or voltage output check - functional etc. Significantly more detail would be required here ifit were an electrical question - and for this you should refer to module 3 and the books in this series on electrical systems.

- 47 -

moodull l A-1491

If only the toilets to one tank are not working you should check that power is available to that toilet system (similar to above). Check the level sensors (These may need cleaning - by flushing/servicing the toilet system. If the toilet works but there is no suction check the vacuum blower C/B's. If they have tripped check the reasons why, rectify the cause, reset C/Bs and retest. If a n electrical fault rectify as above. If C/Bs and elect system OK change blower and retest.

If any valves fail to work (flush valve, rinse valve etc) check the electrical supply and appropriate valve as above. If supplies and switching are OK change the valve.

Record all the work done and sign the CRS.

SAQ 4. All appropriate C/Bs should be pulled and 'gagged' to ensure that they are not reset whilst work is going on. Also, operate maintenance 'break-outs' to de- clutch the motors to prevent inadvertent movement. Take care, lifts move fast and can be dangerous.

- I r -- - - I I

1 I \

1 I 1 i I I

>>>>>>>>~>, l> l>>t , ;> I ! ' 1

- -- /' I - - , I -- -' I i I 1

I

/

- -

CONTENTS

Page

Basic principles BITE Central Maintenance Computer - introduction Fault finding Central Maintenance Computer Flight phase logic FDEs Fault correlation Fault history -

- - .

CMC - hdditional functions \

Air/ ground simulations Data loading I

Printing I -

Condition monitoring - - -

ACARSl link \

\ \

DFD R facility I , Maintenance pages

,

On-board manual systems , Answers tb-self assessment questions -

G l o s s m f t c m s - -

rnoodull l A-I493

BASIC PRINCIPLES

Clomputers are used extensively on all systems on modem aircraft. Almost all electro-mechanical and electronic systems will have some form of microprocessor control/monitoring unit and this will be connected to the Central Maintenance Computer (CMC) - figure 1.

We are not going to cover the function of these computers here as this is dealt wlth in the books on Digital Technology in the LBP series module 5 EASA66. However, we are going to look at the operation of the system as a whole including how the various systems interface with the CMC and how the ground crew and flight crew use it to indicate the fault status of the aircraft.

I t might be a good idea, if you have already studied module 5, to have another

( look a t it - just to refresh your memory as to how a computer works. If you have not studied module 5 yet you should-be-able to complete this book anyway, but would'need to read certain sections more thqn once. - -

I

I \ I 1 I To help in your studies I have inclAded a glos,sary of terms at the back of the book. This should be used as reference. i I ! I 1

' I , 1

The system is designed so that all a i r q d t systems "report!' their-status to the CMC yi th provision made to 4ey any faults. There is also an-extensive BITE system: \ 1 ' I : ! ~

I

I

Built-In Test Equipment (BITE)- - - _,

-- - - - - -

Computers used on aircraft are designed to perform a certain function or functions. For example, in a fly-by-wire system, an input signal by the pilot via the control column is transduced into an electronic signal and sent to the fly-by- wire computer which will "work" on the signal and as a result will sent an analogue signal to a n electric valve of the PCU (Powered Flying Control Unit) to move the controls.

It will do this after taking inputs from other computers - air data, engine data, altitude, the PCU itself etc - and processing ("working on") that data.

From this there is a basis for system monitoring. If the inputs are monitored and measured and the outputs are checked that they are within limits then we have, essentially, a BITE system.

- 1 -

moodull lA-1494

BITE can come in one of three forms:

* Star tup. * Interruptive. * Continuous (called Cyclic on some systems).

Start u p BITE is limited in its ability to detect failures resident in the system. At power-up BITE will check the power supplies are good, looking for correct phase rotation, frequency and voltage.

It will carry out checks of the processors and the areas of RAM in the computer, ensuring all are clear and functioning properly. It will then, through its Built In Operating System (BIOS) invoke the loading of the main operating software and any airline modifiable software. This process is very similar to the 'boot up' sequence of a personal computer.

Intequptive BITE, is something the engineer (or the pilot in some cases) can initiate. In its simplest form it is a press-to-test, but with on-board maintenance systems it is usually much more. Generally what happens is the inputs and outputs bf the computer are electronically (ie, not physically) ,disconnected. The BITE now starts to systematically inject a signal into each channel or lane and look for fhe corresponding outpGt.

I I

I t will injLct signals across the entire design range of the system, with generally, these e$reme functions never being met with in normal operations. Sometimes this foiTn of BITE is called Ground Tests and we shall look later on how this is initiated!

--

SAQ 1

Considering what happens during interuptive BITE what precautions might you and the designer take before initiating such tests?

Continuous BITE, as the name implies, is carried out all the time. Inside the Line Replacement Unit (LRU), which in this case is the computer, there are (at least) two channels that work out the information and provide outputs. These two channels are called Command and Monitor. (If there are three channels the third is called Standby).

All channels receive the input signals and calculate the output solution. The command channel will output the signal and send it onto the unit being operated (whatever that is), the monitor channel will use its output figure to cross-check that the command channel is functioning correctly.

Brief History of BITE (for information only - not in the CAA syllabus)

As soon as computing became available on aircraft with very basic niicroprocessors, it was possible to monitor the outputs against given inputs. Hence BITE was formed. Initially BITE was a box bolted onto the side of the LRU that was used in much the same way as ground test equipment is used.

The BITE was a microprocessor that provided input and output monitoring. Access to the LRU was required to initiation the BITE (by pressing the test button). Typically LED's would come on to indicate system faults. These LED status indication data would be cross referred to the maintenance manual that would give suggestions as to the fault.

I n time, the BITE box was incorporated inside the LRU it was testing but initiation was carried out in the same way. Continuous faults would be shown on the front

i of the LRU by looking for illuminated LED's. - -- - - - - .

A further development came when the-LED'S were moved down to a centralised location, bften close to the re-fudlidg panel. A bank of LED's &as available for checking with the aircraft on the ground. As the BITE picked ufi failures it sent a signal 20 illuminate the LED. 1 , I 1 , I

I /

- -

As you may know computers car) startto wbrk for no apparer?t reason, in much the samelway as they fail - without warning.'so in this case the LED &uld go out

I I and the kaintenance engineer would have no ,indication that anything had ever gone wrong. ~hi~ ' ; laced the enginLer in the situation where hjelshe would have to change1 items 'on spec' without a sure indication that the item was actually faulty. Not a-very-acceptable approach to good-maintenance practice:

- - - - -- / , , -

What the designers did to help to overcome this problem, was to allow the LEDs to stay illuminated, until they were reset manually. What would happen now is; the aircraft lands and the engineer checks the maintenance panel noting all illuminated LED's. The reset button is now pressed and the LEDs that remain illuminated are noted. Appropriated rectification action would be taken in accordance with the AMM, the work signed for and all remaining illuminated LEDs cleared.

These types of failure gave rise to new terminology, ie:

* Soft failure - where there has been a fault that has subsequently cleared.

* Hard Failure - where the fault is still apparent (ie LED still illuminated after reset is pressed).

This gave the engineer a better indication of faults that had perhaps cleared but would still provide a good chance of accurate fault diagnosis.

- 3 -

rnoodull I A-1496

SAQ 2

Why might a fault not be apparent to the engineers, but the flight crew constantly complain about the defect? How do you think BITE might help u s detect this?

