Aircraft Training Manuel LET 410 UVP-E

/ Nepod%hj ziEdci161n~ AIRCRAFT TRAINING h/WNUAL

FOR THE AEROPLANE L 41 0 UVP-E, L 41 0 UVP-E9,

L 41 0 UVP-E20.

Issued January 21,1998

LET, a.s. 686 04 Kunovice

CZECH REPUBLIC

LET, as . KUNOVICE, CZECH REPUBLIC

Approved by:

........*........................ Ing. M. PeSak

Chief Designer

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, E20

January 21, I998

We cannot correct an error unless we know of its existence, threfore, it is essential that you do your part. Comments, corrections regarding this manual are welcomed and should be sent to:

Documentation Department LET, as. 686 04 Kunovice Czech Republic EUROPE

or fax us to:

Documentation Department +42063261352

or e-mail us to:

O 1998 LET, a.s., 686 04 Kunovice, Czech Republic

All rights reserved.

No part of this manual may be reproduced or transmitted in any form or by any means, electronic or mechanical, including photocopying and recording, for any purpose without the express written permission of LET, as. I


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AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E20

- Rev. No.

-

-

RECORD OF REVISIONS

Numbers of PE

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)perator's Logging leference and Date

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Revision

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- Rev. No.

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INTRODUCTION L-410 UVP-E, E9, E20 AIRCRAFT TRAINING MANUAL contains brief technical description of the

aircraft and its systems, makes one familiar with maitenance system, checks and aircraft

servicing. More detailed technical description and aircraft maitenance are given in the

L-410 UVP-E, E9, E20 Operating Manual and L-410 UVP-E, E9, E20 Maintenance Manual.

L-410 UVP-E, E9, E20 AIRCRAFT TRAINING MANUAL has been compiled mostly from the

Documentation Department documents. Taking into account the fact that L-410 UVP-E, E9, E20

is currently modified the text may differ from the real state of the Aircraft.

There is not described individual modification of aircraft (for example L 410 UVP-E3,

L 410 UVP-El 0)

The L-410 UVP-E, E9, E20 AIRCRAFT TRAINING MANUAL is intended for training purposes

only, it is of informative nature and this may not be used as documentation for aircraft operation.

That is why it is not permitted to distribute this TRAINING MANUAL whithout consent of LET

Kunovice Training Center Anthonty.

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9. E2O

Intentionally left blank

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9. EZO

CONTENTS CHAPTER 1 STRUCTURES

1 .l. General

1.2. Fuselage

1.3. Doors

1.4. Windows

1.5. Wing

1.6. Engine nacelles

1.7. Stabilizers

1.8. Technical summary

CHAPTER 2 FLIGHT CONTROLS

2.1. General

2.2. Elevator control

2.3. Rudder control

2.4. Aileron control

2.5. Elevator trim tab control

2.6. Rudder trim tab control

2.7. Aileron trim tab control

2.8. Spoilers

2.9. Automatic bank control tabs

2.10. Flaps

2.1 1. Autopilot servos

CHAPTER 3 LANDING GEAR

3.1. General

3.2. Extension and retraction

3.3. Main wheel braking

3.4. Steering

3.5. Position and warning

CHAPTER 4 AIR - CONDITIONING

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, €20

--

4.1. General

4.2. Distribution of air

4.3. Heating

4.4. Temperature control

4.5. Indication

CHAPTER 5 ICE AND RAIN PROTECTION

5.1. General

5.2. Pneumatic deicing system

5.3. Air intakes deicing

5.4. Heads pressure deicing

5.5. Windshield and wiper unit deicing

5.6. Propellers deicing

5.7. Deicing detection

CHAPTER 6 HYDRAULIC POWER

6.1. General

6.2. Main hydraulic system

6.3. Normal control

6.4. Emergency control

6.5. Hydraulic tank pressurization system

6.6. Indicating

CHAPTER 7 FUEL AND OIL SYSTEM

7.1. Fuel system - general

7.2. Fuel distribution

7.3. Fuel tank venting

7.4. Fuel tank interconnection

7.5. Fuel drainage

7.6. Fuel indicating

7.7. Oil system - general

7.8. Oil system indicating


CHAPTER 8 FIRE PROTECTION

8.1. General

8.2. Detection

8.3. Extinguishing

8.4. Fire protection of AC generator cooling air intake

CHAPTER 9 ELECTRICAL POWER

9.1. General

9.2. AC generation

9.3. DC current sources

9.4. External power

9.5. Electrical load distribution

9.6. Electric system of airplane systems

9.7. Lighting system

9.8. Flight compartment lighting

9.9. Exterior lighting

9.10. Emergency lighting

CHAPTER 10 COMMUNICATION AND NAVIGATION

10.1. Communication - general

10.2. VHF transceiver

10.3. SW transceiver

10.4. Passenger address system

10.5. lnterphone

10.6. Device for sound record

10.7. Static dischargers

10.8. Navigation - general

10.9. Gyro horizons and turn and bank indicators

10.10. Compass

10.1 1. Gyro magnetic compasses

10.12. VORIILStMKR navigation sys'em


10.1 3. ADF - automatic direction finder

10.1 4. Radar altimeter

10.1 5. DME interrogator

10.16. Transponder

10.1 7. Weather radar

10.18. EFlS - Electronic Flight Instrumentation System

10.1 9. Flight recorder BUR - 1 - 2G

10.20. Autopilot

CHAPTER 11 PlTOT - STATIC SYSTEM

1 1 .l. General

CHAPTER 12 EQUIPMENTI FURNISHINGS

12.1. General

12.2. Flight compartment and pilots seats

12.3. Instrument and control panels

12.4. Passenger compartment and passengers seats

12.5. Portable oxygen equipment

12.6. Toilet

12.7. Baggage compartments

12.8. Emergency equipment

CHAPTER 13 POWER PLANT

1 3.1. General

13.2. Fixing of the engine

13.3. Fireseals

13.4. Drains

13.5. Engine controls

13.6. Emergency shutdown

13.7. Indicating

13.8. Engine starting


CHAPTER 14 PROPELLERS

14.1. General1 4.2. Propellers controlling1 4.3. Indication

CHAPTER 15 WATER INJECTION

15.1. General

15.2. Water distribution

15.3. Dumping and purging

15.4. Indicating

CHAPTER 16 SERVICING

16.1. Lifting and shoring

16.2. Levelling

16.3. Weighing

16.4. Minimum turning radius when aeroplane taxiing on the ground

16.5. Towing

16.6. Parking and mooring

16.7. Exterior marking

16.8, Interior placards and markings

16.9. Airfield servicing

16.10. Servicing in emergency situations

16.1 1. Ground equipment and tools

16.1 2. Airplane maintenance

CHAPTER 17 AIRCRAFT CHARACTERISTICS

17.1. Basic Performance

17.2. Flight Technical Performance

17.3. Operation Performance


Intenticmalty left blank


CHAPTER 1

STRUCTURES

1.1. General

1.2. Fuselage

1.3. Doors

1.4. Windows

1.5. Wing

1.6. Engine nacelles

1.7. Stabilizers

1.8. Technical summary



1 . A . GENERAL The L 41 0 UVP-E aircraft is intended for the transport of passangers, mail and cargo and is able

to take-off and land on hard as well as unpaved runways (the aeroplane can be thelivaled in flying

ambulance and parachute version by customer request).

The L 410 UVP-E is all metal semimonocoque upper wing design. Two turboprop

WALTER M 601 E engines with V 510 five blade props are built in the aircraft.

1.2. FUSELAGE

The fuselage is an all-metal panel structur and consist of the following independent units: nose

section, centre section, rear section and landing gear nacelles.

In the nose section there is a cockpit separated from the passenger cabin by folding safety

partition having the wings that can be secured only from the cockpit side and locked by a special

key from the passenger cabin during the parking. At customer s choice the aircraft can be

equipment with bullet proof armoured door.

The fuselage nose section is closed with removable covers for securing the access to the

equipment located in this section.

The front baggage compartment is located in the fuselage nose section and is accessible trom the

outside.

The passenger compartment is located in the centre section of the fuselage. It is fitted with seat

for 19 passengers It is separated from the passenger compartment by a fixed partition with folding

wings.

The baggage compartment and toilet are located in the fuselage rear section. The fuselage is

closed by a removable plastic tail cone and ventral fin.

Fuselage nose section

In the nose section the main frame is composed of frames No. I , 2, 3, 4, 5, 5A, 6 and 7, of system

of riveted grids between the frames and of system of continuous longitudinal stringers divided at

the frame No. 5. The longitudinal stringers are divided in the connection in the centre section at

the frame No. 8. At the fuselage outline the longitudinal stringers are riveted together with the

grids (floor and cabin ceiling grids).

A part of the main frame is the window frame of the flight compartment in the nose section,

between the bulkheads No. 4 through 7. This window frame is manufactured of bent duralumin

profiles. The structure parts of the fuselage nose section are mutually connected together with

rivets and screw connections. On spot welding is used some subassemblies there (with panels).

On the fuselage RH side there is placed the emergency exit closed with door (Fig. 2) suspended

on two hinges with vertical pins and is openable outwards. The nose section is provided with re

movable covers for securing the access to the equipment placed in the fuselage nose section. At

the frame No. 2 an external power supply connector is located. Below bulkhead No. 1 tiltable with

sight hole of organic glass for searchlights,is placed radom shaped for function of meteorological

radar installed on bulkhead No. 1.


Fuselage centre section

The fuselage centre section main frame is compcsed of frames No. 8 through 18 out of which the

frames No. 8, 12, 14, 18 are the basic elements of the stressed structure, system of floor grids,

system of logitudinal stringers and door frames for the passengers and cargo.

The frames 12 and 14 are composed of milled segments interconnected with milled spar,

extended on both sides of the fuselage with sliding sleeves at the ends of the spars in which the

landing gear is placed. On the upper parts of the segments there are fixed steel suspensions for

attachment of the fuselage and the wing by screw passing through shims and the skin.

The frames 8 and 18 are interconnecting frames for connecting with the fuselage nose resp. rear

section. They are composed of rolled profiles and sheet webs.

The other frames make parts of subassemblies of the panels of the fuselage centre section. In the

part between the frame 15 and 18 on the LH side of the fuselage there is a doorframe formed by a

dura limin profile of U-shape, composed of four parts.

The floor grids are laid-out asymmetrically along the fusetage axis, they are fixed through the

longitudinal stringers to the skin, in the top part they are reinforced with T-profiles which serve to

fix the seats and the floors. At anchoring points for attachment of the seats the grid sheets are

mostly provided with vertical pressed grooves and reinforced with a stiffener from the other side

with vertical pressed groove too. This arrangement increases the stability of the grid webs as well

as floor sheets and makes it possible to use bolts of any length for attachment of the seats. At the

bottom and,the longitudinal stringers are divided according to the division of floor grids; at other

places they are divided according to the skin panels.

Fuselage rear section

The rear section main frame is composed of a system of longitudinal stringers and bulkheads in

which these framework elements are interconnected to form respective parts of the fuselage rear

section. The bulkhead No. 21 closes the cabin space and is provided with an opening for the toilet

location. After removing the toilet a manhole for entering the fuselage rear section is available. In

the axis of the frames 25 and 26 there are milled pylons which make as well the suspension of

horizontal and vertical tail unit and with further structure of the tail superstructure it make fairing

between the fin and fuselage outline .The rear section framework is closed with the frame No. 27

provided with angle irons around the circumference with 24 riveted nuts for attachment of glass

reinforced plastic tail cone.

Landing gear nacelles

Landing gear nacelle is formed by a framework, skin, removable covers and landing gear doors.

After assembling these parts there arises characteristical shape of aerodynamic profile with

leading and trailing edges.

The framework of the main landing gear nacelle is formed by a system of logitudinal and lateral

ribs. The lateral ribs are divided to the upper and bottom ones. To fix these ribs there are used the

spars of the main landing gear which at the same time form the centre part o f t he lateral ribs.

The spars of the main landing gear are connected together with a reinforced platform to which the

bracket for fixing the main landing gear lock is attached in the middle at the fuselage. Longitudinal

ribs are divided to front, centre and rear ones and are rigidly connected with lateral ribs.

The flank and the bottom parts of the landing gear nacelle are provided with door that extends

from the front spar up to trailing edge of the landing gear nacelle.

FIG. 1-1 DRAWING OF THE FUSELAGE SYSTEM

(Circled numbers indicate the frames numbers)

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, EZO

1.3. DOORS

Main door

The aeroplane can be boarded through the main door on the LH side of the fuselage behind the

single seats. The main door is divided: One part of the door can be tilted upwards by means of a

lever which also fixes the door in open position, the other part of the door is secured closed and is

used in aeroplane cargo modification only. Boarding the aeroplane is enabled by boarding stairs.

Emergency exit

The front emergencyexit (emergency exitdoor) is situated in the cockpit on the right hand side of

the fuselage front section (in the section between frames No.6 and 8). Emergency exits in the

passenger compartment (emergency exit door) are situated on the left and right hand sides of the

passenger compartment between frames No. 13 and 14 (if installed).

Luggagee compartment door

The front luggage compartment door is situated on the fuselage left and right hand side between

frames No. 2 and 4. The rear luggage compartment door is situated on the fuselage right hand

side between frames No. 19 and 21 (if installed).

Landing gear doors

To cover the nose landing gear bay (the landing gear in retracted position) there are doors at the

bottom side of the fuselage nose part. Tilting doors hinged on the left and right landing gear - - - - - - - - - - - - - - - ~ ~

nacelles serve to cover partlythe main landing gear bays.

Toilet compartment door (if installed)

The toilet compartment door is placed in the rear part of the passenger cabin.

Door on the frame No. 21 (if installed)

The door hinged on the LH side of the frame No. 21 by three hinges provides acces to the

fuselage rear section.


3 f-7 I f i n s t a l l e c

T TURN 3. PULL

I f installed /

Fig. 1-2 THE WAY OF OPENING THE MAIN DOOR, COCKPIT EMERGENCY EXIT

DOOR,PASSENGER CABIN DOOR AND REAR BAGGAGE COMPARTMENT

DOOR OPENED FROM OUTSIDE


C LCS ED

3. TURN

OPEN

1 . REMOVE 2. PUSH AND HOLD

I f insLaI led 3

PUSH AND HOLD

I

A - LOCKED B - UNLOCKED

Fig. 1-3 THE PROCEDURE OF OPENING THE MAIN ENTRANCE DOOR,

EMERGENCY EXIT DOOR, COCKPIT AND PASSENGER CABIN DOOR FROM

INSIDE


Door warning

When the main, emergency exit, front baggage compartment doors, and rear baggage

compartment door (if installed) are not shut it is indicated in the cockpit by the illumination of the

DOOR signal on the central warning display - see Fig. 1-4. The main door is provided with one or

three (if installed) terminal switches which control the main door warning circuit. The movement of

one of there terminal switches is derived from the closed position of the door handle, the

movement of the remaining two switches derives from the movement of the locking rods when

they slide into the door frame. Locking rod position indicators (if installed) provide for visual check

of the locking mechanism. There indicators, located on the inside door upholstery panel, are in the

CLOSED position when the locking rods are slid in the door frame, and in the OPEN position

when the rods are shifted out of the door frame. The front emerggency exit door is provided with

three (if installed) trminal switches which control the door warning circuit. The movement of one of

these switches is derived from the closed position of the door handle, while the remaining two

switches are actuated by the locking rods as they move into the door frame. Locking rod position

indicators (if installed) provide for visual check of the locking mechanism.