The final steps in BITE technology (so far) came when the BITE box was moved to a central location. In fact, the box actually becarne a computer in its own right, and Boeing call this computer The Central Maintenance Computer (CMC). It forms the basis for all on-board maintenance systems and its introduction has given rise to a new ATA chapter being formed - ATA 45.

Airbus (and other manufacturers) have the same philosophy but call the unit the Centralised Fault Display System (CFDS).

Access to tests and checks are now via a centralised menu and can even be controlled by a 'windows' style software package that is very similar to any other com$ter. We shall spend some,time discussing this type of system.

I I

Central Maintenance Computing - Introduction

Figure 1 gives details of a basic system. Note the two way flow of information from all the lother computers to the CMC with this data being converted into graphical symbols for transmission to the CRTs (or flat panel displays) on the flight deck. (CRT = Cathode Ray Tube.)

- .. -

Aircraft system

00~00 Display system n

Fig. 1 CMC SYSTEM - BASIC LAYOUT

- 4 -

rnoodull l A-I497

The Graphics Generator (may be called a Symbol Generator) has stored infonnation on the symbols to be used on the flight deck displays. It takes the electronic signals from the CMC and generates the appropriate syrnbol for the display system.

Data is sent from all the aircraft systems to include the:

APU. autoflight system. Doors. ECS. Electrical system. Engines. Fire protection systems Flight controls. Fuel systems. IceIrain protection-systems._

* , - Indicatinglrecording systems. " -

* Landing gear. \ \ 1

* I Navigation systems. 1 1 I I * I

I Pneumatics. ) I , ; I

/ , I ' The aircrkft system computers afe ico~tinubvsly working and cah-ying out calculatidns as defined by their lodded soi'twhre. Each cornputer-knowd what the inputs and outputs should be and these are continuously monitored. If any paramiter goes outside of limits k i message is bent to the CMg, through a serial data highlway. The CMC will now start to do many things ie:

\- / , i

* It checks-the reported fault against-known criteria. It interrhgates the faulty LRU.

* It checks the fault is in a 'reportable period' (more of that later). * It checks another failed unit is not causing the reported fault. For example,

if the IRS failed this would show hundreds of faults in all the systems that use attitude data from the IRS.

* Providing all this defect correlation is correct, it will indicate the failure to the flight deck in a number of ways.

* It will store the infonnation in data banks (called fault history) for later (long term) analysis.

Some new terminology as used by Boeing (B) and Airbus (A):

Status Message. (A & B). Is a message that is placed on a status display screen. A status message gives the engineers a true indication of system failure. Aircrews do not have immediate access to the status page, but would look at it when 'cued' to do so. For an aircraft to be dispatched (ready for flight) there should be no status messages (or a pretty good reason for having one).

c - 3 -

rnoodull lA-1498

Maintenance Message. (B). A maintenance message is set by the CMC after it establishes that a fault does exist. It takes the form XX-YYYYY. XX' indicates the ATA chapter that it relates to, eg 27- flying controls. This message is never displayed to the aircrew as it is only ground engineers that use the information - we shall discuss how, shortly.

Fault Codes. (A & B). A fault code is very similar to the maintenance message. It is set by the failed LRU and the CMC will go through the process of correlating it all together. As with maintenance messages this information is not available to aircrew, as it is only engineers that need it for fault diagnosis reasons.

Maintenance Memo. (A). Airbus use the term Fault Message. A maintenance memo would be set if the failure is very insignificant, and there was sufficient back-up to mean the failure could almost be ignored. Here, the aircraft can fly on quite safely until its next scheduled maintenance when all these minor defects can be rectified. This keeps maintenance costs down yet does not compromise safety;-

\

Memo Message. (A & B). You might hear reference to this, although it is not strictly related to the CMC. A memo message is displayed for the aircrew to see and it gives them reminders of systems that are in operation eg APU running, Park Brake Set etc.

For fault finding reasons, they afe of no use to maintenance engineers, but you may hear them referred to. Obviously the engineer would use them at the appropriate time eg, when setting the parking brake.

~ a u f t I?inding - The Digital Way-

Access to the CMC is provided in a number of ways, the most common currently is through the flight deck Central Display Unit (CDU). The Line Select Keys (LSKs) are used to input requests for information. In this system (which is common to airbus and early Boeing) the requested information is displayed on the multi- function displays that normally display flight information.

(The most modern system uses a separate screen in the flight deck and requests for information are by way of a Windows style menus. Display then being given on the same screen. This type of system is fitted to the B777 and uses fibre optic interfaces .)

The CMC will display the fault code and also the maintenance message, and these will be correlated to a Flight Deck Effect (FDE). So the first indication that there i s a problem will probably come from the crew - either verbally or recorded in the sector record page of the Tech Log. That will give you the engineer a FDE to work from. You can now check the CMC present leg faults pages and find the appropriate maintenance message.

--- , Fig. 2 FIM CHA*ER SUEUECTS ' 1 - - _ _ --'

-1

Tlne engineer takes these numbers to the Fault Isolation Manual (FIM) (an essential tool in today's integrated aircraft and supporting systems design). Sometimes a fault in one system will show itself as a failure in another system and this type of 'fault hiding' can render traditional fault diagnosis techniques useless.

(The FIM is not very well excepted by the older generation of engineers and some see it as taking the skill out ofthe job. It is important that all engineers use ALL the tools that are available and the FIM is just that, a tool.)

The FIM is split into ATA chapters. Figure 2 shows the breakdown of a typical chapter, in this case ATA2 1, which is air conditioning.

Notice the index titles (two of these should be familiar to you).

21- Fault Codes Index, is the numerical index of all the possible fault codes. Fault codes, as you might recall, are set by the failing LRU and correlated by the CMC.

rnoodull l A-1500

2 1-Maint Msg Index, is the numerical index of the maintenance messages that the CMC sent. The CMC sent these in response to the failing LRU's 'cry for help'.

From the interrogation of the CMC pages on the aircraft you will certainly have one, and possibly both of these numbers. Figure 3 shows what these index are like once opened. Mere we are assuming that we have the fault code and one of the maintenance messages that are displayed. Notice the FIM tasks in the right hand column.

With reference to figure 2 you will see the other two sections of ATA2 1 were FIM tasks and task support. The FIM task is the actions you would take given a specific maintenance message. These tasks lead you through a 'most likely' fault finding approach; figure 4 shows you an example of a FIM task.

I

1 I

Fig. 3 FAULT CODE INDEX

213 147 42

blank

GO TO FIM TASK

- 8 -

rnoodull l A-1501

I I I 1 I MAINT MESSAGE FAULT CODE

I

PRESS RELIEF 'VLV R (EICAS A D V I S O R Y )

I

I

FAULT JDESCRIPTION ,

21 -01 391 21-01393 21 -01402 21-01 451

21-31 TASK 803 21-31 TASK 804 21-31 TASK 809 21-31 TASK 899

TASK NUMBER AND DESCRIPTION

A. f b l n t.~k 1 s lor ..lnl.n.nr. ..m..p.: l ? - I Z W .

B. ,"l,t., cr.,u.r,rn < I ) I 1 Ih. M r .~DII I ( I I I5 r l l h Ih. .#lnlen$n(. ..s%ap.. I h n , O Ih.

MAINTENANCE MESSAGE V ~ u l t I . o l . l l a Vcorc&. b r l w .

< I > 8 , a h . -1 .h-. L I T ' N l * r # % b Lh. m.1nl.n.n~. ..s.mp., Iln & tL. Iz-Pr<.,ry Illp,., <a 4 l . h . I ~ , u n * o . \ l l a Ir.".e., ,.,I a ,L.

(31 I 1 Ih. MI ~lcr . P l I l r P , t h r m terr.rllu. a c l l a 1% ~.....ry. < I > I f th. N T 'hw. r A t l l l , do Ik l u l s I . o l . l l m Pro..&. b<p.

INITIAL EVALUATION' PARAGRAPH

lull ,laL..la ,roc.-. L1) n.k. svr. I h m t Ih. ( a t r o t cat- 11 a t Ih. -lr.L 0 o . l l l a . ( 2 1 b I l t a . . c l < m . r l o r *U711& l r o l the I c I I 1 c a l m w s l t l m

,r.o,bn... " I?I , I . 13) Bs a chack fw , V b < o h # 8 d * et c-ct- 8*7t36. ( I> I f %hare 1% I IU b.iu.n .Ins .nd 9 a# W 1 1 Y . do lllil charkt

(.> nalw. ch. r.slsr.n.as bet-- ,he*. o.1.m of ~ l n n 01 th. < a c t o r m th . trmrdi..,, "zltl': I ) PI" 1 I. 01" *, . ~ i t l l . l ... 1.l.n.. I.-- oh.. 21 ? I n I ~s .in 6 . 1 w l l 1 . d r.lltlulc. (<-m &a ,I PI" I ,..%" 9. ,P..Ll,,dr..l.l.r.c. cr-,, oh..

I b ) I , th. r..l.l.nc. 13 ml In Ib. ~1.0. 1mc0ll.d for .arb p l l r 01 D l n l , r . o l l ~ # ih. l e f t I cot- m.lilcn tr.nsb.r.r, 1111114 1.m 2,-,!.0'~40l~.

I t 1 I 1 th. re%l.lmnca 1 1 In the rmg. % o l r l O e d I- .arb o.lr o l olns, & t h i s 'ha&: I ) W n 1h.s. clrrull br..L.r. m lh. I e I I Plr. m < l O l :

.I I, .c< m 61 LJ .I, *r* I, 41 ILPs* PI" d) AT LL<Y sCU

Finally thkre is, sometimes task support data,, that provides wiring diagrams and locations of-components. This section is not,always available, however.

I -- - : - 1

To recap on the sequence of events:

- - - ~ - ~ ~- ~

W I R I N G CHECK , .-

i

!

Flight crew report FDE from EICAS/ECAM (EICAS - Boeing, ECAM - Airbus). The ground engineer (you) interrogates the CMC pages for present leg faults relating to that FDE. Cany out any quick testlvisual inspection first to unlatch/rectify the fault etc. (This has not been covered yet but it is realistic to assume that before you march the '/z mile or so to the office where the FIM is kept it would be a good idea to try any 'quick fixes7 first - a visual inspection is almost essential). The latest CMC's do actually provide a quick guide as to what might be wrong. Note the fault code and maintenance message(s). Take those codes to the FIM. Check against the fault code index for the codes you have and get the appropriate FIM task number. Work through the FIM task number to its conclusion. Restore the aircraft (carry out rectification iaw FIM/AMM) to normal and complete all relevant paperwork.

-- - -' ,/ ! I , 4 I I ,

, , ~ -- - \\ ; - '

i : 1 ' ,

8 , i

'\ : , ' !

, , I I j ;

1 I) 8Z srls ,111 p, B' I l l i O f s l B s l (L, .) L ) 86 E L ~ Y r c x l < 11 81 .UO 11111 L C I I I L I

ZI ..u~. th. 1.1, I ME, a u z l 1.0. t h CI-5 .bait . Jl 10 I c m l l w l t l chert b.8n.n ,la,. vlnl 01 c a r t o r

nU7$>< 8 , t k E l - 5 z h c l l n r d c-c<or b4'7'J a t th. t l - S .h*I, <"t" ? r - , , - t , , :

W 7 I S ' o'LT11

p'n, -..--.--.--.------rln ,2 .I" r 10

----

.I,, * --1; .I" I,

-- ~

,, I r ) T G - , l y . At- I" ~ h a driq, L th.,. ' _

--

, , ,

I , , i Paw 207 ,

i JY 15/95 ~ 1

Before we move off from the FIM - sometimes you don't get a fault code and maintenance message. I n these cases you would use the observed faults index or the ECAM/EICAS index that enables you to develop a fault code. Not a s quick and clinical bu t it does get you there eventually.

If you work in heavy maintenance - a word of caution. The FIM is written assuming that the aircraft is otherwise serviceable, basically it is written for line' operations. Now that doesn't mean it has no value in the hangar. It is still a very useful tool, but you need to be aware that it does have its limitations. If other parts/systems/components of the aircraft are removed/made inoperative for any reason then the CMC might not be working and/or FIM would not be appropriate.

That concludes this introduction to on-board maintenance. We have covered a lot of ground and to some this will all be new. It is up-and-coming technology and it is something we will all see more of, so consequently it is a system you need to be increasingly familiar with.

I ,

CENTRAL MAINTENANCE COMPUTER

Figure 1 kave details of how the fault information is passed from the user systems to the CMC and on to the output device. So far we have not explained much of the detailsi of this process.

SAQ 3 :

Consider for a moment when we will NOT want to report defect informationj7om the CMC for onwards display and subsequent storage in the fault history?

From SAQ 3 we should start to see the requirement for a couple of parameters to be set. We need to have 'phases' of the flight so that we can work out when to filter out maintenance induced faults. Secondly we need some way of identifying when faults happened previously. These are called Flight Phase Logic and Flight Leg Logic.

Flight Phase Logic

The CMC divides a flight into flight phases as follows:

(1) Power on / power up (2) Engine start (3) Taxi (4) Take off (5) Initial climb

- 10-

rnoodull l A-1 503

Climb Cruise Descent Approach Landing Taxi in Engine shut down Maintenance

The CMC is receiving infomation from every system on the aircraft so can work out which phase it is in.

At power-up the CMC resets and puts itself into the first phase. It now waits for the engine start signal from the Electronic Engine Control - EEC (the digital device that controls the engine). When it receives this signal the CMC moves into the next phase.

I - - - -- - - -

Now it loo,ks for a brake release signal.- \ C _

I I

' ! ' 1 Take off phase is set when take-dff power i; selected.

1 I

Initial cli&b phase looks for a weight-~ff~wheels signal (weight switch, ground/air sensor). I - -- I I -- -dl

1 , I ,

\ \ I - -

Climb dhase is set-by a combinatioh of FMC dnd time delay fiP& take-off. I i I

Cruise is given by a combination06 the FM,C $nd the IRS gividg stable attitude for % _ I a set period; - -

I -- -- -- . - 1- -- A

Descent is given by the FMC alone. Landing is given by the weight-on-wheels signal.

Taxiing is given after the thrust reversers are set from reverse to forward thmst arid idle power has been selected for a period of time.

Engine shut-down is given when the run/cut off switches are set to off.

Each phase change requires the previous phase to have been set before moving or1 .

The 'odd ball' phase is maintenance, which is a phase that ground engineers can in'duce before any specific functions are carried out. The CMC will consider what flight phase the aircraft is in before it signals a defect. This is called flight phase screening.

Basically, any defects that occur in the flight phase will be recorded and hence stored in the fault history. Faults occurring outside the flight phase are displayed as existing faults only.

rnoodull l A-1 504

This means that all the faults 'we' induce on the ground do not clog up the fault history. This keeps the fault history for real defects not maintenance induced defects (such as removing a component for replacement).