These indicators can be then on the inside door upholstery panel either in the CLOSED position

(when the locking rods are slid in the door frame) or in the OPEN position (when the rods are

shifted out of the door frame). The front baggage compartment door (both LH and RH side) is

provided with two terminal switches which control the door warning circuit. The switches are

actuated by springs which are compressed by the door lock hooks when the door is closed. The

passenger compartment emergency exit doors located between frames No. 13 and 14 (if

installed) are also provided with locking rod position indicators which provide for visual check of

the locking mechanism These indicators cam be scen on the inside door upholstery panel wither

in the CLOSED position when the locking rods are slid in the door frame, or in the OPEN position

when the rods are shifted out of the door frame.


M l T C R ( i f Inmtallod)

I

CARCdbOOR

(if inatalld)

TltR)(INAL MITCE ( i f i n s t a l l 4

FIG. 1-4 SCHEME OF THE MAIN DOOR WARNING SYSTEM

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E20 -

1.4. WINDOWS

The windshield of the pilot cockpit consists of two front windows, two side rear windows and two

outlook windows on both sides of the pilot cockpit located between the front and rear side

windows. The front windows are equipped by electric-heated glasses. The organic glass of

various thickness is used. The frame of the windows, placed in the place between the frames 4

and 8 is made from duralumin profiles joined mutually by spot welding and by riveting. The upper

and lower part of the frame is joined by columns made from duralumin sheet. The columns are

joined with both frames by means of riveting. Interconnection of the frame with the fuselage

construction is made by menas of riveted joints sealed by sealing cement U 30-M.

Passenger cabin windows

Outlook for passengers is ensured through the windows on both sides of the passenger cabin.

The windows are doubled. For outer glass the 5 mm thickness organic glass is used (6 windows

between the frames 9 and 12) and 3 mm thickness (the windows between the frames 12 and 18).

The inner glasses are made from the same material of the thickness 3 mm (between the frames 9

to 12) and thickness 2 mm (between the frames 12 to 18), see Fig. 1. The windows are protected

against moisture by their own atmosphere of dry air which is dried by means of heating in the case

that the heating system is switched on. The heated air passes below the upholstering. In the

centre and side panels a basic frame from duralumin sheet of the thickness 1 mm is made for

windows glassing, which is connected to the outside panel flange by means of spot welds. The

internal rim of the basic frame is connected in its bottom and upper part with stiffened horizontal

ribs by means of riveted joints. The ribs stiffen the wea kening of the outside skin due to the

window cut-outs. On the circuimference of the internal rim of the base frame 12 special nuts with

M4 thread are placed, which are predetermined for fastening of the pressure frame used for

window glass. The inner and outer windows are spaced by means of a welded frame onto which

both glasses are sealed by the FAKO-tape. The air expansion in the sealed space between

glasses (which occurs due to flight level changes) is facilitated by a hole of dia 3 mm, which is

placed in bottom part of the internal frame. The all assembly is supported from the outer side by

a rubber profile and from the inner side the assembly is pressed into the space of the stiffened

fuselage frame by 12 screws pressing onto the supporting frame.

WING

The wing is of a classic all-metal construction. It is double-beam, double-cavity, continuous

through the wing span. Torsion boxes form the cavities which are bounded by the skin of the

leading edge and the webs of the front beam and the skin of the part between the beams and

webs of the front and back beam. The center cavity is continuous through the whole wing span,

the leading cavity is ended at rib No. 3, which means that it is disrupted in the axis of symmetry.

The skin of the leading a nd center part is abundantly reinforced by longerons, so that it is capable

to carry a considerable part of the bending load.

The part in the back of the back beam is not supporting, it only has an aerodynamical function.

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, EZO

The engine bed is suspended on the wing on four suspenders, from which two are on the noses of

ribs No. 8 and 10, two are protruding from the bottom contour of the wing in the back of the front

beam also on ribs No. 8 and 10. On the front and back beam 3 suspenders of the win g tip fuel

tank are placed in the area of rib No. 31. The front torsion box is not disrupted in the area of the

engines. The fixed part of the engine nacelle is riveted on the angles, which are fastened from the

outside to the lower skin of the wing.

The cover of the leading edge of the wing between the engine nacelle and the fuselage is

equipped with 6 rectangular openings, which are determined for the disposal of harmful

substances during fire extinguishing of the engines and-for-thedisposal of used air from the- - - - - - - - - - - - - - - - - - - - - -

passenger cabin.

In the area between the front and back beam there are created areas for accommodating the

nubber fuel tanks. The lower covering in the area above the fuselage is sealed against penetration

of liquids and in the lowest point (in the axis of symmetry) there is a drain sump, which collects the

accumulated liquids, which are carried away under the fuselage by piping.

In the area in front of the front beam control rods for the aileron and wing flaps control, engi ne

control cables, the hydraulic system, the electrical installation, the air conditioning system piping

and the de-icing system are placed. The fuel system (with the exception of venting and

interconnection of the wing tip fuel tank with the external fuel tank in the wing, which is directed

through the central part of the wing) and part of the electrical installation are placed in the area in

the back of the rear beam. In the back of the rear beam the interceptor control is also placed.

~ - - - The venting of the fuel-system is led through-thecentral part above the fuel tanks-and exits from - - - - - - -

the wing on the lower cover before the rear beam between ribs No. 14 and 15.

The wing is connected with the fuselage by 4 hinges. It is equipped with all-metal double-slot wing

flaps with both slots changeable. The wing flap is in its depth divided into two mutually swivelling

connected parts, which are as a whole suspended on the wing. Both parts of the wing flap and the

wing are tied by a unique kinematic c oupling, which controls the deflection of the rear part of the

wing flap in relation to the deflection of the front part. Directly controlled is only this front part of the

wing flap (slot). Along the wing span the wing flap is divided into two parts - the internal wing flap

and the external wing flap.

The a~lerons are of fabric - metal combined construction. They have a metal elliptic leading edge

and a considerably backwards shifted axis of rotation. They are statically balanced to 100% by a

cou pled weight, which is placed in the nose of the aileron. They are connected to the wing by four

- - - suspenders, fiom which the suspender on rtb 24 is constfuckd for transfering otthe axial toad. -

The let? aileron is equipped with aileron trim tab. The interceptors are clamped on the hinge

between the I I th and 20th rib in the back of the rear beam of the wing.

The ABC tabs are between the 27th and 31st rib on the upper side of the wing, in the area of the

rear beam of the wing

The usual strength connections on the whole supporting system are carried out by riveting, the

exceptionally loaded centers are screwed together by fitted bolts.


With regard to aerodynamical requirements dictated by the used wing profile the surface of the

wing is quite smooth and identical with the theoretical contour, with the exception of bands on ribs

No. 3, 8, 10, 15, 31, which fulfil the function of flanges of these ribs and bands connecting the

covering on the lower side of the wing in the area of ribs No. 12 and 21, which are positioned

above on the covering. All riveting on the surface is carried out with countersunk rivets. Lentil

head rivets are used in ribs of the wing flaps and of the aileron and on the lower side of the wing

for riveting of the center ribs in the section between the fuselage and the engine nacelle.

FIG 1-5 SYSTEM DRAWING OF THE WING

(The numbers, in circles indicate the numbers of ribs)

A - System of the front beam B - System of the rear beam

C - System of the auxiliary beam D - Axis of rotation of the ABC tab

E - Axis of rotation of the aileron F - Axis of rotation of the wing flap

G - Leading edge of the aileron H - Leading edge of the wing flap


FIG. 1-6 SYSTEM DRAWING OF THE AILERON

(The numbers in circles indicate the numbers of ribs)

A - System of the main beam of the aileron

B - System of the auxiliary beam of the aileron

C - System of the beam of the trim tab


FIG. 1-7 SYSTEM DRAWING OF THE WING FLAP

(The numbers in circles indicate the numbers of ribs)

A - System of the beam of the slot

B - Leading edge of the wing flap

C - System of the beam of the wing flap

D - Trailing edge of the slot


FIG 1-8 SYSTEM DRAWING OF THE INTERCEPTOR

A - System of the rear beam of the wing

B - Axis of rotation of the interceptor

1.6. ENGINE NACELLES ~ ~

Engine nacelles serve as aerodynamic fairings for the engines and engine installations. They are

fixed directly on the wing and on the engine bed. The engine nacelle is composed of engine cowls

(in the front part) and of a fixed rear portion which is connected to the wing. As to increase safety

against fire propagation the engine nacelle is separated by a fire wall from the other parts of the

aeroplane

1.7. STABILIZERS

Horizontal stabilizer

They are located on the fuselage built-up structure. The stabilizers consist of the horizontal

stabilizer and elevator. The elevator consist of two halves mutually interlinked by the system of the

height controls. Each half is provided with a trimming tab.

Vertical stabilizer

They consist of the vertical stabilizer and rudder. The rudder is provided with a trimming tab.

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E2O

FIG. 1-9 ENGINE NACELLE

(1) Lower engine cowl, (2) Front upper cover, (3) Upper engine cowl, (4) Side

engine cowl, (5) Lower cover, (6) Rear fairing, (7) Side engine cowl, (8) Wing-to-

nacelle fillet, (9) Fixed portion of engine nacelle, (10) Locks of engine cowls, (1 1)

Locks of engine cowls, (12) Air intake for starter generator and alternator

cooling, (13) Cover, (14) Piano hinge, (15) (16) Cover (on the nacelle LH side

only), (17) (18) Guide pin, (19) Sealing, (20) Cover

A, B, C - designation of modifications of the cowl lever locks

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E. E9, E20

FIG.l-10 SYSTEM DRAWING OF THE HORIZONTAL STABILIZERS

A - Skin stiffener

B - Aircraft axis

C - System of the stabilizer front girder

D - System of the stabilizer rear girder

E - Axis of the elevator swivelling

F - System of the elevator main girder

G - System of the elevator auxiliary girder

H - Axis of the trim tab swivelling


FIG. 1-1 1 SYSTEM DRAWING OF THE VERTICAL STABILIZER

A - The system of vertical stabilizer front girder

B - The skin stiffener

C - The system of the vertical stabilizer rear girder

D - The system of bracing

E - The axis of the rudder swivelling

F - The axis of the trimming tab swivelling

G - The system of the rudder auxiliary girder

H - The system of the rudder main girder

1.8. TECHNICAL SUMMARY

MAIN DIMENSIONS

Length

Height (on the ground)

Landing gear spacing

Landing gear wheel base

Wing

Span:

- Configuration with tips

- Configuration with wing tip tanks

Wing area (without the tips)

Wing tip area

Mean aerodynamic chord

Aspect ratio

Taper ratio

Wing twist:

- aerodynamic

- geometric

Sweep back angle (at 250h centerline)

Angle of incidence at root

Dihedral angle

Ailerons

Span

Area

Aileron deflection:

14 424 rnrn

5 829 rnrn

3 650 rnrn

3 666 mm

19 479 rnrn

19 980 rnrn

34.860 rn2

2 x 0.16 rn2

1918 rnrn

10.45

2 x 3 822 rnrn

2 x 1.448 m2

UP 27O +I- lo

down 14O +I- 10

NOTE: Max. aileron deflections are marked by gauge lines on the airplane

Trim tab span (left aileron only) 1 030 rnrn

see fig. 1-1 3

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, E20 - - - - -

Trim tab area

T r~m tab deflect~on

Win flaps

Span

Area

Deflection of outer wing flap:

0.1966 m2

UP 220 +I- 20

down 22O +I- 2O

take-off

landing

Deflection differentiation of innerwing flap:take-off

landing

Ground spoilers

Span

Area

Deflection

ABC tabs

Span

Area

Deflection

Stabilizer and elevator

Span

Total area

Sweepback angle (at quarter-chord)

Dihedral angle

Taper ratio

Aspect ratio

Mean aerodynamic chord

Angle of stabilizer setting

2 x 4 830 mm

2 x 2.96 m2

I8O +I- lo

42' +I- lo

not specified

+ 100 +I- 3 0

Elevator

Area 2 x 1.575 m2

Max. deflection: UP 30° +/- lo

down 14O+ l o / - o 0

NOTE: Max. elevator deflections are marked by gauge lines on the airplane - see fig. 1-14

Elevator trim tab area 2 x 0.192 rn2

Elevator trim tab max. deflection: up l o 0 +I- lo

down 16O +I- lo

Fin and rudder

Height 3 310 mm

Total area 7.3 m2

Sweep angle (at quarter-chord) 35O

Aspect ratio 1.5

Taper ratio 0.5

Mean aerodynamic chord 2 285 mm

Rudder

Area 2.814 m2

Max. deflection (to both sides) 17O + 0°1- 30'

N0TE:Max. rudder deflections are marked by gauge lines on the airplane - see fig. 1-15

Rudder trim tab area 0.433 m2

Max. rudder trim tab deflection: 1 O0 + 0°/-1 (to the left side)

l o 0 + oO/-1°30' (to the right side)


FIG. 1-12 MAIN DIMENSIONS


Fig. 1-13 GAUGE LINES MARKING MAX. AILERON DEFLECTIONS

(1) Gauge lines in red colour marking max. aileron deflections

(2) Aileron control rod shroud varnished in white colour


FIG. 1-14 GAUGE LINES MARKING MAX. ELEVATOR DEFLECTIONS

(1) Gauge lines in red colour marking elevator max. deflections

(2) Edge of stabilizer lower skin varnished in white colour

FIG 1-15 GAUGE LINES MARKING MAX. RUDDER DEFLECTIONS

(1) Gauge lines in red colour to check max. rudder deflections

(2) Gauge line in white colour


CHAPTER 2

FLIGHT CONTROLS

2.1. General

2.2. Elevator control

2.3. Rudder control

2.4. Aileron control

2.5. Elevator trim tab control

2.6. Rudder trim tab control

2.7. Aileron trim tab control

2.8. Spoilers

2.9. Automatic bank control tabs

2.10. Flaps

2. I I . Autopilot servos




2.1. GENERAL

The flight control system is provided with dual rudder, elevator and aileron controls. Each pilot has

a control column of his own with a steering wheel in its upper part. The elevator is controlled by

pushing or pulling the control column. The ailerons are controlled by turning the steering wheel.

The rudder is controlled by means of two pairs of pedals which are also used to control wheel

brakes. The controls of the elevator trim tab, rudder trim tab and aileron trim tab are located on

the front control panel. The controls of wing flaps are located on the central and the RH control

panel. The controls of spoilers and automatic bank control-tabs are located on the central control

panel. The spoiler control push-buttons are located on the steering wheel. The tie rods of the

different control systems are marked by coloured strips as follows:

- aileron controls - 1 strip in black colour

- rudder control - 2 strips in black colour

- elevator control - 3 strips in black colour

- aileron trim tab control - 1 strips in brown colour

- rudder trim tab control - 2 strips in brown colour

- elevator trim tab control - 3 strips in brown colour

2.2. ELEVATOR CONTROL

The elevator is controlled from the control column by a system of pull-rods and levers. The column

parts are glued together with the ARALDIT AU I glue. The whole unit is suspended to the fuselage

structure in four suspension fittings by means of bolts. The lever is fixed by means of screws

between flanges of the left-hand and right-hand control column. The control motion is transferred

from the lever through the pull-rod to the bellcrank installed on the lower counter shaft and

through the p ull-rod to the bellcrank installed on the upper counter shaft.

The control motion is then transmitted by means of 9 pull-rods to the lever installed on the rear

counter shaft. From the rear shaft, the control motion is transmitted by means of levers and pull-

rods to the left-hand and right-hand elevator control levers. The two elevator halves are

kinematically joined by means of the rear counter-shaft. The pull-rods in the fuselage are guided

by levers.

The displacement range of the control colm n is limited by adjustable stop screws in the brackets

attached to the fuselage structure. The stop is a spring-type loader. The stop-contact area is

provided on the lever by means of which the motion from the control column is transfered to the

elevator control system.