FDE's

Flight Deck Effects (FDE's) are any effect that is noticeable on the flight deck. Certainly the 'Status' messages that we have spoken of will be FDEs. But also missing flags or failure bars on primary flight displays are also FDEs.

Fault Correlation

Fault correlation is a process for linking together FDE's and maintenance messages. We as ground engineers need to be able to relate a flight crew reported defect (they will quote the FDE) and link that to a CMC generated message. If we can'kdo that the system is next to no help at all. Fault correlation does that and a little more besides, consider figure 5. ,

I

MESSAGES D I S P L A Y S REPORTS

Fig. 5 FAULT PROCESSING

- 12 -

rnoodull l A-1505

Input Processing. All the aircraft systems send information to the CMC and it - arrives a t the input processor. Input processing will interrogate the failing LRU alnd possibly initiate BITE checks. It will check the information that it is receiving i:s correct, valid and accurate.

Input processing works a t very high rates (typically processing u p to 5,000 faults per second and receiving about 5 times that number). Providing the faults are accurate, the messages are passed onto the next sub-routine.

Inhibits and Special case BITE. When the aircraft is powered for the first time or - starts an engine, we would expect many failures to happen. We don't really want to be processing all of these as we know they happen every time so this sub- routine filters these failures out. It also removes failures caused when the engtnes are shutdown or power is removed from the aircraft.

Cascaded Effects. If a key LRU were to fail, for example the IRS. The IRS provides - irlforrnation to virtually all systems-on-the-aircraft and if the IRS failed, it would induce Inany failures in the other user-systems. These failureg would be reported to the CMC but this sub-routine would detect the cause and ?e$son for the failure a n d filtkr them out. I I I I

I I ,' ' 1

Consolidation. Often several maintenance messages will be generated when - sorneth;ng fails. This sub-routink brings thkse related failured tbgether., It is not like cadcaded effects, which was about remoriing faults that were caudd by other failures. This is grouping together similar faults.

1 1 I I

' I

Correlation. A FDE needs to be linked to a failure message. ~Ae/correlation sub- - routine has details of all the fault codes, maintenance messageskand FDEYs and is able taplink-ttheiii together where appropriate. On some systems tdis sub-routine categorises the failures:

* Class 1 - Operational consequence and those that would require referral to the MEL.

* Class 2 - No immediate operational consequence, but ones that will require moderately speedy rectification.

* Class 3 - No consequence on aircraft safety and can be left to the next hangar scheduled visit.

NOTE: Airbus tend to use this style of categorisation, Boeing tend to simplqy matters and state that all defects must be referred to the MEL and the MELprovides the categorisation. In this type of system the ground engineer doesn't even see the minor defects. Each system has its advantages and disadvantages. You need to become familiar with your aircraft, so check on how it works and compare with these notes.

If this sub-routine is unable to correlate for some reason, it will still allow the failure to be displayed and stored (if appropriate) but now it is up to the engineer to try to establish any link.

1 3 - LLJ -

moodull l A-1 506

Storage and Display. This is the sub-routine that uses the flight phase information and decides if a fault is in a reportable phase or not. If it is in a reportable phase it will move the fault into the fault history and usually displays it in a section called present leg fault (sometimes termed inbound faults).

Present leg faults records ALL faults that have happened on this flight leg in the reportable flight phases. This would include active (hard) and latched (soft) faults that have and have not been correlated together with a FDE. If a fault reaches this sub-routine and has not occurred in the reportable phases it will not go into the fault history, it will not be displayed in present leg faults, instead it will be displayed in existing faults for as long as it stays active (hard).

As soon as the fault clears it will be removed. From that you should be able to say that 'existing faults' displays real time information. This makes existing faults very useful for us as maintenance engneers when we are about to dispatch an aircraft. Using this we can see exactly the failures the aircraft has.

- \ -

Fault History (Data Recording) I I I

The fault history will store all defects that have occurred in the reportable phases (the flighlt regon) and will keep them for future analysis.

I

I ' --

You sYould remember from your Air Legislation studies, that we have to keep all aircraft iecords and part of that is to keep a 'reasonable' amount of history on- board ithe aircraft - often as part of the Tech Log. That information is sometimes invaluable to engineering when trying to locate defects that have only occurred in the air. It is also a useful tool for spotting a trend of failure, leading u s to a more correctfault diagnosis first time. -

Much in the same way as engneers look through the aircraft Logbook for the past ten or so sectors, we have the same ability with the electronic information that the CMC has been collecting.

All fault codes and maintenance messages are stored in a memory that is volatile ie, it is lost when power is removed. That would be true, but for the fact that we supply a 28vdc supply direct from the aircraft battery. This supply will keep the memory of the fault history alive for usually about two weeks.

SAQ 4

What is the process for replacing an aircraft main battery? How might that now need modvying in the light of what you have just read?

- 14 -

rnoodull l A-I507

In addition to the change of battery procedures, we must also consider long term parking of the aircraft and hangar maintenance. Long-term parking will drain the battery so we must record the fault history as part of our preparation for parhng atnd/or storage routine.

Hangar maintenance will cause all sorts of faults to be recorded in the fault history for various reasons. Again we need to record the fault history prior to starting in-depth hangar work.

We use the fault history information for fault analysis and to provide assistance when fault finding. The information is displayed, as it was when it first happened. Figures 6 and 7 give two examples, but this information changes from aircraft to aircraft so there is no need to study them in detail.

Fig. 6 FAULT HISTORY OVERVIEW

blank

l r ; - - L O

moodull l A-1508

You must have completed an approved course on the aircraft type before you are given authorisation and EASA PT 66 endorsement (to sign for work done), but in general it is worth noting that the pages all show:

* Date and time of event. * Flight leg or legs (see below). * Fault code. * Maintenance message number (including ATA reference). * Short description of fault. * Possibly flight phase on later systems. * Possibly type of failure ie, hard (active) or soft (latched).

Fig. 7 FAULT HISTORY PAGE

1

When considering fault history it is always useful to know when the fault occurred. For example - the crew reports problems with 'the aircraft flying with constant rudder trim'. We carry out the interuptive BITE check and this reveals nothing - what then? Well, we could check the fault history and could find that this problem has occurred on the last, say, ten flight sectors.

EXTENDED RAINTENANCE

- 16 -

rnoodull l A-1509

OTHER FUNCTION$ HELP

TIME/DATE )/{SOURCE F I E L D

INSTRUCTIONS F I E L D

( 5 1 - NUMBER OF ITEMS

FAULT SUMMARY /'/ GROUP

,'FAULT OCCURRENCE

~ 7 7 7 0 1

I=]}-/? System ARI l lC 629 Bus ( l q f t ,

This Fau l t has occurred i n th. l o l l w i n g Legs: FAULT HISTORY-ATA [I ~ e g 0 Descent 19301 07JUL97 Hard DIALOG BOX, n a i n t c n n s c n.ssage 21-11012

FAULT HISTORY-LEG Buck Cargo V.ntilatlon Fan doer not t o l l a u carrnand

DIALOG BOX (m or o f f ) ar l a overheated

ThIs Fau l t has occurred i n the l o l l au ing Iegr: -

Cl Lap 0 Cruira 15l61 07JUL97 Cl ~ . g - 65 I n l t i a l Clinb 0323r OZAPRW

Haint.naoce nassagr 21-22106 2on. A i r Tmpcratura In.? $ (Zone E ) 3Ign.L 11 SINGLE MAINTENANCE out o f rang.