FIG. 2-1 ELEVATOR CONTROL

I - Control column

II - Lower and upper control countershafts with levers

Ill, IV - Control levers in fuselage

V - Rear control counter shaft with levers


2.3. RUDDER CONTROL

The rudder control is performed by means of two separate foot control assemblies, interlinked

kinematically. Each foot control assembly is connected by two pull-rods to a bellcrank, interlinked

by means of pull-rods with a segment and the opposite bellcrank. The motion is transferred from

the segment by means of two front cables and two rear cables to the rear segment. From the rear

segment, the motion is transferred by means of the shaft with lever to the rudder. Between front

and rear cables, turnbuckles are situated. The cables are guided by a system of rollers on ball

bearings. Some rollers are placed on common countershafts, while the other ones on individual

brackets. The stops in the rudder control system are arranged so that the supporting surfaces on

bellcranks, contact the adjustable stop screw which is screwed into the bracket fixed to the

fuselage structure. The left-hand bellcrank serves in the capacity of the left stop, the right-hand

bellcrank in the capacity of the le ft one. The contact with the stops is achieved by applying a force

of 980 N + 49 N (100 kgs + 5 kgs) to pedals.

The foot control assembly consists of two pedal levers that swing on a common shaft, In the

upper part of pedal arms, there are sliding bearings for pedal tubes. Attached to the pedal tube

end by means of two tapered bolts is a lever by means of which the brake valves are controlled.

When swinging in the sliding bearing, the pedals are in function as brake pedals. The shaft is

fixed by means of tapered bolts in lateral fittings that are attached by means of screws to the

fuselage structure. The slots for the pedals are provided with brush screens in detachable covers

of the pedal control.

2.4. AILERON CONTROL

The aileron control is performed by means of the steering wheel which is placed on the control

column. The steering wheel is attached with screws to the shaft flange and to the adapter flange.

The adapter is seated in the column head on two ball bearings and secured with a nut and a

washer. Between the adapter flange and the column head of the control system, there is a spacer.

On the rear shaft part, there is a toothed wheel which is fixed by means of a nut and a washer.

The control motion is transmitted from the toothed wheel through the toothed wheel situated on

the shaft in a casing with bearings to the chain wheel. The chain wheel is fixed by means of a nut

and a washer.

The shaft is attached to the control column head by means of a nut and a washer.

On the chain wheel, there is a cable with a chain. The cable with the chain rolls over working

surfaces of individual pulleys and ends in segment, where the adjustment can be carried aut. too.

Mechanisms of ailerons control, situa ted on the left and right control column, are connected

mutually with a pull rod and levers, fixed to the shaft by means of two conical bolts. On the rear

shaft end, a lever is fixed by two conical bolts and the motion is transferred from this very shaft

end by means of pull rod to the bell crank.


FIG. 2-2 RUDDER CONTROL

I. Foot control assemblies with pull-rods, levers, bellcranks and rollers,

II. Rollers, turnbuckles and guides,

Ill. Rollers, segment and shaft


The rotating pull rod compensates for the influence of the inclination of the control column.

The control motion is further transmissed by a pull rod to the bell crank, situated on frame No. 7

and by means of the pull rod to the bell crank upon the upper countershaft and then, by means of

two pull rods with a different length to the bell crank which secures the transmission of the control

motion from the fuselage into the wing.

The direction of the pull rod motion in the fuselage is given by the guide bell crank.

The transmission of the control motion within the wing is performed by means of 5 kinematically

mutually combined levers to the bell crank and from this very bell crank, by means o f the pull rod

to the lever of the aileron driving mechanism. The direction of pull-rods within the wing is obtained

from the lever.

The kinematics is of a standard type. When turning the steering wheel to the right (clockwise), the

right aileron moves upwards and the left one downwards. When turning the steering wheel to the

left (counterclockwise), the right aileron moves downwards and the left one upwards.

Limiters of the ailerons control are such so that there is a lever on the shaft of the left control

column attached there limiting the control movement in such a way that it moves between two

adjusting screws, screwed in the risers of the control column case.

The limiting screws are adjustable.

2.5. ELEVATOR TRIM TAB CONTROL The elevator trim tab is controlled mechanically. The handwheels are attached to the shaft by

means of the pins and fixed in the brackets by means of the ball bearings. The brackets are

attached to the front control panel. The control movement is being transferred from the handwheel

by means of a system of cables running through the fuselage and resting on pulleys. The cable

system consists of cables and turnbuckles.

The control movement is being transferred finally from the cables to the drums.

The structure of the elevator trim tab drive is shown in. The elevator trim tab deflection is indicated

by the mechanical indicator situated next to the handwheel. The mechanical indicator of the

elevator trim tab deflection consists of a toothed disk which is driven from the shaft by means of

the pinion. A part of the indicator disk projects above the outline of the front control panel. The

zero-position is marked by a mark on the disk face.

When turning the handwheel in the PUSH direction, the elevator trim tab deflects upwards and

vice versa.

The movement of the elevator trim tab control is limited by the duralumin stops attached to the

cables and by means of the screws.

The stops area is made by the stop bracket that is fixed to the frame No. 25 by means of screws.

The adjustment of elevator trim tab control stops is performed by changing the position of stops

on the cable.

The movement of the elevator trim tab is kinematically independent, i.e. it does not depend on the

elevator movement. This means that with the movement of the elevator the trim tab remains in an

unchanged position which reference to the elevator.


FIG. 2-3 AILERON CONTROL

I - Control column

II - Attachment of bell crank

1 1 1 - Lower and upper counter shaft with bell cranks

IV - Control lever in fuselage

V - Counter shaft with bell cranks

VI - Control lever in wing

VII - Control lever in wing

FIG. 2-4 ELEVATOR TRIM TAB CONTROL

I - Handwheel forelevator trim tab control with mechanical indicator and pulleys

Il - Guide

1 1 1 - Pulley with turnbuckles

IV - Stops, pulleys and drum assemblies with pull-rods

A - Detail view showing the way of cable winding on the drum

a - 4 114 turns

b- Cable centre (clamp tube on the cable)


2.6. RUDDER TRlM TAB CONTROL

The rudder trim tab is controlled by means of the LEFTlRlGHT TURN change-over switch and the

UT 6D electromechanical strut.

Setting the lzver of the change-over switch to the left causes the left turn of the aircraft and the

rudder trim tab deflects to the right. Setting the lever of the change-over switch to the right causes

an opposite action.

The electro-mechanical strut is attached to the bracket formed by the nose ribs of the rudder. The

motion of the electro-mechanical strut is transferred by means of a bellcrank to the counter shaft,

placed vertically in the rudder. The turning of the bellcrank controls, by means of the counter

shaft, also the LUN 1688-8 trim tab position transmitter. The LUN 1688 trim tab position

transmitter together with the LUN 1687-8 trim tab position indicator serves for the indicator of the

rudder trim tab position. The rudder trim tab shall deflect from its neutral to its extreme position

within minimum 10 sec and maximum 18 sec (under the main vol tage of 28.5 +I- 1 V).

The control system of the rudder trim tab is actuated by the LEFTlRlGHT TURN change-over

switch with the TRlM TABS circuit breaker switched on. The LUN 1687-8 trim tab position

indicator together with the LUN 1688-8 trim tab position transmitter indicate the position of the

rudder trim tab.

2.7. AILERON TRlM TAB CONTROL

The aileron trim tab is located on the left-hand aileron. It is controlled by means of the BANK

LEFT-RIGHT change-over switch and the UT-6D electromechanical strut. Its motion is transferred

by a bell crank and a pull-rod to the aileron trim tab lever. The aileron trim tab control system is

actuated by the BANK LEFT-RIGHT change-over switch with the TRIM-TABS circuit breaker

switched on. The neutral position of the aileron trim tab is signalled by illumination of the signal

lamp.

2.8. SPOILERS

The spoiler control system is of electro-hydraulic-mechanical type

Electric part of spoiler control system

The electric part of the spoiler control system includes the following items:

The SPOILERS circuit breaker which activates and electrically protects the feeding circuit of

spoiler control. The VG-15K-2s switch which activates the spoiler control circuit. The KNR push-

buttons, actuated by a pressure cap, which activate the GA 184 Ul2 solenoid valve. The push-

buttons are protected against unintended depression by a safety latch. The extension of the

spoilers is signalled to the crew by the S POILERS signalling cell on CWD activated by the

LUN 3159.02-7 terminal switch.


BUS BAR

T' I r ' CENTRAL W A R N I N C ~ I DISPLAY

AIRFRAME

I._._.__._. I I

I TRIM TABS I

FIG. 2-5 RUDDER TRIM TAB CONTROL

FIG. 2-6 AILERON TAB CONTROL


Hydraulic part of spoiler control system

The hydraulic part of the spoiler control system includes the extensionlretraction circuit of the

spoilers which is connected by means of the GA 184 Ul2 solenoid valve to the constant pressure

circuit. The devices and the piping of this circuit are located in the middle part of the wing. The

GA 184 U12 solenoid valve is provided with four necks connected to pipings which connect the

solenoid valve with the constant pressure circuit, the return line and with the LUN 7138-8 spoiler

actuator (2 necks). A throttle on the solenoid valve inlet is inserted into the constant pressure

circuit to slow the spoiler extension.

Mechanical part of spoiler control system

The mechanical part of the spoiler control system consists of pull rods and levers which connect

the spoilers into one system. Linked to this mechanical system by means of an adjustable eyebolt

is the LUN 7138-8 spoiler actuator which ensures the extension and retraction of the spoilers. The

spoiler actuator is fastened to the bracket located on the rear wing spar. The lever controls the

LUN 3159.02-7 terminal switch through a pull rod located on the bracket. The stops in the spoiler

control system are arranged in such a way that the spoilers are provided with striking surfaces

which run against adjustable stop screws on the rear wing spar.

Spoiler control system operation

With the circuit breaker spoilers and the switch spoiler in ON positions and after remowing the

mechanical catch, the joint push-buttons on the left-hand steering wheel shall be depressed. The

GA 184 Ul2 solenoid valve opens the supply of hydraulic liquid from constant pressure circuit into

ground spoiler actuator LUN 7138-8. The spoilers are out while holding the push-buttons

depressed. Spoilers being out position is signalled by lighting the signal SPOILERS on the

signalling panel.


FIG. 2-7


2.9. AUTOMATIC BANK CONTROL TABS

The system of automatic bank control is electro-hydraulic- mechanical.

NOTE: At flight speeds above 11 1 kt (205 kmlhr) the extension of the automatic bank control tabs

is blocked.

Electric part of the ABC tab control system

The electric part of the ABC tab control system includes the following items:

The PROP. FEATHERINGIAUT. BANK CONTROL circuit breaker which switches and electrically

protects the feeding circuit of automatic bank control.

The AUT. BANK CONTROL switch which switches the control circuit of automatic bank control.

The TKE 52 PODG and TKE 54 PODG relays which activate the GA 184 U12 solenoid valve of the

corresponding automatic bank control tab and turn on or off the signalling circuits and the circuits

of limiters.

The extension of ABC tab is signalled by the AUT.BANK CONTROL signalling cell. To enable the

proper operation of the system during landing (i.e. when power control lever are in a position

below 89 +I- 1% n ~ ) the engine is equipped with 0.05K LUN 1492.01-8 pressure switches which

limit the operating range of this system.

For the system reliability each engine is equipped with two pressure switches.

Hydraulic part of the ABC tab control system

The hydraulic part of the ABC tab control system consists of the ABC tab extensionlretraction

circuit which is connected by the GA 184 U12 solenoid valves to the constant pressure circuit. The

components and piping of the ABC tab extensionlretraction circuit are located in the wing. The

GA 184 U12 solenoid valves are provided with four necks by two of which they are connected to

the constant pressure circuit and return circuit and by the remaining two the hydraulic fluid is led to

the LUN 7134-8 automatic bank control tab actuator.

Mechanical part of the ABC tab control system

The mechanical part of the control system of the automatic bank control tabs consists of rods and

levers.

Connected to this mechanical system are the LUN 71 34-8 actuators which supply the power

necessary for the extension and retraction of the automatic bank control tabs. The actuators are

fastened to brackets located on the wing ribs No. 28.

Operation of the ABC tab control system

The system is switched on by the PROP. FEAT.HERINGlAUT. BANK CONTROL circuit breakers

and the AUT. BANK CONTROL switch. If the power control levers of both engines are in the

position corresponding to 89 +I- 1% nG (92 +I- 1% n ~ ) and higher, the system beccomes ready to

operate within 5-7 sec. after switching on. This is signalled by the illumination of the green AUT.

BANK CONTROL signal in the AIRFRAME section of CWD.

AIRCRAFT TRAINING MANUAL L 410 UVP-E. E9. E20

The system is ready to operate up to the airspeed of 11 1 kt (205 kmlhr) only. At higher speeds the

system is switched off, which is signalled by the extinguishing of the AUT.BANK CONTROL

signal. The signal for automatic switching off proceeds from the LUN 11 73.11-8 airspeed

signallizer which at speeds lower than 11 1 kt (205 kmlhr) turns on the relay E 55, this activates

the relay E 145 which through its closed contacts, when the conditions for the automatic bank

control tab extension are fulfilled (switching on of the ist stage of the inoperative sensor M 309,

M 310), turns on the soleno id valve B 147, B 148.

At speeds above I I I kt (205 kmlhr) the relay E 145, which returns to rest position, activates

through its closed contacts 4 and 5 the relay E 142 which disconnects the power supply of

signalling and control. When the torque of one of the engines drops to approx. 24%, the

autofeather sensor M 309, M 310 transmits a signal to the relay E 143, E 144 which while closing:

- opens the feeding circuit thus blocking the extension of the automatic bank control tab on the

side of the inoperative engine. The automatic bank control tab remains retracted.

- extinguishes the green AUT. BANK CONTROL signal

- switches on the yellow AUT. BANK CONTROL signal

- activatos the solenoid valve E 147, E 148 which causes the extension of the automatic bank

control tab on the side of the operating engine.

When the airspeed of 11 1 kt (205 kmlhr) is reached, the ABC tab retracts automatically and the

yellow AUT. BANK CONTROL signal extinguishes.

If after the ABC tab extension accompanied by automatic propeller feathering the speed

decreases below 11 1 kt (205 krnlhr), the ABC tab does not extend again and the system does not

become operative. The removal of the operative condition of the automatic feathering and the

ABC systems is allowed only after successful re-starting of the inoperative engine or after a

check.

It is carried out by simultaneous switching off and on of both PROP. FEATHERINGIAUT. BANK

CONTROL circuit breakers and the AUTOFEATHER and AUT. BANK CONTROL switches.

If the air pressure after the compressor drops below the level limited by the pressure change-over

switch E 152, E 154, its contacts disconnect. The control circuit of the relay E 155, E 157 or

E 156, E 158 connects positive signal to the relay E 141 or E 142 of the system:

- opens the feeding circuit thus blocking the extension of the automatic bank control tab on the

side of the inoperative engine. The automatic bank control tab remains retracted.

- extinguishes the green AUT.BANK CONTROL signal

- switches on the yellow AUT.BANK CONTROL signal

- activates the solenoid valve E 147, E 148 which causes the extension of the automatic bank

control tab on the side of the operating engine.


2.10. FLAPS The aeroplane has an electro-hydraulic-mechanical of wing flap control. The wing flap controls are

located on the central and the right-hand control panel. The LUN 2690.05-8 wing flap position

indicator signals the flap position. The FLAPS 18' signalling cell cautions the pilot to extend the

flaps before the take-off.

NOTE: At airspeeds above 11 1 kt (205 kmlhr) , the possibility of the extension of wing flaps to the

maximum angle is blocked.

The electrical part of the wing flap control system consists of the electrical devices described

below. The WING FLAPS circuit breaker switches and electrically protects the feeding circuit of

the wing flap control. The D 701 terminal switch eliminates the possibility of wing flap extension

with an opened entrance door. The OK l(8)wing flap control unit, assembled on the base of a

443 853 067 722 three-position change-over switch provides the control of the GA 163 TI16

solenoid valve through the contacts of the KPK 3(8) wing flap terminal switch. When the wing flap

control unit is in its upper position, the flaps are retracted. The other two positions of the central

unit correspond to the 18O and 42O deflection of the wing flaps. In each of these positions the

control unit is fixed by an arrester.

The position of the flaps is signalled to the crew on the LUN 2690.05-8 wing flap position indicator

through the actuation of the D 701 terminal switches in the KPK 3(8) wing flap terminal switch

assembly which senses the wing flap motion by means of a rod driven by the wing flap actuator.

Functional check of the wing flap position indicator lamps is performed by actuating the 2 KNR

pushbutton labelled SIGN.

When actuating the SIGN, button (with the CENTRAL WARNING DISPLAY-AIRFRAME and

CENTRAL WARNING DISPLAY-ELECTRO circuit breakers ON), the lamps of the wing flap

position indicator are powered and by their light up their correct function can be checked. The

FLAPS 18' signalling cell informs the pilot about the necessity to extend the wing flaps, if the

landing gear is extended.

The hydraulic part of the wing flap control system consists of the wing flap extensionlretraction

circuit which is connected by means of the GA 163 TI16 solenoid valve to the constant pressure

circuit. The components and pipelines of this part of the wing flap control system are situated in

the middle part of the wing. The GA 163 TI16 solenoid valve has four necks: by means of two of

them the valve is connected to the constant pressure circuit and return circuit and through the

remaining two the hydraulic fluid is supplied via the LUN 7183.04-7 or LUN 7183.05-7 emergency

lock actuator and the 12 LUN 7573.4-7 throttle valve to the LUN 7231.02-8 wing flap actuator. The

LUN 7231.04-8 wing flap actuator contains,also the LUN 7543.02-8 double hydraulic lock and the

LUN 7547 03-8 shuttle valve. The LUN 7543.02-8 double hydraulic lock locks automatically the

spaces on both sides of the flap actuator piston as soon as the hydraulic fluid supply stops. The

shuttle valve secures the operation of the wing flap actuator with the main or emergency hydraulic

system.


CENTRKL WARNING DISPLAY PROP. FEATHERING AIRFRAME AUT. BANK CONTROL

PROP. FEATHERING AUTO 1

a@-

L/' AIRSPEED SlCNALlZER /SWITCIIES ABC TAB CONTROL OFF AT 205 km/h [AS/

FIG. 2-8 ABC TAB CONTROL


CONSTANT PRESSURE CIRCUIT


Through the LUN 7183.04-7 or LUN 7183.05-7 emergency lock actuator the hydraulic fluid is

supplied from the main hydraulic system via the GA 163 TI16 solenoid valve or from the

emergency hydraulic system via the LUN 6577-8 hand-operated valve.

If connected to the emergency hydraulic system, the hydraulic fluid procceds from the emergency

lock actuator directly into the return circuit.

The mechanical part of the wing flap control system consists of pull-rods and segments combining

all four flaps into one system. The pull-rod is led by the guide Connected to this mechanical

system is the wing flap actuator which develops the force which is necessary for the extension

a n d retraction of the wing flaps. The wingflap actuator is attached to two brackets in fronl ofthe-

main wing spar. The kinematics of wing flaps is based on the inter-connection of the sub-systems

on either side by a through piston rod of the wing flap actuator. The synchronization of wing flap

extension and retraction is secured by the adjustment of pull-rod lengths or segment rotation

angle.

Operation of the wing flap control system - normal operation mode

With the WlNG FLAPS circuit breaker switched ON, set the OK l(8) control unit from the initial to

the selected position. The signal cell of the initial position on the wing flap position indicator goes

off, the GA 163 TI16 solenoid valve connects the constant pressure circuit with the double

hydraulic lock so that the hydraulic fluid is supplied to one side of the wing flap actuator

whereas on the other side of the actuator it is being discharged through the solenoid valve into the

return - - - - - - - - - - - - - - - - - - - - - - - - - - - -

circuit. After the piston of the actuator reaches the selected position, the terminal switch

disconnects the corresponding electric circuit, which makes the piston stop. Simultaneously, the

corresponding signal cell on the wing flap position indicator illuminates. The stability of the wing

flap actuator piston position after the extension is secured by means of the double hydraulic lock.

If the wing flap control unit is set to the upper position, the corresponding signal cell signalling the

extended flap position extinguishes, the solenoid valve secures the connection of the second neck

of the actuator with the constant pressure circuit and of the first neck with the return circuit. The

double hydraulic lock secures the hydraulic fluid supply from one side of the wing flap actuator

only and the piston of the actuator returns to its initial position. As soon as it has reached this

position, the terminal switch breaks the feeding of the solenoid valve, the actuator piston stops

and, at the same time, the signal cell, signalling the retracted (initial) position of the flaps on the

wingf~ppo~sitionindicatorilluminates. - - - - - - - - - - - - - - - ~

Operation of the wing flap control system - emergency operation mode

The control of the emergency extension system of wing flaps is performed by means of the

LUN 6577-8 hand-operated valve labelled EMERG. EXTENSION-WING FLAPS (with the WlNG

FLAPS circuit breaker switched ON).

The handle of the hand-operated value is fixed in its upper position with a seal, which means that

the system of the emergency flap control is normally out of operation (closed).


hand-operated valve into the lower position. This opens the supply of the hydraulic fluid to the

emergency extension system.

Note: The LUN 7183 05-7 emergency lock actuator is cancelled from the 23.rd series.

BUS B A R

FIG. 2-9 WING FLAP CONTROL SYSTEM

CONSTANT PRESSURE CIRCUIT

-

I . -

1 -

SlhlPLlFlED WIRING 1)IAGRAM

WING FLAP

I

EXTENSION

EMERGENCY CIRCUIT I . - . - . - . . i

1 I WING FLAP ' ACTUATOR

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9, E20

The hydraulic fluid is supplied from the LUN 6100.03-8 hand-operated pump through the

LUN 6577-8 hand-operated valve into the LUN 7183.04-7 or LUN 7183.05-7 emergency lock

actuator. By the pressure of the hydraulic fluid the shuttle with the piston rod are displaced so that

the hydraulic fluid may flow further through the double hydraulic lock to the LUN 7231.02-8 wing

flap actuator.

At the same time the LUN 7183.04-7 or LUN 7183.05-7 emergency lock cylinder secures the

connection of the second neck of the actuator with the return circuit through the double hydraulic

lock. The signal cell, signalling the original position of flaps, extinguishes at this very moment, too,

and as soon as the extension cycle finishes, the signal cell, signalling the extended position

illuminates and the double hydraulic lock locks the actuator piston in this position.

2.1 1. AUTOPILOT SERVOES

Aileron servo

The aileron control by means of the autopilot servo is controlled by KSA 372 X aileron electrical

servo located underveath the cockpit floor between frames No. 6 and No. 7. The control

movement is transferred from the servo drum by cable to the lever on the segment. The cable

tension is adjustment can by carried out by tension nuts located on the segment. The motion is

transferred from the lever by a pull-rod to the lever by which the aileron primary control is

performed.

Rudder servo

The rudder control by means of the autopilot servo is controlled by KSA 372 electrical servo

located on the frame No. 24. The control motion is transferred from the servo drum by cable

through the pulleys to the rudder primary control cables. The cables are adjustment can be done

by tension nuts. The cables from the servo are fixed to the primary control cables by means

clamps.

Elevator servo

The elevator control by means of autopilot servo is controlled by KSA 372 elevator electrical servo

located at frame No. 26. The control motion is transferred from the servo drum by cable through

the pulley to the elevator primary control system. The cable tension is adjusted to 147 N (15 kp).

The adjustment can be done by tension nu's.

Elevator trim tab servo

The elevator trim tab control means of autopilot servo is controlled by KS 272 A elevator trim tab

electrical servo located on the frame No. 23. The control motion is transferred from the servo

chain wheel to the chain cable of the elevator trim tab primary control. The cable tension is

adjusted to 147 N (15 kp). The adjustment can be done by tension nuts.



AIRCRAFT TRAINING MANUAL L 41 0 UVP-E; ES, €20

CHAPTER 3 LANDING GEAR

3.1. General

3.2. Extension and retraction

3.3. Main wheel braking

3.4. Steering

3.5. Position and warning


lntentionaly left blank


3.1. GENERAL

The aircraft is equipped with a three-wheel retractable landing gear of swinging type. Nose landing

gear is installed in the fuselage front section between 4th and 5th frames and retracts in flight

direction. The main landing gear (left-hand and right-hand legs) is installed in landing gear

nacelles located on outer side walls of the fuselage center section between frames No.12 and

No.14 and is retractable inward perpendicularly to the flight direction, towards fuselage.

All the three landing gear legs are equipped with shock-absorbers. Retraction and extension are

performed by hydraulic jacks. The nose landing gear is equipped with a servo-control cylinder

which provides two functions:

- shimmy damper

- nose wheel steering servo

The rectracted and the extended position of the landing gear are signalized by light and

mechanical indicators. Apart from that and acoustical divice in installed in the cockpit warning the

pilot about the necessity of the extension of the landing gear.

Main landing gear wheels are equipped with hydraulic brakes.

The main landing gear is equipped with a K38-1100-7 wheel and a K38-1200-7 brake and a

tubeless electrically conductive BARUM 12.50-1 0, model 4 tyre. The nose landing gear is

equipped with a K39-1100-7 wheel and a tubeless electrically conductive BARUM 9.00-6, model 4

tyre. Brakes are controlled through a hydraulic system which generates pressure in brake

cylinders on the main landing gear wheel brake namely:

- during normal braking with brake pedals: 0 - 4.4 + 0.3 MPa (0 - 45 + 3 kp/cm2)

The pressure of 2.45 + 0.49 MPa (25 + 5 kp/cm2) is used for parking the aircraft on apron.

For engine test the pressure of 4.9 + 0.49 MPa (50 + 5 kp/cm2) is used. Parking brakes can be

used also for emergency and for normal braking during landing by generating the pressure of

0 - 4.4 + 0.3 MPa (0 - 45 + 3 kp/cm2) with smooth movement of the hand-operated pump lever.

At that case both wheels are braked uniformly.

CAUTION: AT THIS CASE BRAKES CANNOT BE USED TO CHANGE THE AIRCRAFT

MOVING DIRECTION, THE ANTIBLOCKING DEVICE IS INOPERATIVE, AND THE

WHEELS CAN GET BLOCKED IF THE PRESSURE OF 45+3 kp/cm2.

The aircraft is equipped with an antiblocking device which.switches the wheel brakes off closely

before their potential blocking. Retracted and extended position of the landing gear are signalized

by light and mechanical indicators. Apart from that an optical and acoustical device is installed in

the cockpit warning the pilot about the necessity of the extension of the landing gear.

FIG. 3-1 MAIN LANDING GEAR

(1) Carrying tube, (2) Sealing, (3) Suspension, (4) Swing arm, (5) Shock

absorber, (6) Wheel axle, (7) Wheel with brake, (8) Tyre, (9) Tilting door,

(10) Main landing gear hydraulic jack, (1 1) Hinge, (12) Bracket, (13) Tie rod, (?cj

Swing lever, (1 5) Fixed lever, (16) Tie rod, (7) Main landing gear mechanical

lock, (18) Bushing, (19) Support, (20) Pull rod, (21) Bracket,

(22) landing gear position mechanical indicator, (23) Inertial sensor set for the

L.H.part of UA 28A-13 or UA27A-13, for t he R.H.part of UA 28A-14 or UA 27A-

14, (24) Fastening strip with button, (25) Bolt with nut and washer, (26) Bonding

strip, screw.

FIG. 3-2 NOSE LANDING GEAR

(1) Shock absorber, (2) Servo control of the nose gear wheel, (3) Swinging lever,

(4) Nose landing gear door LH + RH, (5) Rear nose landing gear door,

(6) Wheel axis, (7) Wheel with tyre, (8) Countershaft, (9) Nose landing gear jack

(10) Safety pin, (1 1) Suspension point, (12) Bracket, (13) Suspension, (14) Tie

rod, (15) Suspension, (16) Tie rod, (17) Tie rod, (18) Flange,

(19) LUN 3159.01-7 terminal switch, (20) Nose landing gear position mechanical

indicator, (21) Nose landing gear mechani cal lock, (22) Hinge for the landing

ski limiter.

NOTE: Nose landing gear doors are shown in the open position which can be

achieved by disconnecting the pull rod (16) after having removed the

safety pin (10).


3.2. EXTENSION AND RETRACTION The control system of the extension and retraction of the landing gear is electro-hydraulic one.

The electrical part of the system controls the GA 163Tl16 electromagnetic valve, while the

hydraulic part of the system controls landing gear jacks and lock release cylinders.

Description

The electrical part of the landing gear extension and retraction control system consists of the

following: LANDING GEAR circuit breaker which interrupts and protects the circuit of'the landing

gear extension and retraction control.

Landing gear OP(8) change-over switch on the basic of the 2 PPG-15K-2s change-over switch

with two positions, which controls the GA 163 TI1 6 solenoid valve by means of terminal switches

in mechanical locks and by means of terminal switches on landing gear jacks. If the landing gear

change-over switch is in its DOWN position, the landing gear is lowered. The landing gear

retraction comes about when the change-over switch is set to the UP position, In both positions,

the change-over switch is secured by means of a stop. When pilots leave the cockpit, an

additional fixing of the landing gear change-over switch is performed in the position LOWERED by

means of a safety pin. B 073 571 N delay circuit improves the landing gear retraction.

The hydraulic part of the landing gear extension and retraction control system consists of a

landing gear extension and retraction circuit which is connected through the GA 163 TI16 solenoid

valve to the permanent pressure circuit. GA 163 TI16 solenoid valve has 4 necks. By two of them,

it is connected to the permanent pressure circuit and the reverse circuit and through the other two

necks the hydraulic fluid is supplied under pressure through LUN 7515.10-7 , LUN 7515.1 1-7,

LUN 751 5.12-7 landing gear emergency valves t o the LUN 7108.1 1-7, LUN 7108.1 2-7 LH and

RH main landing gear jacks and to LUN 7233.04-7 the nose landing gear jack. Before entering the

LUN 7233.04-7 working cylinder of the nose landing gear jack, the hydraulic liquid flow is throttled

by the B 057 840 N throttling valve which prevents eventual damaging of the nose landing gear

due to hydraulic shocks at its rapid extension and retraction. By means of the LUN 7183.04-7 or

LUN 7183.05-7 emergency lock actuator the hydraulic liquid is supplied und er pressure from the

main circuit via GA 163 TI1 6 solenoid valve or from the emergency circuit via LUN 6577-8 hand-

operated valve. The LUN 7183.04-7 or LUN 7183.05-7 emergency lock actuator interconnects the

landing gear control circuit with the reverse circuit. The nose landing gear leg is secured in the

lowered position by the nose landing gear mechanical lock. In the retracted position the nose

landing gear is secured by means of a ball lock which is a part of the nose landing gear jack. The

main landing gear is secured in its extended position by locks with segments which are

component part of the main landing gear jacks. In the retracted position, the main landing gear is

held by the main landing gear mechanical locks.

LUN 7515.10-7, LUN 7515.1 1-7 , LUN 7515.12-7 emergency valves form the component parts of

individual landing gear jacks. The emergency valves provide for landing gear extension, retraction,

and emergency extension, as well as hydraulic locking of landing gear jack pistons in extended

position. LUN 7185-7 lock release cylinder is the component part of the nose landing gear

mechanical lock, LUN 7188-8 lock release cylinders are component parts of main landing gear

mechanical locks.