MESSAGE DISPLAY This ;mutt has occurred I n the t o l l w i n g lag,:

FAULT SuMblARY F I E L D T I T L E 1

MAINTENANCE ' MESSAGE NUMBER '' I MAINTENANCE MESSAGE SYMPTOM I

REPORT T A I L I D

{rput f r a A i r Supply Cabin Presourr con t ro l le r (R) on

The t ime 7s OlLSr OBJUL97 T h i s d a t a Is from L e f t CMCF

S e l e c t t e x t o f a leg, then s e l e c t t h e

NALNTENAHCE MESSAGE DATA b u t t o n to g e t more data. --

'21 c a b i n Temperature C o n t r o l System

J tm ln tmance messap. 21-22101 lare A i r 1smper.tura Snar 1 (FUght Deck) * I p s ( i, out of rang. -, This faul t has occurrrd in the fo l touing Legs:.

~ L J ~ e g u Approach I l?rsr f ~ntarmittent ( 1 ) ) L J ~ b g - I XOLLOUI I L ~ ' I I UOJIJLY~ nara [I Leg - 31 Cruiaa 05261 OSJW97 In tc rm i t ten l (14)

H e i n t ~ n a n s r flc*sage 21-29434. 2

cabin Tenprratura c o n t r o l l r r ( R ChanZ) ha8 no

4 y

A check of the maintenance records might reveal that one of the flight computers was changed after that flight for some other reason. We have used the information in the fault history to accurately analyse a fault that we otherwise could not trace.

Flight Leg Logic

'We can now see that there is a use for fault history. We now need to understand :how the flight legs are defined. This is done electronically. First of all, what is a flight leg? A flight leg is a sector, (that is: doors closed, engines started and aircraft airborne). The flight leg ends when the crew shut-down the engines after ].anding.

All those parameters are sent to the CMC and the CMC then calculates the flight leg. Figure 8 shows the flight leg logic for the Boeing 777. Similar logic is used on other aircraft.

- -- - -

\ - --, - - - ' -

r - .

I

NEXT FLIGHT LEG ' TOBEGIN,WHEN 1 @ 5 ION

, I PARKING BRAKE RELEASED T E = TRANSITION

ENABLE ' 0-- 1 I @ = ;s;:;;:T::;

MANUAL-LEG I N H I B I T - -

Fig. 8 B777 FLIGHT LEG LOGIC

The aircraft's computers count backwards so the current sector is flight leg zero, the one before that is - 0 1, then - 02 and so on.

Looking at the flight leg transistion logic, with the aircraft on the ground and the engine started a new flight leg is started. As the flight leg is started notice that an input is provided to the bi-stable RSQ, that prevents the next transition by removing the logic 1 from the AND gate.

To regain the Enable logic 1, the aircraft has to take-off (park brake released AND e:ngine at take-off thrust AND groundspeed of 80kts). This provides an input to the 'S' leg of the bi-stable and then resets the enable logic.

That concludes this section on the workings of the CMC. It might be worth having a re-read of the text as some of the concepts introduced do make more sense on the second read.

CMC - ADDITIONAL FUNCTIONS

Most people who have a computer or uses one usually come to the conclusion that they never quite get to grips with all its intricacies. Just as you think you have mastered a topic then a new program (software) is introduced or you find out about some additional function that you've never used before. Well, the same holds true for the CMC.

So far in this book we have just discussed the make-up of the CMC and how we can use it to fault find. I n fact we use the on-board maintenance system for many other functions to include:

-

* ' Ground test (interruptive BITE). * +ir/ Ground simulations. * special functions. * Data loading. I

* Printing. ! * Condition monitoring - or ;structural analysis. * Engine fan blade balancing. * ACARS link. I

* DFDR facility. * Maintenance pages (providing real and snapshot information).

Looking at eackone in turn.

Ground Testing

The CMC can be used for a wide variety of testing duties. The CMC is programmed with a series of interuptive BITE checks. A system can be selected for testing and specified interuptive BITE checks carried out.

For this testing to be allowed the aircraft must be on the ground and often airspeed has to be below a certain speed (80 kts). The reason is simply the interuptive BITE test actually electrically disconnects the inputs and outputs and then injects a whole series of inputs and monitors for the correct output/s.

Interuptive BITE has the advantage that it will test the system beyond the 'normal' expected range of inputs. It tests the system to its design limits.

- 18 -

rnoodull l A-151 1

Ground tests are initiated through the normal CMC interface, which as we've already discussed is a CDU (using LSK) or a MAT (Maintenance Access Terminal) using a Windows based menu driven application. Indication of test pass or failure will be given through the normal screen display. Results can be printed off (more of this later).

[The MAT terminal of the B777 consists of a screen, tracker ball (same function as n mouse), and a select button, with the keyboard stowed within the console. It is situated on the flight deck-deck just behind the crew.]

u PMAT

Fig. 9 THE MAT TERMINAL OF THE B777

(With power on the tracker ball is moved to get the MAT screen to come on and selections are made using the selector button controlling a cursor. There is a help menu).

It is true that we had all this when we had intemptive BITE, just by pressing the button on the box. But now we have distinct advantages:

* All system tests (within the scope of the CMC) start and operate from the same point.

* Changing the diagnostic software loaded into the CMC can change the tests, very quickly. (Useful when updated LRU's come along).

I have not included any details of how to initiate ground tests as they vary greatly between aircraft. It is important to understand the principle, however.

Air/ Ground Simulations - -- --

Many systems use airlground sensing for safety and operational reasons. As ground ehgineers, when we test thkse systems we need to put the aircraft in the 'air' withhut actually starting the engines. Traditionally this used to be performed by raisink the aircraft on jacks ok insertingblocks in the WOW (Weight On Wheels) micro switches, or pulling the appropriate C/Bs to the switches to simulate the airborne condition. -

(The WOW micro switches/proximity switches are fitted to thk landing gear shock absorbers to make/break a circuit when oleo deflection takes place so the aircraft 'knows' when it has taken off or landed. Sometimes called squat switches, ground/air sensors etc - see the book in this series on Landing Gear.)

As all systems report to the CMC for diagnostic reasons, we can simulate the air mode from the CMC. From the usual input device (CDU/MAT) we select the simulation screen. This will prompt us to check many things such as we would normally check before a 'traditional' simulation method; these might include, but not limited to, the following:

* Pitot heater CB's tripped. * Drain (water from wash hand basins etc) heaters not active. * RAT (Ram Air Turbine) deployment mechanisms de-activated. * Automatic APU (Auxiliary Power Unit) start systems de-activated. * Spoiler systems de-activated. * Ground lock pins fitted and secure.

etc

- 20 -

rnoodull l A-1513

The use of this function is usually limited to one of the CMC input devices, called the master. This is to ensure that there is no inadvertent operation of the simulation, which could prove expensive, not to mention dangerous. A s a point of nlote, the CMC master input device has to be out of the "mainstream" flight crew use. It must not, for example, be at the CDU that they normally use.

Special Functions

Special functions include a host of various features that, as ground engineers, can be useful. The airlground simulation is a form of special function. Special fu~nctions change from aircraft to aircraft but can include such tasks as:

* Flying control safety device override. Some flying controls are locked on the ground for safety reasons, alternate flaps is a specific example. To get these to operate this function is used.

* Door lock activation. Modem aircraft have flight locks on the passenger -doors to prevent them from being opened when the aircraft is airborne.

* Airyground target calibration. \ 'i I '

* Proximity sensing device calibration. All proxirnity-sensing devices have a 'range' where they come inf .~ /~roximi t~ , ' (o~era te ) . Using this type of function we can set up exactly whe$ the devi6e hctivates, removing the grey area.

, , --- l 1 - \

-- - The list cbntinues, but you should ibe starti,ng to get the hangof-the principle.

1

Data Load ,, I , \ _ - -- --

r //

I 1

1- -

A data-loading-facility is required for a-number of reasons: - -

* To load new software into various system LRU7s. * To extract variable data from the CMC such as fault history data. * To speak directly with the FDR if required. * To load route information. * To load navigational database(s).

Data loading is in fact read or write information, not simply a loading (write) facility.

Originally data load facility was provided by an external test set that plugged into a test connection, often on the front of the main aircraft computing system. Data transfer would then be from/to a magnetic tape in the test set. The set would convert the magnetic information into coded decimal and often something like

code was used.

Clock Pulse

Data Word r Power spike cancelled by transmitting reflected -+I

Data word wave form.

I 1 Stretched portion of tape. If no clock pulse this would indicate 4 logic 1's instead of the correct 2 logic 1's

Fig. 10 DATA TRANSMISSION

This type of data transmission was always transmitted in parallel (see figure 10) to ensure that any stretches of the tape were not seen as logic zero. This method was cumbersome, slow and required large test equipment.

This method is still used today on some aircraft but most now use other systems ,such as the 1.44MB diskette (floppy disk) that we see in use on all PC systeqs 'today.

1 1

somewhere on the flight deck (on the MAT terrninal for example), close to the master CDU (usually the centre one for maintenance reasons) there will be a data-load facility. 1.44MB diskettes can be loaded into the drive and then through the normal input of the CMC we can install the data to the correct location.

A s data uploads sometimes introduce operational effect changes, a normal procedure would be to load the software across all the fleet a t the same time (if the operator has many aircraft of the same model and type). The loaded software is held in the CMC storage device until such time as it is required and then it is transferred to the LRU concerned.

This addition provides us with the benefit that we have all the software to hand on-board should something happen that corrupts a particular program.

(Any one who owns of has a lot to do with a computer knows have annoyingly easy this can happen.)

Prior to loading any software we need to ensure certain parameters are cleared. Firstly we must satisfy ourselves of the origin of the data and the fact that it is virus free. We must then check that the software we are about to load is the correct part number and is compatible with the LRU we are about to load it into.


Finally after load we must ensure that the correct entries are made in the aircraft log to ensure all data is updated and correct.

Incidentally all aircraft that have a centralised maintenance facility have to have a copy of the correct software on the flight deck. This is usually carried as part of the Tech Log.

Data loads for the other functions are performed in much the same manner. Fault history can be read and copied from the CMC and reloaded again, this is useful for the reasons we stated when we looked a t fault history. Navigational d-atabases provide information on the airports and navigational aids worldwide.

You may also see other (minor) information loaded into the CMC in the same nnanner.

F'rinting

Most aircraft with these type pf systems have an on-boardprinter. . 1 - \ -

' I-- - -- ' On a ~ M C generated screen, whether thdt i's through a multi-function display or a sp&cific screen, the information can be printed if requirkdr This ynables hard copy information to be take; away from the aircraft to re,search fablts in the FIM etc. Figures 11 and 12 show-a-typical thermal and schematic diagram. i I - - \ \

1 1 I - - / \

I I I - - The p in te r is usdally located on the flighd dLck centre stand. Both flight crew and the kngineers have access 40 it. ~nfonhation is sent to the printer from the C!MC (NOTE thk diagram shows the ~ 7 7 7 ' l a ~ o u t and on that aircraft the CMC residzs wTthin AIMS). This i n f o ~ m & t i o n ~ ~ ~ u a l l y comes in ARINCA429 b i n a g coded decimal form, but you Gy see it in other forms also.

Inside the printer the processing board converts the data into the language the printer uses. This is fed to the interconnection board, which controls the printing process. At the correct sequence the head moves and is heated, this passes over thermally sensitive paper, which is moved by a stepper motor.

Defects of the printer are notified by way of lamps on the front of the panel. The printer normally doesn't fault report to the CMC; it is part of the CMC and so is not really necessary. Paper rolls are loaded in a number of ways but the instructions are always located in the AMM arid often on the inside of the printer top cover. A s the paper is thermally activated it must be kept away from heat and direct sunlight.

blank


r-

Fig. 11 FLIGHT DECK PRINTER I ,

SUPPLY

Fig. 12 PRINTER SCHEMATIC

'-)A - L - r

rnooduil l A-1517

Condition Monitoring - or Structural Analysis

Condition monitoring was introduced to satisfy the demands of the regulatory authorities for information and proof of reliability for ETOPS (Extended range 'I'win engined aircraft Operations). The authorities required information primarily about engine reliability, but also about other systems and the structure of the aircraft. So the Aircraft Monitoring System (ACMS) was born.

The CMC is already receiving information from all the systems on the aircraft and the engines are no exception. Various parameters and triggers are set and this information is collected by the CMC and sent onto the ACMS. The type of information that the ACMF will take depends on how it is being used but is likely to include:

* In flight engine shut-down information. * Filter changes - yes the CMC can work out when an engine filter is

changed by checking a data chip inside it. * Structural loading (if appropriate). * Routes flown. - - -- -- - -- --

1

* Airfr-ame hours and lanaings. - I - -

I \ I

This list is by no means cornplete'as virtually any piece of dAta that the CMC has can be requested by the AC'MF, but stofage tends to be a limiting fitctor. When the ACMF has the data it requires' it will transmit this to an optical recorder, known as the Quick qcqess Recorder (QAR). A typical 'penny and Giles7 unit is shown in figure 13. Maintenan'ce of this systek ip really limited to replacing the diseevery ramp c p c k as spkcified in the AMS (Aircraft ~a in t endnce ~chkdule) .

/ I l 1 ,

The procedure for replacing the d i s ~ i s i , / dl I -_ - - - - -

Draw a new item from stores, checking its applicability, validity and stores release documentation. Check also that it is empty. Wear lint free gloves. Remove the new disc from the case and hold it carefully. With aircraft power applied press the eject button and remove the 'used' disc and place it in the case, in which you have received the new one. Remove the spare disc and place that in the active slot. Place the new empty disc in the spare slot. Remove gloves and label the used disc stating the aircraft serial number, the date of removal and the reasons why. Return the disc through the correct channels to the data analysis unit. Carry out the necessary aircraft documentation.

As yet, no civil aircraft have fatigue meters and so this data is not requested.

LOW CAPACITY

EJECT BUTTON

, SPARE DISK

OPTICAL OAR

PENNY &

G I L E S

/

M A D E IN U.K.

ACTIVE DISK SLOT (BEHIND DOOR)

SLOT POWER ON INDICATOR-,

I Fig. 13 OPTICAL RECOIZDER

C PRESS SPARE DISK

Before we finish this topic, I shall cover some functions of the CMC that we can only describe as bonuses. The system was never designed around these but they do prove increasingly useful.

- - -

Engine Fan Blade Balancing

The engine EEC sends information to the CMC through the normal course of a flight. Engines have vibration motoring systems installed and this information is available to the CMC. The CMC is programmed to check the 'high7 points of the engine acceleration and deceleration and from that it can calculate mathematical where balance weights are needed if required. It is recommended that you read the gas turbine engine notes and particularly those relating to engine balancing and instrumentation.

ACARS Link

All the information the CMC has can be transmitted to anywhere in the world either via VHF or Satellite communication networks. This means that should the aircraft have a significant failure the aircrew can select this function and let the ground crew at the receiving station know details of the failure prior to the aircraft touching down.