Lock release cylinders serve for releasing these mechanical locks. LUN 7561 -7 stabilizing valve

protects LUN 71 88-8 mechanical locks release cylinders against pressure peaks, which arise at

the end of the landing gear retraction and which could cause the repeated release.

FIG. 3-3 DIAGRAM OF THE LANDING GEAR EXTENSION AND RETRACTION

CONTROL SYSTEM

Legend to the fig. - see the following page


Legend to the fig. 3-3:

(1) GA 163 TI16 solenoid valve, (2) LUN 7183.04-7 or LUN 7183.05-7

emergency lock actuator, (4) LUN 7108.1 1-7 main landing gear jack, left-hand,

(5) LUN 7233.04-7 nose landing gear jack, (6) LUN 7108.12-7 main landing

gear jack right-hand, (7) LUN 7185-7 lock release cylinder, (8) LUN 7188-7 lock

release cylinder, (9) LUN 7515.10-7 landing gear emergency valve, -

component part of LUN 7108.1 1-7 the main landing gear jack, (10) LUN

7515.1 1-7 landing gear emergency - component part of LUN 7108.12-7 the

jack, (1 1) LUN 7515.12-7 landing gear emergency valve - component part of

LUN 7233.04-7 the nose landing gear jack, (12) I3 057 120 N discharge valve,

(1 3) LUN 7561 -7 stabilizing valve.

a - permanent pressure circuit

b - return circuit

c - emergency landing gear extension circuit (see the fig. 3-4)

Legend to the fig. 3-4:

(1) LUN 6577-8 hand-operated valve, (2) LUN 7183.04-7 or LUN 7183.05-7

emergency lock actuator, (4) LUN 7108.1 1-7 LH main landing gear jack,

(5) LUN 7233.04-7 nose landing gear jack, (6) LUN 7108.12-7 RH main landing

gear jack, (7) LUN 7185-7 lock release cylinder, (8) LUN 7188-7 lock release

cylinder, (9) LUN 7515.10-7 landing gear emergency valve - component part of

LUN 71 08.1 1-7 the main landing gear, (10) LUN 751 5.11-7 landing gear

emergency valve - comp onent part of LUN 71 08.12-7 the main landing gear

jack, (1 1) LUN 751 5.12-7 landing gear emergency valve - component part of

LUN 7233.04-7 the nose landing gear jack, (12) I3 057 120 N discharge valve.

a - emergency control circuit

b - return circuit

c - landing gear extension and retraction circuit (see the fig. 3-3)


FIG. 3-4 EMERGENCY LANDING GEAR EXTENSION SYSTEM DIAGRAM

Legend to the fig. - see the previous page


Operation - normal control

Retraction of the landing gear

With the circuit breaker LANDING GEAR on, set the handle of 2 PPG-15K-2s landing gear

selektor to the up position. GA 163 TI16 solenoid valve connects the permanent pressure circuit

with the retraction circuit. The hydraulic liquid under pressure goes through the LUN 7183.04-7 or

LUN 7183.05-7 emergency lock actuator which allows the hydraulic liquid to flow under pistons of

individual landing gear cylinders and simultaneously, the other side of pistons is connected with

the return circuit. Connected to the tube leading to the LUN 7233.04-7 nose landing gear jack is - -

- - - - - - - - - - - - - - - - - - - -

- - - - -

LON 7185-7ralease cylinder, which releases the nose landing gear mechanical lock and terminal

switch switches off the middle green lamp on the landing gear position indicator and switches on

the red signalling that the nose landing gear is in interposition. The hydraulic liquid under pressure

flows to LUN 7515.12-7 emergency valve of the nose landing gear jack (which hydraulically locks

the extended pis ton position of the nose landing gear jack) and due to its repositioning the nose

landing gear jack is connected with the return circut.

The piston in the nose landing gear jack rnaves in its other end position, in which it is locked by

the mechanical ball lock placed inside the jack. The terminal switch which is connected with this

ball lock, signals by switching on the central red lamp that the landing gear is locked in its

retracted position. The circuit of LUN 71 08.1 1-7 and LUN 7108.1 2-7 main I anding gear jacks

operatcs in similar way. It is connected in parallel to the circuit of the nose landing gear jack. The

main landing gear jacks are locked in their retracted position by means of mechanical locks. - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -

Terminal switches signalize this extended position with green lamps on.

After the delivery of the hydraulic fluid under pressure, the segment locks of main landing gear

jacks release and simultaneously, by changing over the terminal switch on main landing gear

jacks the landing gear interposition is signalled by means of red lamps. LUN 751 5.1 0-7 and

LUN 751 5.1 1-7 emergency landing gear valves are parallelly connected to this circuit (they lock

hydraulically the piston in the extended position) and due to their shifting, the main landing gear

jacks are connected with the return branch. After the piston reaches its limit position, the landing

gear is locked by main landing gear mechanical locks and terminal switches on these locks signal

by switching off the red lamps that the landing gear is locked in its retracted position.

Extension of the landing gear

With the circuit breaker LANDING GEAR on, move the handle of the landing gear switch in its - - - - - - - - - - - - - - - - -

lower position. The GA 163 ~/16solenoid valve connects the permanent pressure circuit with the

working circuit. The hydraulic fluid under pressure flows through LUN 7561.7 stabilizing valve and

LUN 7515.12-7 landing gear emergency valve into the nose landing gear jack where it releases

the mechanical ball lock of the jack and the terminal switch on the nose landing gear jack signals

by red signal light on the landing gear position indicator that both the piston and the nose landing

gear leg are in the interposition. After shifting the piston in its lower limit position, the mechanical

lock of the nose landing gear locks the nose landing gear. The terminal change-over switch on the


nose landing gear mechanical lock signals by switching off the red signal lamp and switching on

the green signal lamp that the nose landing gear leg is locked in the extended position. The

hydraulic fluid under pressure proceeds towards LUN 7108.1 1-7 and LUN 7108.1-7 main landing

gear jacks via downlock cylinders and LUN 751 5.1 1-7 and LUN 751 5.1 0-7 emergency landing

gear valves and moves the piston in the lower limit position. LUN 7188-7 lock release cylinders

which form parts of main landing gear machanical locks and release the landing gear legs with

terminal switches signalize by switching on red signal lamps that both main legs of the landing

gear are in the interposition. After the pistons reach the lower limit position, segment locks of

these cylinders mechanically lock the piston and simultaneously change over terminal switches

signal by switching off, the red signal lamps and switching on the green signal lamps that both

legs of the main landing gear are locked in the extended position.

Operation - emergency extension

The circuit of the emergency landing gear extension is controlled by LUN 6577-8 hand-operated

valve, marked as EXTENSION GEAR. The lever of the hand-operated valve is fixed in its upper

position by a seal which means that the circuit of the emergency landing gear extension is closed.

In the case of an emergency landing gear extension, the lover of LUN 6577-8 hand-operated

valve has to be shifted in down position. In this way, the supply of the hydraulic fluid into the

emergency landing gear extension circuit is opened. By means of LUN 6100.03-8 emergency

hydraulic pump, the hydraulic liquid is pressed through the LUN 6577-8 hand-operated valve to

LUN 7183.04-7 or LUN 7183.05-7 emergency lock actuator. The pressed hydraulic fluid shifts the

shuttle with the piston rod and comes into the emergency circuit to LUN 7188-7 lock release

cylinder LUN 751 5.1 0-7 and LUN 751 5.1 1-7, LUN 751 5.1 2-7 landing gear emergency valves and

out from here, into landing gear jacks. Simultaneously the shuttle of LUN 71 83.04-7 or

LUN 71 83 .05-7 emergency lock actuator connects the discharge of jacks directly with the return

circuit. Lock release cylinders release mechanical locks of the main md ing gear which starts the

signallig of the main landing gear extension and by releasing the mechanical lock of the nose

landing gear jack, the signalling of the nose landing gear extension starts.


LAND. GEAR

EMERGENCY VALVE

TERMINAL SWITCH f - d r t t

EMERGENCY

M TERMINAL SWITCH

TERMINAL SWITCH

EMERGENCY VALVE "t

LAND. GEAR

DOWN

FIG. 3-5 LANDING GEAR CONTROL


FEEDING C I R C U I T

--- RETURN C I R C U I T T

I I

u P

SOLENOID VALVE

- - - EMERGENCY EXTENS ION

LAND. GEAR

EMERGENCY HYDRAUL. PUMP

EMERGENCY C I R C U I T I - 0 - * * * f a * -


3.3. MAIN WHEEL BRAKING Normal main wheel braking with brake pedals

Description

The main wheels brakes and parking of the aircraft are controlled with a hydraulic circuit of main

wheels braking. The braking circuit is separated from the steady pressure circuit with a

674 600 B(8) non-return valve the LUN 6953.02-8 separate hydraulic accumulator with which

supplies the pressure energy in case of the steady pressure circuit failure from. The capacity of

the hydraulic accumulator provides et full pressure at least 25 fully braking with detached

antiblocking device. The hydraulic fluid is supplied through a LUN 7514.02-8 hand operated valve

to four

LUN 7367.03-8 brake valves that generate the pressure for braking. They are controlled by a lever

gear from the foot operated steering pedals. The brake valves deliver the hydraulic fluid with

smooth reduction from 0 to 4.4 + 0.3 MPa (0 to 45 + 3 kp/cm2) to the brakes proportionaly to

intensity of pressing the foot operated steering pedals. The LUN 751 4.02-8 reduction valve keeps

the steady reduced pressure of 4.9 + 0.5 MPa (50 + 5 kp/cm2) in the main wheel circuit.

CAUTION: WHEN REPLACING THE LUN 7367.03-8 BRAKE VALVE IT IS NECESSARY TO

INSPECT AT FIRST THE IDENTICAL MARKING ON ALL FOUR BRAKE VALVES ON

BOTH STEERING BLOCKS ,NAMELY LUN 7367.03-8.

The braking of the main wheel is doubled. To prevent the simultaneous braking from both pilots

two LUN 7368.01-8 shuttle valves are installed to the circuit. They supply the hydraulic fluid from

valves to brake cylinders. SP 1 self-sealing couplings (8) are determined to prevent the loss of

hydraulic fluid during brake disconnection. The main wheel brake system has an antiblocking

device consisting of UA 27A-13 or UA 28A-13 (or UA 27A-14 or UA 28A-14) inertial sensors and

LUN 2575-7 or LUN 2575.01 -7 electromagnetic distributors.

FIG. 3-6 DIAGRAM OF MAIN LANDING GEAR WHEEL BRAKING SYSTEM

(1) LUN 7514.02-8 Reduction valve, (2) LUN 7367.03-8 Brake valve,

(3) LUN 7547.03-7 Shuttle valve, (4) SP 1 (7) Self-sealing coupling,

(5) LUN 1446.02-8 Dual pressure gauge, (6) LUN 7368.01-8 Shuttle valve,

(7) Brake cylinders, (8) LUN 2575-7 or LUN 257'5.01-7 Electromagnetic

distributor.

a - steady pressure circuit

b - reverse circuit

c - brake drain

d - into emergency tank

e - parking brake circuit (see fig. 3-7)


Operation

The hydraulic fluid under the pressure of 14.7 MPa (1 50 kp/cm2) is reduced by the

LUN 7514.02-8 reduction valve to the steady pressure of 4.9 + 0.5 MPa (50 + 5 kp/cm2). If the

pressure in the reduction valve exceeds 5.9 +/- 0.19 MPa (60 +/- 2 kp/cm2) the safety valve slide

releases the excessive fluid into the reverse circuit. The hydraulic fluid under the pressure of 4.9

MPa (50 kp/cm2) is supplied to four LUN 7367.03-8 brake valves. When braking the aircraft by

smooth pressing of brake pedal the force is transmitted by a regulation spring to the valve. The

valve connects the pressure fluid intake with the brake circuit. After equalization of the pressure in

the brake circuit and behind the pressed regulation spring, the valve closes further hydraulic fluid

supply. After releasing the brake pedal the valve connects the brake circuit with the reverse

branch and releases the brake. Higher or lower pressure in brakes can be obtained by changing

the intensity of brake pedal pres sing. The control force depends on the lift and changes smoothly

within the range of 169 N through 490 N (20 kp through 50 kp). The hydraulic fluid is delivered

from the brake valves to LUN 7368.01 -8 shuttle valves. The slide controlled by the regulation

spring opens and closes the hydraulic fluid supply from the first or the second pilot respectively.

That prevents simultaneous braking from both pilots.

The hydraulic fluid is supplied from shuttle valves via the LUN 2575-7 or LUN 2575.01 -7 elect

romagnetic dis tributor to the LUN 7547.03-7 shuttle valve. Through this valve flows the hydraulic

fluid freely to brake cylinders. In case of normal circuit failure it opens the way for the fluid from

the parking brake circuit.

SP 1 (7) self-sealing coupling is placed between the shuttle valve and the brake cylinders. In

screwed state the sealing cone presses away the connection valve and the fluid passes freely.

During dismantling the brakes the connection valve is pressed back by the spring and closes the

flui d supply. LUN 1446.02-8 dual pressure gauge is connected parelelly to the left-hand and right-

hand brake circuit indicating the brake control pressure. The excessive fluid from the

LUN 7367.03-8 brake valves is drained either directly to the main hydraulic tank connected to

which is the emergency tank.

The LUN 2575-7 or LUN 2575.01 -7 electromagnetic distributor which is controlled by an inertial

sensor, switches the fluid supply from the brake cylinders to the reverse branch in case of braking

a nd simultaneous wheel blocking. The electrical circuit of this antiblocking device is switched on

permanently, but the pilot can decide about its switching off using the manual switch ANTISKID

placed on the central control panel.

Emergency braking of the main landing gear wheel

Parking brake circuit can be used for emergency braking the main wheels. At that case both main

wheels are braked simultaneously. As higher pressure can be supplied to brake cylinders via the

parking brake circuit than under normal braking it is necessary to watch the parking brake

pressure gauge (on instrument panel) not to exceed the pressure of 4.41 + 0.29 MPa

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, EZO ~-p-~~

(45 + 3 kplcm2) and to prevent the brake blocking. 4 to 5 cycles of the hand operated hydraulic

pump are required for emergency braking. During emergency braking, the antiblocking device is

out of operation.

FIG. 3-7 DIAGRAM OF PARKING BRAKE CIRCUIT

(1) LUN 6578.02-8 Hand operated valve, (2) LUN 7557.01-8 Safety valve,

(3) LUN 6900-8 Brake accumulator, (4) LUN 7547.03-7 Shuttle valve,

(5) SP 1 (7) Self-sealing coupling, (6) MA 100 Pressure gauge, (7) Brake

cylinders, (8) Non-return valve 674 500 B (8).

a - emergency control circuit

b - main wheel braking circuit (see fig. 3-6)

c - reverse circuit


FIG. 3-8 WHEEL BRAKING SYSTEM


Parking braking of the main landing gear wheel

Description

Parking brake circuit serves to the permanent braking of aircraft on the ground and its braking

through engine test.

LUN 6100.03-8 hand operated hydraulic pump is a source of pressure for the parking brake

circuit. Hydr~ulic fluid is pumped from an emergency tank. The LUN 6578.02-8 hand operated

valve serves to control the proper parking. It is separated from the emergency circuit by a 674 500

B non-return valve (8). The hydraulic fluid is supplied from the hand operated valve to the

LUN 7547.03-7 shuttle valves and to brake cylinders. The LUN 6900-8 parking brake accumulator

and the LUN 7557.01 -8 safety valve are parallel connected to the circuit. The pressure in the

parking brake circuit is checked by MA 100 pressure gauge.