- %(j -

moodull lA-1519

'This means that fault diagnosis can begin and in some cases be completed prior to aircraft landing. Spares and equipment can also be prepared so rectification and turn-round can be accomplished with minimum disruption of the flight schedules.

Some software faults can be cleared whilst the aircraft is in the air with the aircrew in communication with engineers on the ground. Several airlines have Faults Cells manned by engineers around the globe, say one a t LHR and one at San Francisco, where data on airborne aircraft faults is sent and a n y necessary actions taken.

Actions can include discussions with the aircrew on how to rectify the fault and forward planning can be carried out a t the destination airfield re spares availability, manpower availability etc.

IIFDR Facility

( This is-similar to the ACMF. The Digital-Flight Data Recorder has-to be carried on allcom&ercial air transport aircraft; it is sometimes referred to as the 'black box' - though they are usually 6right orange for crash scen6 identification purposes. They are located at the back of'the aircraft and have son&-locating beacons attached to aid undedat'er recovev.

I / I ' I

The F Q R ~ ~ simply a recording devic~thdkecords a whole y a y of parameters like thiottle settings, thrust conkhanded,\flight control positions,-aircraft attitude, altitudgttc. The A N 0 states thatlit .shall be in use from the beginning of the t4e-off y n to the end ofithe landing run and that thk dperator shall preserve ht least the last 25 hours of recokdihg.

i - - \ - -- - /

,J

- -

All this information is by defaUE ai&liiGe at the CMC and itpis %ow a case of converting the CMC language into BCD language that the recorder understands. The CMC converts the signal and sends it to the FDR.

Another advantage of this system is that the FDR is now also in direct communication with the CMC and can put up failure codes to inform the crew if the FDR is not functioning properly.

hlaintenance Pages

This is not strictly speaking a function of the CMC, but I would suggest that it does fall into the category of on-board maintenance. The screens in front of the flight crew are divided u p according to function. The outer two are always Primary Flight Displays (PFD); the next two in are used normally for Navigational Display (ND), but are really called multi-function displays.

The two displays in the centre, the top one is the ECAM/EICAS which provides information regarding the engines, and operating parameters of the aircraft that the flight crew in particular are interested in.

- 27 - moodull 1A-1520

The lower screen is another multi-function display that is used for systems such as hydraulics, pressurisation, fuel, electrical supplies etc.

Figure 14 show the instrument panel of the A320.

All screens are in colour.

Master warning It ,- Side stick priority It Master caution It] , Stby ATT - EFlS CTL

LDG GEAR IND AUTO BRK/A-SKID CTL

Chart holder Optional / , I Clock \ metric ALT LDG

I GEAR I

DME / VOR CTL LVR

DDRMI I

Optional ADF RMI - Possible exchange of images' I

Fig. 14 FLIGHT DECK INSTRUMENT PANEL - A320

Some of the displays are interchangeable, but the PFD and the EICASIECAM have their function assigned and cannot be used for anything else. In fact their importance is such that they will 'steal' another screen automatically should they fail.

Weather radar displays can also be superimposed on the same screens.

We can call up a variety of synoptic pages on the systems and figure 15 shows an example of two displays - one for the hydraulic system and the other for the electrical system. These pages show a diagrammatic representation of the system with indications of parameters such as pressure, flow, temperature, voltage, frequency etc as appropriate.

We can also, through the input device to the CMC, call up maintenance pages (see figure 16). These give actual values for inputs and outputs to various systems. These screens can provide a good insight into a system when fault finding.

E55 TR EnER GEN v

TRT +I9 'C

HYD SREEN - BLUE YELLOH

9 Q

i R T +19 ' c G . W . 60 3.r SRT + 1 8 ' ~ I CMT 2 3 ~ 5 6 1 C . ; . 28 r

Fig. 15 ECAM DISPLAYS - EXAMPLES --

- - 1- -

- - I I

. \ -

So much so that the CMC hold; a series of trigger values and i f any input moves beyond the trigger the screen is snapshot. This is stored in memory us to view and/or print when the aircraft lands.

I -

I I /

SYSTEM DATA '

for

1 I I

. 1

I J AUTO EVENT HESSAGE F I E L D

l 1 -

Fig. 16 MAINTENANCE PAGES

I

UmLLBua nus

TITLE \

u1 c I

YES 1 YES YES NO NO n o

AUlO AUTO 1 0 7 4 Nun NORM HORN

_- -- , -

YES FAST

YES

UP NO

no YES YES YES YES

YES YES

- AUTO -

,, ULLar PG 211

M,U.

L

a 12.03 r

ID 12.03 12.03 n -T2\ I< 12.03 12.03

rr 12.03 I

+3.U t3.36 I +13M +ISM

a.75 1r +0.75 ' rr n ts t I rl620

\

ulmLmun - 3 0

.or, trrc -0.70 trr nm +2.39 rnnat 2l iW *OL$ l * ~ c -0.18 -I 0' t5.23 CII

10l.y 257

LCC ~1 ~2 L

I 1 .!ru..n M.15 '0.15 t0.15 4 . 1 5

c : t r U . L I;,': ".Y

wrlroll I +5.36 +5.57 -0.42 +5.36 tJ.57 4.42 2 G.36 r3.57 4.42 t5.36 +3.5? -3.42 3 C5.36 +3.57 +5.36 63.57

-

P YES UP P YES UP

17 JAN 95 mc 1 8 : 4 4 ?r

1-c6 3 +0.05 M.05 rO.05 W.05 2 +O.OS rO.05

IRm M 1 n-r IW 1111 ~ 4 w I IL YPI.EUII.

rorrlr- l t9.W a.05 UJ.03 n l r l ln +1.2 a C0.W *c 05 rO.05 m +1.2 a t0.m 4 t0.W)

P f t s wc. m 1 1 0 1 M O T NODE A C T I V I T Y : NORMAL DISEHGAGED OVERSPEED M O T ACT

,1620 n -1620

1 7 nNA1c4 .WITIDI l l l l 1 - z - l

+2.35 +I230 +2.35 11 +iZJ0

I 17 JAN 95 vrc 1B:U:J~ \ I

FLARE C C W BANK ANGLE PRO1

orrt 17 JAN 95 w< l 8 : 4 4 : 3 & 1

/

DATE AND T INE /

There is usually a facility to take manual snapshots by pressing a record button located on the flight deck somewhere. The flight crew a t our request can press this if we are trying to trace a fault that seems to only appear in the air. Also, we can use the system ourselves to record information during an engine ground run for example.

On-board Electronic Manual Systems

With the use of Windows based maintenance applications on the access terminals, it is becoming possible to load and use maintenance manuals electronically through the CMC. Currently no aircraft has the facility in use due to complexities associated with its installation, but CDs for AMM, SRM, FIM etc are available for use in desk top/lap top PCs and belt worn PCs.

The belt worn PC is a recent innovation with the computer strapped to the belt and a one inch liquid crystal display strapped to the head. The computer is voice commanded which leaves both hands free to 'get on with the work'.

'On-boqd' manuals will afford u s the opportunity to look at the FIM and the AMM without actually leaving the aircraft. Caution needs to be taken when using this style of information that it is in fact current and that any temporary revisions are also incorporated. That may well mean a trip to the technical library in any case. I

H E A D STRAP

EAR PHONE

COMPUTER UNIT

Fig. 17 THE 'TREKKER' NOTEBOOK COMPUTER

We need to be careful, with all this technology, that we don't become blinkered during fault diagnosis. In a way it is like using a calculator, we must have some idea of what the expected result should be. If the calculator gives u s a result that is wildly different from what we expected then we should consider the possibility if mis-use of the calculator - incorrect keying etc. The same with electronic fault diagnosis, we should think what the computer is telling us what to do and it should make sense with our own logical reasoning.