Operation

For parking the LUN 6578.02-8 hand operated valve is switched to STOP position (upwards). By

pumping with the LUN 6100.03-8 hand operated hydraulic pump the hydraulic fluid is supplied

through the 674 500 B non-return valve (8) and the LUN 6578.02-8 hand operated valve to the

LUN 7547.03-7 shuttle valves which, by moving the slide, let the hydraulic fluid flow to brake

cylinders. Pump by the hand operated hydraulic pump until the required pressure for parking is

achieved. The pressure is ch ecked by the MA 100 pressure gauge. The parking brake circuit is

protected against overloading by the LUN 7557.01 -8 safety valve which is set to the relief

pressure of 5.88 MPa (60 kp/cm2). The LUN 6900-8 brake accumulator prevents pressure drop in

the parking brake circuit caused by the compensation of temperature and leakage of the circuit.

Parking brake release is performed by over switching the LUN 6578.02-8 hand operated valve to

the initial position (downwards). Thus the parking circuit, whi ch is under pressure, is connected

with the reverse branch, excessive hydraulic fluid is drained into the ,everse branch and parking

circuit remains without pressure.

In the parking brake circuit there is a 0.7 S LUN 1492.04-8 pressure switch for blocking of the

voice recorder erasing circuit after the aircraft has stopped (if installed).

3.4. STEERING Description

Steering (i. e. nose wheel leg steering) is an electro-hydraulic-mechanical system and can be

hand-operated by a lever on the steering column or foot-operated by pedals of the foot-operated

steering. Manual nose wheel steering anables taxiing of the aircraft. The foot-operated steering

controls the nose wheel within small angles and is used for take-off and landing only.

The electrical part of the system controls the GA 184 U/2 solenoid valve. The foot-operated nose

wheel steering controls also the LUN 2550.02-8 electromagnetic clutch.

Hydraulic part of the system supplies hydraulic fluid into the nose wheel steering servo cylinder.


Mechanical part of the system controls the nose wheel steering servo which is a part of the nose

landing gear leg.

The steering servo has two independent functions in the aircraft:

- when connected into the hydraulic system through the switch located on the central control panel

(signalling lamps in the PEDAL STEERING, MANUAL STEERING window on the signalling

panel are on it works as a steering servo.

- when not connected into the hydraulic system, i.e. the switch on the control panel is off

(signalling - - lamps - in the PEDAL STEERING, MANUAL STEERING window are off) it works as a

shimmy damper. The steering servo must be switched off when towing the aircraft by a tractor.

Operation of the nose wheel steering system when hand-operated

With LANDING GEAR circuit breaker on and NOSE WHEEL STEERING switch in MANUAL

position the GA 184 U12 solenoid valve lets the pressurized hydraulic fluid flow into the nose wheel

steering servo circuit. Simultaneously the MANUAL STEERING lamp on signalling panel lights up.

By moving the distribution slide valve in steering servo cylinder through the lever on the steering

wheel the pressure fluid is delivered to one side of the nose steering servo piston. The other side

is at the same time connected to the waste branch. Thus the wheel starts turning to one side. The

wheel turning is not proportional to the shifting of the lever on the steering wheel. Centering

springs in the slide valve body return the slide valve automatically into the central position,

meaning that the lever must be held-duringthe-whole period of turning in displaced position. After - - - - - - - - - -

positioning the NOSE WHEEL STEERING switch into neutral position the GA 184 Ul2 solenoid

valve connects the supply inlet with the waste branch and the signalling lamp for MANUAL

STEERING on the signalling panel lights down. Manual steering controls the nose wheel within

50°to each side with a tolerance of -5'

Operation of the nose wheel steering system when foot-operated

With LANDING GEAR circuit breaker on and NOSE WHEEL STEERING switch in PEDAL

position and through motion of rudder control pedals the neutral positions of the nose wheel and

the rudder become identical. At that moment the D 701 terminal switch switches on arid closes

the LUN 2550.02-8 electromagnetic clutch. The GA 184 Ul2 solenoid valve lets the pressure

hydraulic fluid flow into the nose wheel steering servo circuit.

Simultaneously - - - - - - - the PEDAL STEERING signalling lamp lights up~on the signalling panel. By

moving the slide valve in the servo steering cylinder with foot-operated steering pedals the

pressure fluid is delivered to one side of the piston of the steering servo. The other side is

simultaneously connected to the waste branch. Wheel turning is proportional to the motion of the

foot-operated steering pedals. It enables the nose wheel to turn within the range of 4O30' +I- 1°30'

to the right (and to the left) the nose wheel deflection of + 4'30' corresponds to full rudder

deflection.


Centering springs in the slide valve body return the slide valve automatically into the central

position. After moving the NOSE WHEEL STEERING switch into neutral position the GA 184 Ul2

solenoid valve connects the supply inlet with the waste branch and the PEDAL STEERING

signalling lamp lights down on the signalling panel. The LUN 2550.02-8 blocking electromagnetic

clutch is automatically opened following the impulse during the landing gear retraction. In case the

clutch remains closed even after the landing gear retraction the flight safety is not endangered.

FIG. 3-9 DIAGRAM OF THE NOSE WHEEL STEERING SYSTEM

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, E20 -- -

3.5. POSITION AND WARNING

The arcraft IS provided with the

- light

- acoustic

- mechanical

indication of the landing gear position

Landing gear position light indication - - - - -

Light indication of the landing gear position can be indicated by the LUN 1694-8 landing gear

position indicator connected to the LUN 3170-7 terminal microswitches located on LUN 7108.1 1,

LUN 7108.12 main landing gear jacks and to KP 517lU terminal switch located on the nose landing

gear mechanical lock. Extended and secured position of each landing gear leg is indicated on the

landing gear position indicator by lighting up of 3 green lights. Intermediate position of the landing

gear legs is indicated on the indicator by lighting up of 3 red lights. Retracted and secured

position of the landing gear is not indicated by light. There is no light signalization of the retracted

landing gear position on the landing gear position indicator.

Landing gear position mechanical indicator

Mechanical indicator of landing gear position indicates the extended position of landing gear by

three extended mechanical indicators. In retracted-position the indicatars are retracted in. - -

- - - - - - - - - - -

Landing gear position acoustic signalization

Acoustic signalization of the landing gear position is indicated by the H 118 horn from AM 800 K

terminal switches The acoustic signalization of landing gear operates:

-when both POWER control levers are in IDLE position, the speed is lower than 205 kmlh. IAS

and the landing gear is retracted or in intermediate position,

-when wing flaps slide out to landing position and the landing gear is in intermediate position.

The acoustic signal by a horn is always accompanied with lighting of the EXTEND LANDING

GEAR block on the warning display. The electrical circuit of the acoustic signalization is protected

by LANDING GEAR ACOUSTIC SIGNALIZATION fuse.

Afterextendmgand locking the landing gear (sliding i~ the wing-flaps is not sufficient) orafter -

resetting one of the two POWER engine control levers forward, the signalization of necessity to

extend landing gear stops the operation.

NOTE: The ZO-1 S retarding circuit prevents the wrong signalization within the landing gear

extend circuitry with wing flaps slided out to 18O.

AIRCRAFT TRAINING MANUAL L 410 UVP-E E9, E20

FIG. 3-10 DIAGRAM OF THE SYSTEM OF THE LANDING GEAR POSITION

INDICATION


- - - - - -


CHAPTER 4

AIR - CONDITIONING

4.1.Generalp

4.2. Distribution of air

4.3. Heating

4.4. Temperature control

4.5. Indication




4.1. GENERAL The air-conditioning system is intended for:

- the heating of the pilot and the passenger cabin within the limits satisfying physiological demand

for heat,

- the heating of the storage battery bay,

- sufficient ventilation of the pilot and the passenger cabin during flight and landing.

- the cooling of radio navigation equipment.

4.2. DISTRIBUTION OF AIR Hot air for heating purposes is brought by means of a piping to the annular ejector where it is

mixed with cold air brought from the outside by means of an air intake. The air intake is provided

with a flap controlling the cold air flow rate. The heating or ventilating air is then conveyed by

hoses and pipes from the annular ejector via the silencer to the air ducts situated along both sides

of the passenger cabin. From the ducts, the air gets into the passenger cabin through slots.

Located downstream of the silencer is a distribution piece for fanning the side windows and pilots'

legs. The heating or ventilating air is supplied to the distribution piece through a hose. The

distribution piece is provided with a common control flap for side window and pilots' legs fanning.

In addition, the distribution piece comprises a flap for a separate control of the air for side

windows fanning. The flow rate of the air for the heating of pilots' legs is controlled by independent

flaps situated at the fanning outlets. The air intake also comprises a branch for inverter cooling.

Apart from the system described above, the aeroplane is equipped with a system of individual cold

air showers for each passenger and pilot. The cold air for passengers enters through airscoops

situated in the fuselage skin. From there, the air is led to air channels provided with air showers.

The cold air for pilots, enters through airscoops situated in front of the windshield and is further led

by a piping to air showers located on the left-hand and the right-hand control panel in the pilots'

compartment. The cold air intakes are provided with a drain piping going outside the plane. The

venting of exhausted air from passenger cabin is secured by its design. Furthermore there is a

KA 33 blower installed in the cockpit between frames 4 and 5 which serves for the cooling of radio

- navigation equipment.


FIG. 4-1 DISTRIBUTION OF AIR - SCHEMATIC - - - - - - - - - - -

- - - A -FOR HEATING AND VENTILATION

B - TO AIR SHOWERS

a - cold atmospheric air intake

b - hot air supply

(1) Air intake with flap, (2) Annular ejector, (3) Silencer, (4) Distribution piece,

(5) Air duct, (6) Outlet for pilots' legs fanning, (7) Outlet for side window

fanning, (8) Airscoop, (9) Air channel, (1 0) Air shower, (1 1) Air intake, (1 2)

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9, E20 -- - - - ~

4.3. HEATING

The hot air for the heat~ng is tapped from both engines after the last (radial) cormpressor stage.

The hot air is conducted by a piping provided with thermal ~nsulation. In the area of the engine

nacelles, the piping comprises a branch from which the hot air IS taken for engine air intake

deicing - see Chapter 5. The pipings running from the left-hand and the rlght-hand englne join in a

mixer in the wing centre section. Located on the pipings before their entry into the mixer are non-

return valves, a comperisator and a pipe coupling. The rnixer cor~~prises a branch where the hot

air is tapped for airframe deicing - see Chapter 5. Welded to the rnixer is also a test pressure

connection with a blinding plug. The hot air from the rnixer is led to a shut-off flap. The shut-off

flap controls the hot air flow rate. From the shut-off flap, the hot alr is led by a piping with

compensators to an annular ejector. The annular ejector incorporates a branch through which the

hot air is supplied to the storage battery bay.

FIG. 4 2 k1EKTING SYSTEM - SCHEMATIC

( 1 ) Non-return valve, (2) Shut-off flap, (3) Compensator, (4) Mixer,

(5) Compensator, (6) Storage battery bay heating.


4.4. TEMPERATURE CONTROL

The heating and ventilation system is controlled by the:

- VENTILATION lever,

- HEATING lever,

- COCKPIT AIR CONTROL levers,

- COCKPIT AIR CONDING.

The temperature of the air in the heating and ventilation system is indicated by a set of

instruments. The VENTILATION control lever opens and closes the atmospheric air intake flap.

The HEATING control lever opens and closes the shut-off flap of the hot air tapped from the

compressors. If it is necessary to heat the passenger cabin, the two control levers are opened so

that the temperature in the air ducts does not exceed 80°C. As soon as the passenger cabin

temperature reaches the selected value (the recommended temperature is from 22OC to 24OC),

the two control levers are reset to maintain the selected temperature without its further increasing.

The pilots' cabin and the passenger cabin are ventilated with cold air if the HEATING lever is in

the fully closed position and the VENTILATION lever is opened. Apart from the heating and

ventilation system described above, each passenger or pilot has an air shower at hislher disposal

which allows for adjusting the cold air flow rate individually. The cold air flow trom the shower is

available after turning the lid of the shower counterclockwise and deflecting its spheric body in the

desired direction. The COCKPIT AIR CONTROL levers control the flaps regulating the flow rate of

the air blowing on the side windows of the cockpit and the pilots' legs. The air for fanning the

pilots legs and the cockpit side windows is only available when at least one of the

HEATINGNENTILATION control levers is in the open position. The fanning proper is then

activated by moving the COCKPIT AIR CONTROL lever to its open position (upward). If the right-

hand COCKPIT AIR CONTROL lever is in the lower position, the air is conducted to the outlets at

pilots' legs where each pilot can individually adjust the air flow rate by control lever marked

COCKPIT AIR CONDING. If the right-hand COCKPIT AIR CONTROL lever is in the

WINDSHIELD position, the air is brought to the side windows ot the pilots' cabin. If t l w lever is in

the middle of its travel, the air is distributed evenly between the windows and the pilots' legs.

NOTE: The passenger cabin can be heated faster if the cockpit air control is out of ~,lwration (the

left-hand COCKPIT AIR CONTROL lever is in its closed position 1.e. down).

4.5. INDICATION

Air temperatures 1r1 the Ileatiny and ventilation syslem of the airplane are ~nd~cated by means of

an indication system. The air-conditioning syste111 temperature is rnonitored on the LUN 5610.01.8

dual air temperature indicator, its bottom scale shows the air temperature in the heating air ducts,

the top one that in the passenger cabin. The air-conditioning system ternperature indication is

turned on by switching on the CENTRAL WARNING DISPLAY-AIRFRAME circuit-breaker.


FIG. 4-3 TEMPERATURE CONTROL SYSTEM - SCHEMATIC


u

LUN 5615-8 LUN 5616-8

2 7 V v 1

I I I

CENTRAL WARNING DISPLAY i I AIRFRAME I I

0 L-, I I I I I I

FIG. 4-4 TEMPERATURE INDICATION - SCHEMATIC

AIRCRAFT TRAINING MANUAL i 410 UVP-E, E9, E20

CHAPTER 5

ICE AND RAIN PROTECTION

5.1. General

5.2. Pneumatic deicing system

5.3. Air intakes deicing

5.4. Heads pressure deicing

5.5. Windshield and wiper unit deicing

5.6. Propellers deicing

5.7. Deicing detection



5.1. GENERAL

The following systems have been installed on the alrcraft front sections as the protection against

the ice formation:

- the pneumatic deicmg system for deicing the leading edges of wings and tail unit

- the system of hot-air heating of leading edges of the engine air intakes

- the system of the electric heating of the Pitot tubes, static and ram pressure heads

- the system of heated front windshield including the installation of the wiper unit for removal of the

hoarfrost, snow, ice and dust from the p~lot's cabin windshields

- the system of the electrlc heating of the propeller blades leading edges for deic~ng of the

propeller blades lead~ng edges

- the system of the ice detection and signalization

5.2. PNEUMATIC DEICING SYSTEM

The pneumatic deicing system serves for the ice removal from the leading sections of wings and

tail unit. The function of the de~cer is based on the mechanical action of the flexible rubber boot,

which is fixed to the leading edge of wings and tail unit. In this deicer there is a series of small

cells, and their quick filling with air causes the fracturing of the ice accretion layer, which in this

way is released from the surface and then removed by means of the air stream.

Description

Hot compressed air is taken from the flange of the mixer (27) - see the fig. 5-1, which is a part of

the air conditioning system (see the chapter 4). On the mixer, there is the connection of the

external air source (29) intended for the deicing test at the engines standstill. This connection is

blinded. Air from the mixture is fed through the compensator (15) anr .,le elbow (21) to the

reduction valve (9). The plug (22) w~th 3 outlets is fixed on the reduction valve.

Through the first out let air is fed by the piping to the air pressure transmitter (10). By means of the

next two outlets air is fed through the piping to the single solenoid valve (1 1) and to the double

solenoid valve (12). With the single solenoid valve in open position air is fed through the piping to

the coupling (26), from which it is fed by the piping to the deicers (5), (6 ) , (7) mounted on the tail

unit. With the double solenoid valve on, air is fed by means of the T-connections (23) and the

piping to the deicer s ( I ) , (2), (3) and (4) installed on the wing. The small plates (30) on individual

DEICERS serve as the conductive interconnect~on between the conductive surface of the deicers

and the wing or the tail u n t


FIG. 5-1 LOCATION OF COMPONENTS FOR DEICING THE WINGS AND TAIL UNIT

(1) P 20-1 Rubber deicer, (2) P 20-2 Rubber deicer, (3) P 25-1 Rubber deicer,

(4) P 25-2 Rubber deicer, (5) P 24-1 Rubber deicer, (6) P 24-2 Rubber deicer,

(7) P 26-1 Rubber deicer, (a), (9) LUN 6656-8 Reduction air valve,

(10) LUN 1562-8 Air pressure transmitter, (1 1) LUN 2477.01-8 Single solenoid

valve, (12) LUN 2477.02-8 Double solenoid valve, (1 3) LUN 3294-8 Control box

of the airframe deicing, (14) LUN 3295-8 Electronic timer of the airframe deicing,

(1 5) Compensator, (1 6) Circuit breaker DEICING-AIRFRAME, (1 7) Bridge,

(18) Clamp, (19) Baffle wall, (20) Clamp, (21) Elbow, (22) Plug,

(23) T-connection (24) Piping, (25) Loop, (26) Coupling, (27) Mixer, (28) Beam,

(29) Connection of the external air source with a blind, (30) Plate, (31) Bracket,

(32) Water separator.


Operation

CAUTION:-THE SYSTEM OF THE PNEUMATIC DEICING SHOULD NOT BE TURNED ON AT

THE TEMPERATURES LOWER THAN -300C. NO DAMAGE OF DEICERS IS

CAUSED UP TO THE TEMPERATURES OF -60°C.

- DEICERS CORRUGATION IS NO DEFECT UNDER CONDITION THAZ THEY ARE

SUCKED TOWARD WING CONTOUR IN 5 MINUTES AFTER ENGINE START.

The actuation of the pneumatic deicing system is carried out by turning on the DEICING-

AIRFRAME circuit breaker (16) mounted on the overhead panel.

The main switch is turned to the position ON hence putting the deicing system into operation. On

the control box of the airframe deicing 3 bulbs light up. The selector of the operation is in the

AUTOM. position. The cycling is accomplished by help of the electronic timer (14), this switching

on separate solenoid air valves and in this way the air stream is fed to the separate sections of the

rubber deicers (sections A, 6, C) and the cells in the deicers are inflated. Upon completing the

deicing, the electronic timer sets individual solenoid air valves so, that the Venturi tube starts

operating, this being a part of the solenoid air valves. The negative pressure appears in the

system and the cells in the deicers get flat. This is repeated continuously, till the main switch on

the control box of the airframe deicing does not turn off. The filling of deicers with the pressure air

is checked on the pressure gauge inside the control box. Switching over individual sections is

indicated by the lighting up of the control lights on the control box (always for the relevant section

A or B or C). Cycling can either be speeded up or slowed down by the FAST-SLOW speed

selector which is also mounted on the control box.

The diagram of the work cycle for the automatic control

-the speed selector is in the FAST position

- the speed selector is in the SLOW position

A - section A cycle

B - section B cycle

deicers operating

I 1 deicers not operating

C - section C cycle


FIG. 5-2 SYSTEM DIAGRAM OF WINGS AND TAIL UNIT DEICING

AIRCRAFT TRAINING MANU~AL L 410 UVP-E, E9, E2O

@b - D E I C I N G AIRFRAME


In case, we don't want to use the automatic cycling, the operation selector is switched to the

MANUAL position and the cycling is accomplished by switching over individual sections to the

position A, B or C. Even in the case that the system of the pneumatic deicing is not put into

operation, it is necessary to turn on the DEICING-AIRFRAME circuit breaker mounted on the

overhead panel immediately upon starting the engines.

The main switch on the control box is in the OFF position. Due to this the Venturi tubes are

actuated. As the Venturi tubes are a part of the solenoid air valves, the under pressure appears in

the system of the pneumatic deicing and the cells in the deicers are sucked toward the leading

edges.

CAUTION: IF NOT FOLLOWING THIS PROCEDURE (TURNING ON THE DEICING AIRFRAME

CIRCUIT BREAKER WITH THE MAIN SWITCH ON THE CONTROL BOX IN THE

OFF-POSITION) THEN NO UNDER PRESSURE IS FORMED IN THE SYSTEM OF

THE PNEUMATIC DEICING AND THE CELLS IN THE DEICERS ARE NOT SUCKED

TOWARD THE LEADING EDGES HENCE CAUSING THE DISTORTION OF THE

AERODYNAMIC PROFILE OF THE LEADING EDGES.

5.3. AIR INTAKES DEICING The leading edge of the air intake to the engines is hot air heated, the air being taken from both

engines behind the last (radial) compressor stage. The opening of the hot air supply is derived

from the control of the ice separator vanes inside the air channels.

Description

Hot air distribution system

Hot air is fed through the piping via the shut-off cock (1) - see fig. 5-3 - to the leading edge of the

engine air intake.

Inside the piping, ahead of the shut-off cock, there is a branch pipe, through which hot air is taken

for the air conditioning system - see chapter 4.

Heating of the leading edge of the air intake to the engine is carried out along its entire periphery.

From the inner side of the air intake to the engine, the surface is heated to the depth of 110 mm,

from the outer side of the air intake to the engine the surface is heated to the depth of 30 mm.

Heating of this surface is enabled by a gap created between the skin (2) and the inner jacket (3).

Hot air is forced into this gap through the ports, made along the piping periphery (4). The hot air

supply to the piping (4) is provided at two points and created by the branched supply piping (5)

from the shut-off cock. This arrangement secures uniform distribution of the hot air along the

entire periphery of the engine air intake and also the uniform temperature of the heated surface.

The discharge of the used air from the gap space to the engine space is by means of im pressed

channels formed in the rear side along the gap periphery.

FIG. 5-3 HOT AIR DISTRIBUTION TO THE LEADING EDGES OF THE AIR INTAKE

(1) UK 1 Shut-off cock, (8), (2) Skin, (3) Inner jacket, (4) Piping, (5) Supply

piping, (6) Bracket, (7) Coupling, (8) Pipe coupling, (9) Searchlight,

(10) Connection, (1 1) Screw, (12) Sealing ring, (13) (14) (15) Piping,

(16) Sealing, (17) Clamp, (18) Straight coupling, (19) Shim, (20) Bracket,

(21) Lever, (22) Spring, (23) Air intake to the engine, (24), (25), (26) Shim,

(27) Terminal.

a - hot air f - 13 ports o 1.5 mm

b - used air discharge g - 3 ports o 2.0 mrn

c - 11 portso 1.5 mm h - 5 ports o 1.2 mm

d - 5 p o r t s o 1.7 mrn i - p o r t o 1.7 mm

e - 7 portso 1.9 mm k -por t0 1.5 mm


Control system of separator vanes in the air channel

The protection from penetration of ice pieces onto the engine screen with the deicing system on is

provided by a system of resetting vanes ( I ) , (2) - see Fig. 5-4, inside the air channel of the engine

nacelle. Vanes (1) and (2) have coupled motion. During the operation under normal conditions

both vanes are closed, i.e. in the position K. With the deicing on, the vanes are set to the position

m. In this position, the vanes cause, that the path of the ice pieces, passing through the airchann

el, is deflected from the suction space into the engine and due to their kinetic energy they fly

through the rear open vane (2) out from the engine nacelle. - - - - - - - - - - - - - - - - - - - - -

- - - -

The resetting of the vanes is accomplkhed by the MP-100 MT electromechanical strut, this being

fixed above the air channel. The motion from the electromechanical strut (1 1) see fig. 5-5 is

transferred to the vane (1) by means of the angular lever (2) and the adjustable pull rod with a fork

(13). The amount of the vane opening (1) is given by the stroke of the electromechanical strut.

The adjustment of the separator vane consists only in the vane setting to the felt stops (4) in the

upper (i.e. closed) position.

The coupling with a vane (2) is accomplished by means of a push-pull cable (5). Due to the

extensive sucking forces, acting upon the vane (2), its control is accomplished by help of the yoke

(8) with an adjustable lever (10). The yoke and the adjustable lever are covered with a cap (18).

Beside the vanes (1) and (2) the vane behind the oil cooler (3) and the shut-off cock (9) are still

more controlled by the MP-100 MT electromechanical strut and the push-pull cables (6) and (7).

The vane opening (3) is adjusted for high and low ambient air temperatures. - - - - - - - - - - - - - - - - - - - - - -

- - - -

The c6ntrol o f the vane (3) behindthe-oil cooler and hence the increase of this vane opening has

been made to increate the efficiency of the oil cooler with the deicing on, when a portion of air

from the engine air channel passes through the (2) out of the engine nacelle.

The push-pull cable of the shut-off cock is terminated with a removable joint by help of double-arm

levers with rollers. One double-arm lever, with rollers (25) is fixed to the bracket and the push-pull

cable (7) is fixed to it by means of a fork. The second double-arm lever with rollers, to which the

motion is transfered, is firmly fixed to the shaft of shut-off cock (9). The box of terminal switches

(24) is fixed to the bracket in the engine cowl above the air channel, this box of terminal switches

being controlled by an angular lever (1 2). The fulcrum forms a trunnion which is connected by a

clip (20) to the shaft of the box of terminal switches.


FIG. 5-4 DIAGRAM OF THE AIR STREAMING IN THE DEICING SYSTEM OF VANES

INSIDE THE AIR CHANNEL

(1) Separator vane, (2) Vane, (3) Vane, (4) Screen, (5) 011 cooler

a - air inlet

b - air streaming to the engine with the delcing on

c - air streaming to the engine under normal operation (without the deicing on)

d - space of the air suction to the engine

e - air streaming to the oil cooler

f - Ice preces

h - air outlet from the 011 cooler

k - vane posit~ons under normal operation (the deicer off)

m - vane positions with the deicing on


FIG. 5-5 CONTROL OF THE SEPARATOR VANES

(1) Separator vane, (2) Vane, (3) Vane behind the oil cooler, (4) Felt stops,

(5) Push-pull cable, (6) - , (7) Push-pull cable, (8) Yoke, (9) Shut-off cock,

(1 0) Adjustable lever, (1 1) MP-100 MT Electromechanical strut, (1 2) Lever,

(13) Pull rod, (14) Clip, (15) Adjusting screw, (16) Ball joint with the trunnion,

(17) Lubricator, (18) Cap, (19) Fork, (20) Clip, (21) Coupling, (22) Clamp,

(23) Sleeve, (24) SKP 1 Box of terminal switches (8) U, (25) Lever with rollers,

(26) Fork, (27) Nut with washer.

ltem II - up to 22. series aircraft ltem IV - from 23. serie on NC.


Operation

During the no-deicing mode, the electrlc circuits are broken, the shut-off cock of the hot air supply

for heating the leading edges of the engine air intakes is closed, the separator vanes are in closed

position. The SEPARATOR VANE s~gnalling cell is off. By closing the SEPARATOR VANE LH

(RH) circuit breaker and by switching over the SEPARATOR VANE LH (RH) selector to the ON

position, the electromagnetic strut is actuated, the shut-off cock of the hot air supply opens and

the separator vanes are reset to the open position. The course of the intermediate position is

signalized by the intermittent light of the SEPARATOR VANE signalling cell on the central warning

display. Upon - achieving - - - - the extreme position, in which the shut-off cock is open, the separator - - - - - - - - - - - - - - - - - - -

vanes are in the open position, and the SEPARATOR VANE signalling cell hghts permanently.

The signal bulbs are put into operation by the box of the terminal switches, which is controlled by

the angular lever

By switching over th e SEPARATOR VANE LH (RH) selector to the OFF position, the

electromechanical strut starts operation, the shut-off cock of the hot air supply closes and the

separator vanes are reset to the closed position. The course of the intermediate position is

signalized by the SEPARATOR VANE intermittent light of the srgnalling cell.

Upon achieving the extreme position, in which the shut-off cock IS closed, the separator vanes are

in the closed position, the SEPARATOR VANE signalling cell goes off.


CENTRAL WARNING DISPLAY LIi Rff LIi ENGINE R l i ENGINE -

! CENTRAL WARNING DISPLAY I LHENGINE RH ENGINE

SEPARATOR VANE 1 I SEPARATOR

VANE

FIG. 5-6 DIAGRAM OF THE CONTROL SYSTEM OF SEPARATOR VANES


5.4. HEADS PRESSURE DEICING

Pitot tubes, static pressure heads and ram pressure head have their own inner heating by means

of the d.c. current.

Heating of the LUN 1157-7 Pitot tubes is accomplished by means of heating elements of 140 W

power, installed directly inside the Pitot tubes.

In similar way is also heated the LUN 11 56-7 static pressure head (the heating element of 53.2 W

power) and the ram pressure head (the heating element of 45 W power).

Operation

- - - - - -

~ h e h e d i n ~ o f Pitottubes andthestafic pressurelieads isclosed by the PITQT-STATIC-I, IC -

circuit breakers and the PlTOT HEADS I, II push-button switches and the STATIC HEADS I, ! I .

The heating of the ram pressure head is closed by the STALL PROBE circuit breaker and the

STALL PROBE push-button switch.

If the heating is operating, the signalization in the knob of the push-button switch l~ghts up. In case

of the circuit failure the light goes off.

The failure is also indicated by PlTOT HEATING signal light. The signal will light up after switching

on the CENTRAL WARNING DISPLAY - AIRFRAME circuit breaker switch on overhead panel.

After switching on all HEATING push-button switches, on right control panel, the signal light will go

off. In case of circuit failure the signal hght will light up.

The signal light is on also In case, when are not all the push-button switches on.

CAUTION: ON THEGROUND,_WHENTHEAIR_CRAFTIS AT THE STANDSTILL, - THE -

HEATING CAN ONLY BE SWITCHED ON FOR A SHORT TIME JUST FOR

CHECKING ITS FUNCTION (APP. 1MIN.).

The heating of all sensors (heads) is switched off by pressing the small, rectangular push-button

below the push-button with the signalization.


CENTRAL WARNING DISPLAY id$ - DEICING-

AIRFRAhlE PITOT-STATIC STALL PROBE I I

i I I I I

i b@ i CENTRAL WARNING DISPLAY

PITOT TUBE

- m\ i f installed IW I

FIG. 5-7 DIAGRAM OF THE HEATING SYSTEM OF PlTOT TUBES, STATIC

PRESSUREHEADANDRAMPRESSUREHEAD


5.5. WINDSHIELD AND WIPER UNIT DEICING For the removal of hoar frost, snow and this layer of ice from the windshields of the flight

compartment, the system of windshield heating is mounted ,on the aircraft, by means of which the

hoarfrost, snow and ice melt.

Melted ice, snow or hoarfrost are removed by means of the wiper unit.

Operation of heated windshields system

The electric heating of the heated windshields system is a two-stage one. In the 1st stage, the

entire areas of both windshields are heated up to 30% of the power when both AC generators are

operating.

In the llnd stage, the entire areas of both windshields are heated up to 100°/o of the power when

AC generators are operating. In case of failure of one AC generator, the entire windshield heating

is transferred automatically onto the other AC generator. The system is actuated by switching on

the BATTERY I, II, WINDSHIELD HEATING LH, RH switches the AC GENERATOR LH, RH,

CENTRAL WARNING DISPLAY circuit breakers and by switching over the WINDSHIELD

HEATING selector to the 1st heating stage.