Don't forget that ultimately you will still have to leave the flight deck and go to the actual system to carry out what-ever rectification is called for.

That concludes this section on the additional functions of the CMC and indeed our study about on-board maintenance and ATA 45, a new chapter and concept to some of us. Try the following exercises. The answers are found in the text.

7,?7),,>)''7,,),

rnoodull l A-1523

1. What is BITE? 2. What is a central maintenance LED panel and how does it operate? 3. How, basically, does the CMC work? 4. Explain where and how the CMC displays its failure codes. 5. How do you fault find using the CMC and FIM? 6. What is a flight leg? 7. Explain the reportable region and the logic? 8. Why and what do we 'correlate'? 9. Explain the route through the CMC that a fault code takes. 10. Briefly state what faults you would expect to see in present leg

faults (inbound faults) and existing faults? 1 1. Describe what you would expect to see in the fault history. 12. What is a ground test? 13. Name some of the functions of the data loading capability. 14. What type of printers are normally used? 15. How do we change a QAR disc and what sort of disc is it? 16.- What is a snapshot-and when,might we expecpto use one? 17. As with all cornpufer systems the only way to get farnili& with

them is to use the system. If hossible, and with supervision - ' unless you are alreAdy cleared to work on your pikcraft - have a I look a t the CMC pages - if you have access. Get1 to grips with the

basic layout and how to navigate your way arounq. If y,ou do not I have access then re-read-any parts of this book on-areas/you are

unsure. 1 \ \ I I - -

18. Look' in ATA45 for aircraft (if possible) a n d s8e how fault information is correlated and prbcessed. Check t o see what is reportable and what isn't, , I

1 - I - - - - -- - - -


S A Q 1. From a maintenance point of view we need to be mindful of what is going to move, this is what is sometimes termed 'aircraft effects'. For example, if we are carrying out a spoiler interruptive BITE test we need to make sure the spoilers are clear of equipment and personnel as they will move. Also from a design point of view think what would happen to the aircraft if the same spoiler test were initiated in the air? - a serious aircraft disturbance would occur with possible fatal consequences. Therefore, we would expect the designers to incorporate air/ground sensing protection in the BITE routines.

S,4Q 2. Some systems do not operate normally on the ground. Pressurisation for example (and there are many others). Only when the aircraft is airborne do these faults become apparent and by the time the aircraft has landed the fault will have cleared. We should notice these sorts of failures as 'soft' failures and we could then manually initiate a BITE test to check the system out.

SAQ 3 . During flight it is safe to say that we want all information stored, but there are times when we might want to block some information. During maintenance tests for example, we generate hundreds of failure codes that we really do not want to keep on record. Engine start is another time when you might consider blocking some failure signals. There are many others.

SAQ 4. The process for replacing a battery is given in the electrical section of the LBP study boolts. On modern digital aircraft we must consider the fault history. To remove the power and the battery, will erase any history that is stored. We must now change the battery with power applied.


APPENDIX

GLOSSARY OF TERMS

ACARS Aircraft Communication and Reporting System. A system that allows aircraft to report to the ground defects in advance to allow rectification to begin as soon as the aircraft taxies in.

ACMF Aircraft Conditioning Monitoring Function (or System). A process that allows the CMC to check vital signs on the aircraft and report them to the appropriate authorities.

AGS Air Ground Sensing. Many systems on the aircraft must not work in the air or must not work on the ground. Switches on the aircraft landing gear allow the systems to know when the aircraft is airborne or on the ground.

LIMS Approved Maintenance Schedule. A maintenance schedule that has b e e n submitted by the operator-and approved by the-regulatory

authority (the CAA in tKe UK) . - , I I

I

1

ARINC Air Radio INCorporated. A standarb electronic data communication ' I I system used on a i r~ ra f t .~ I

I I ' 1 /

I - - / 1

ATA ~ : r Transport ~ssociatioh. ATA 100'~rovides all aircreft Tr i iua l s with a standard format l recogniieh worldwide. 1 1 -

1 I ' I I

AMM Aircraft Maintenance Manual, w h i ~ h !provides technikq data about the I

a aircraft': - 1 1 I , J

- -- / - -

E3IOS Built In Operating System. Technically software, its function is to start the computer and start the process of loading the main operating system.

EHT Built In Test.

BITE Built In Test Equipment. A system built into a component or system to provide a means of self-testing.

CDU Control and Display Unit. A device in the flight deck that allows personnel to communicate with the various computers on board. Some information is also displayed on the screen and selections are made using the line select keys.

CFDS Centralised Fault Display System. An Airbus system essentially similar to CMC.

CMC Centralised Maintenance Computer. A computer connected to the all the important systems/components on the aircraft that receives data relevant to the airworthiness of the system/component.

CMCF Centralised Maintenance Computing Function. Some manufacturers are now installing the CMC as a computing sub-routine of the main aircraft computing system. Operates in the same manner as the CMC.

DFDR Digital Flight Data Recorder. With the advent of CMC, it became possible to supply the flight recorder by using information that has already gone to the CMC. ARINC 427 communication is often used to perform this function.

ECAM Electronic Centralised Aircraft Monitoring. An Airbus electronic display system similar to EICAS (Boeing).

EICAS Engine Indicating and Crew Alerting System.

ETOPS Extended Twin range Operations. Specific conditions (which are more strict that when operating a four engined aircraft over the same route) apply to twin engined aircraft that fly for extended periods.

FDE- Flight Deck Effect. -

/

FIM Fault Isolation Manual.

! FMC Flight Management Computer. Either a stand-alone, computer, or a sub

routine of the main aircraft computer, the FMC is loaded with all the route information and navigational databases. In addition to navigation, it provides such information as top of descent; some of the flight phases are triggered by the FMC.

I I

FRM Fault Reporting Manual. A reference guide that the aircrew might use to hrovide ground engineering with a fault code. Not used much these days as BITE usually provides that information.

IRS Inertial Reference System. A laser gyro system, that provides position and attitude information.

LED Light Emitting Diode - see basic electronic notes for details.

LRU Line Replaceable Unit. Nowadays, virtually all avionics components can be changed on the line' by the action of removing one box, or one item, and inserting another. These 'boxes' usually have quick connections at the rear and are tray mounted.

LSK Line Select Key. Each CDU will have a series of these. On the display an arrow points to the LSK. To select that particular option press the LSK.

MAT Maintenance Access Terminal. A specific computer provided in the flight deck for maintenance reasons. It has a similar function to the MCDU, but uses a Windows style menu driven operating system and is much more user friendly.

MCDU Master CDU. The main CDU and the only one that can be used for certain maintenance functions. See also CDU.

PCU Powered Control Unit. A hydraulic actuator that converts the pilot's inputs be they electrical or mechanical into hydraulic pressure to move the flying control surfaces.

I T C U See PCU

QAR Quick Access Recorder. An optical device that is used to remove the specific data required by the regulatory authority for aircraft monitoring.

RAM Random Access Memory. A memory in a computer that is used for running programs and storing data. Normally volatile, which means it erases at power-down.

\7HF Very High Frequency communication.

Documents

M11 Aerodynamcis,Structures and Instruments 2 Of2