NOTE: Switching over of the WINDSHIELD HEATING selector to the llnd stage is only possible

after 5 to 7 minutes of the windshield heating at the 1st stage.

The operation of individual heating sections is controlled automatically by means of the

thermoregulator mounted in the windshield and set up to the temperature of 30°C. If exceeding

the set up temperature, the heating is switched off automatically. The closing of the windshield

heating system is signalized by the green WINDSHIELD HEATING signal lights, mounted on the

left and right control panel. The function of the thermoregulator is checked by the TEST OF

WINDSHIELD HEATING switches. When with the heating of the windshield on the WINDSHIELD

HEATING signal light is on, then by switching over the TEST OF V;..qDSHIELD HEATING

selector to the HEATING ON position the WINDSHIELD HEATING signal light extinguishes.

When with the wmdshield heating on the WINDSHIELD HEATING signal light IS not lighting, then

by switching over the TEST OF WINDSHIELD HEATING selector to the HEATING OFF position

the WINDSHIELD HEATING signal light lights up. Any other state means a failure.

Operation of wiper unit system

The windshield wiper IS driven by the GA-21 lA.OO-4 mechanical unit of windshield wiper which is

controlled by the throttle cock of the GA 17116 wiper. Hydraulic fluid is fed through the throttle cock

of the GA 17116 wiper from the circuit of the permanent pressure to the mechanical cleaner of the

GA-21 lA.OO-4 windshield wiper. According to the amount of the wiper throttle cock opening the

velocity of the wiping unit motion is controlled.

TEST OF WINDSHIELD I ~ E A T I N ~ I LII HEATING ON KH 1 I

HEATING OFF

FIG. 5-8 DIAGRAM OF THE HEATED WINDSHIELDS SYSTEM


AC d GENERATOR

i I D E I C I N G I

WINDSlIIELD HEATING I

R H AIRFRAhlE L---


Windshield wipe r

- - - - - - - - - - - - - -

c i r c u i t of t h e main network

r e t u r n c i r c u i t L . - . - . - - - . - . -

FIG. 5-9 DIAGRAM OF THE WIPER UNIT INSTALLATION SYSTEM


5.6. PROPELLERS DEICING

Propellers are electrically deiced by means of the heating elements glued on the leading edges of

the propeller blades The deicing system of propellers consists of two sections: parts rotating

together with the propeller (transfer rings and deicing elements of the propeller blades) and parts

not rotating, installed inside the airframe (timers,collectors, contactors, switches, push-buttons,

fuses and signal bulbs).

Operation

The deicing is performed within the cycle, the interval of which has been set on the PROPELLER

DEICING - STBY - MAIN selector on the instrumental panel.

The selector position I (closer to the neutral) corresponds to the interval of 40 sec., the position II

(more distant) corresponds to the interval of 80 sec.

Under normal conditions the MAlN circuit of the cycler is selected. The correct function of the

propellers deicing is indicated so, that both PROPELLER DEICING signal cells on the LEFT

ENGINE, RIGHT ENGINE signalling block are extinguished. The interval of the propellers deicing

cycle can be shortened to 4 sec. by switching over the PROPELLER DEICING selector to the

MAlN position, and by pressing the TIMER push-button on the left control panel. With the correct

function of the propellers deicing, the PROPELLER DEICING s~gnal cells light up in the shortened

4 seconds cycle as follows:

left - none

right - both

whilst the beginning of the signalling can take place in any phase.

Any other cycling indicates the system failure. By pressing the BLADES push-button on the left

control panel (the PROPELLER DEICING selector switched to the MAlN or STBY position), the

checking of the propeller blades deicing elements is carried out. With the proper function, both

PROPELLER DEICING signal cells on the signalling block must be extinguished Any other

condition indicates a fa~lure.


CENTRAL WARNING DISPLAY Lti ENGINE RH ENGINE

C E T R A ~ W A R N I N C DISPLAY ENGINE

I PROP DEICING I

FIG. 5-10 SYSTEM DIAGRAM OF THE PROPELLERS DEICING

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E20 -- -

5.7. DEICING DETECTION To enable the screw of the aircraft to detect whether the ice is bemg formed and at what ~ntens~ty,

the following detectors are mounted on the aircraft:

- rotary detector

- static detector

Operation

The rotary ice detector is actuated by the ROTARY ICE DETECTOR switch. The ice accretion

growth on the rotating sensor of the ice detector is signalized by the lighting of the ICING signal

cell on the signalling block.

The static ice detector indicates on its front side the hoarfrost or ice accretion growth. This ice

layer, corresponds to the layer of ice on the wings leading edges, on the tail unit and on the

engine air intakes hence drawing the pilot's attention on the necessity of the ice removal from the

aircraft by means of the deicing device. After flying through the zone with the ice accretion or ice

occurrence, it is possible to remove the ice accretion or the ice by switching on the heating of the

STATIC ICE DETECTOR circuit breaker.

I I CENTRAL WARNING DISPLAY ICE DEnC-OR

LUCTRO ROTARY STATIC

FIG. 5-1 1 DIAGRAM OF THE ICE DETECTION



CHAPTER 6

HYDRAULIC POWER

6 1 General

6 2 Ma111 hydraul~c system

6 3 Normal control

6 4 Emergency control

6 5 I-lydraulc tank pressur~zat~on system

6 6 lnd~catmg



- - - - - - - - - - -

AIRCRAFT TRAINING MANUAL L 410 UVP-L:, E9, E20

6.1. GENERAL

The hydraulic system has a steady rated pressure of 14.7 MPa, (150 kplcm2) and is intended for

control of aeroplane moving parts and units.

As the pressure hydraulic energy sources there are two LUN 6102.01-8 automatic hydraulic piston

pumps. The hydraulic pumps are located directly on the engines and work independently of each

other. That is why in case of supply break from one of the pumps, the second pump is sufficient to

supply a necessary amount of energy to the hydraulic system. All the system is protected by

LUN 7545-8 relief valve.

As the emergency hydraulic energy source there is LUN 6100.03-8 hand-operated hydraulic pump

provided with an emergency amount of fluid in a separate emergency hydraulic tank.

The piping is mostly made of light alloys. The pressure piping of Js8 (and greater) inside diameter

and all piping in the engines fire zones are made from stainless steel.

In order to facilitate orientation, all piping is marked with symbols. The symbols are composed of

serial number, letter H and a letter corresponding to the circuit. The hydraulic system is filled with

AMG-10 hydraulic fluid according to the standard GOST 6794-75 (boiling point 200°C, flash pomt

92OC, point of congelation -70°C, density 0.85 kg/dm3) or with Aero Shell Fluid 41 (4) complying

with the MIL-H 5606 standard.

6.2. MAIN HYDRAULIC SYSTEM

The main hydraulic system of the aeroplane consists of normal and emergency control systems

and the hydraulic pressurization system.

The normal control hydraulic system controls retraction and extension of landing gears, wing

flaps, interceptors and banking tabs, nose wheel steering servo, main wheel brakes, parking

brakes and wiper unit.

The normal control system is divided into the following constant-pressure operating circuits:

- constant - pressure circuit

- circuit of retraction and extension of landing gears

- circuit of retraction and extension of wing flaps

- circuit of nose wheel steering servo

- circuit of main wheels brakes

- wiper unit control circuit

- circuit of retraction and extension of banking tabs

- circuit of retraction and extension of interceptors

The emergency control hydraulic system controls emergency extension of landing gears and

emergency extension of wing flaps to landing position, emergency and parking braking.


6.3. NORMAL CONTROL Description

The normal control hydraulic system consists of the constant-pressure circuit to which all working

circuits distributing the hydraulic power to individual units are connected. The instruments and

piping are mostly arranged in the left-hand engine nacelle, left-hand landing gear nacelle, in the

wing and in the fuselage nose part. The instruments are connected with piping, only the hydraulic

pumps are connected to the other parts of the system the by means of high-pressure hoses.

The hydraulic pumps are connected through the suction hose and piping with the hydraulic tank

which is located in the wing, above the left-hand engine nacelle. Uncoupling valves (5) - fig. 6-1

are intended for disconnecting of hydraulic pumps from the hydraulic system without losing

hydraulic fluid. Pressurized hydraulic fluid is delivered by hydraulic pumps via non-return valves

(1 4) and hydraulic filter (1 0) to the constant-pressure circuit.

Suction filling branch pipe (2) is intended for filling the hydraulic system with hydraulic fluid from a

ground source, independently of hydraulic pumps and together with the pressure branch pipe for

testing of functions of individual working circuits on the grbund. 7 hey are located in the left-hand

engine nacelle. Constant pressure circuit has a hydraulic accumulator (16) in the left-hand engine

nacelle to maintain constant pressure in the network, and relief valve (8) which protects the

hydraulic system against pressure overload by connection of the discharge piping leading in to the

hydraulic tank. The brakes circuit has its own hydraulic accumulator (17) located in the left-hand

landing gear nacelle. Dual pressure gauge (47) situated on the instrument panel, together with

pressure transmitters (51) located in the left-hand landing gear nacelle, are intended for pressure

check in the main system circuit and in the brakes circuit. Before pressure transmitters (51),

chokes (55) are inserted.

The return constant-pressure c~rcuit has its own hydraulic filter (1 I ) in the left-hand engine

nacelle, in front of the entry to the hydraulic tank. The return hydraulic pumps circuit has a

hydraulic filter (12) in the left-hand and right-hand engine nacelles (up to serial No 1714). The

following operating circuits are connected to the constant-pressure circuit:

- landing gear circuit controlled by solenoid valve (19),

- wing flaps circuit controlled by solenoid valve (1 9),

- wiper unit circuit controlled by wiper throttle cock (41),

- spoiler circuit controlled by solenoid valve (54),

- banking tabs circuit controlled by solenoid valves (54),

- nose wheel steering servo circuit controlled by solenoid valve (54),

- parking brake circult controlled by hand-operated valve (39)

- main wheels braking circuit controlled by pressure reducing valve (37)


Operation

Two two-stage hydraulic automatic pumps (4) adjusted to pressure 14.4 MPa - 0.39 MPa

(147 kp/cm2 - 4 kp/cm2) suck hydraulic fluid from hydraulic tank (50) and force it via non-return

valves (14), hydraulic filter (10) and non-return valve (13) to the,main circuit.

When connecting the qround source, which is in principle a hydraulic pump provided with an

electromotor, the hydraulic fluid is sucked from hydraulic tank (50) via suction filling connnection

(2) and pressure filling connection ( 3), hydraulic filter (10) and non-return valve (13) to the main

circuit. As a part of the main circuit there is a hydraulic accumulator (16) intended for holding

constant pressure, dual pressure gauge (47) and relief valve (8) protecting the circuit against

pressure overload. If pressure in the circuit rises up to 16.2 MPa + 0.2 MPa (165 kp/cm2

+ 2 kp/cm2), relief valve (8) lets hydraulic fluid flow into hydraulic tank (50) and recloses when

pressure drops to 15 MPa (153 kp/cm2).

To the main circuit particular operating circuits are connected.

The brakes circuit is corlr~ected via non-return valve (15). The hydraulic accurr~ulator (1 7) is

inserted in this circuit which ensures pressure power for braking in case of a fault of the constant-

pressure circuit. Pressure reduction valve (37) maintains constant pressure of 4.9 MPa

(50 kp/cm2) in the brakes operating circuit. Solenoid valves (19) control and the circuit for

retraction and extension of landing gears the circuit for retraction and extension of wing flaps.

Solenoid valves (54) control the nose wheel steering servo circuit, and the banking tabs extension

and retraction circuit, and spoiler circuit. Wiper throttle cock (41) is intended for wipers control. In

the return constant-pressure circuit there is a hydraulic filter (1 1). In the return hydraulic pumps

circuit there are hydraulic thermoswitches (53) and hydraulic filters (12) (up to serial No. 1714). In

the aeroplanes bearing factory No. 171 5 and senther the hydraulic filter is in stalled only in the left-

hand engine nacelle. On aircraft from the 22nd series there is a hydraulic fluid sampling cock (65)

installed in the left-hand engine nacelle.


FIG. 6-1 CONSTANT - PRESSURE CIRCUIT DIAGRAM

Legend to the figure - see the following page ,


Suction filling connection

Pressure filling connection

Hydraulic pump

Disconnecting valve

Relief valve

Hydraulic filter

Hydraulic filter

Hydraulic filter

Non-return valve

Non-return valve

Non-return valve

Hydraulic accumulator

Hydraulic accumulator

Non-return valve

Solenoid valve

Hydraulic tank

Choke 0 1 mn

Reduction valve

Hand - operated valve

Wiper throttle cock

Hand - operated valve

Dual pressure gauge

Pressure transmitter

Hydraulic thermoswitch

Solenoid valve

Choke

Choke

Legend to fig. 6-1:

LUN 7741 -8

LUN 7740-8

LUN 61 02.01 -8

LUN 7366-8

LUN 7545-8

LUN 7614.01 -8

LUN 7614.03-8

LUN 761 3.02-8

LUN 7549-9

LUN 7560-8

674 600 B (8)

LUN 6953.05-8

LUN 6953.02-8

674 500 B (8)

GA 163 TI1 6

B 057 801 N

XL 41 0.4630-42

LUN 7514.02-8

LUN 6578-8

GA 17116

LUN 6577-8

UI 2 - 240 K

ID - 240

LUN 31 92-8

GA 184 U/2

0-002

B 057 049 N

Ser. No. Name Drawing No. or designation

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E 2 O

58 Hydraulic - hand pump LUN 61 00.03-8

60 Emergency tank B 057 430 N

64 Discharge valve B057 120N

65 Hydraulic fluid sampling cock LUN 7386 (from 22nd series)

68 Non-return valve B 057 824 N

Legend to fig. 6-l(cont.):

a -Main wheels braking circuit,

b - Wipe unit control,

c - Nose wheel steering servo circuit,

d - Circuit for landing gears extension and retraction,

e - Circuit for banking tabs extension out and retraction,

f - Circuit for wing flaps extension and retraction,

g - Circuit for interceptor extension and retraction,

h - Circuit for hydraulic tank pressurization system.

6.4. EIERGEBCCY CONTROL The emergency circuit is independent on the constant-pressure circuit. The emergency tank (60)

- fig. 6-1 rs refilled with hydraulic fluid from the brakes return circuit. The hydraulic hand pump (58)

sucks hydraulic fluid from the emergency tank and delivers ~t w~th max. pressure of 9.8 MPa

(100 kp/cm2) to the emergency control circuits of landing gears, wing flaps and parking brake.

6.5. HYMAtLLIC T A M WESSURIZATION SYSTEM The hydraulic tank press~~rization system is intended for ensuring air overpressure in the main

hydraulic tank. By pressurization the hydraulic tank, the optimum conditions are established for

operation of LUN 6102.01-B hydraulic pumps. The instruments relevant to that system are

arranged in the left-hand engine nacelle and in the wing above the left-hand engine nacelle.

Operating pressure in the main hydraulic tank is 0.1 MPa +0.12/-0.03MPa (1 +I .2/ -0.3 kp/cm2)

Description

After starting the engine, pressure of air in the main hydraulic tank reaches the operating values in

about 30 seconds. When the hydraulic system is in function, the air pressure value in the main

hydraulic tank does not vary, except the function of landing gear retraction. After having retracted

the landing gear, the pressure bill be increased by due to a portion of hydraulic fluid returning

back to the main hydraulic tank. This increase in pressure is not relieved by the safety valve (relief

valve).

After extending the landing gear, air pressure in the hydraulic tank drops to the original value.


Increase of pressure in the hydraulic tank also occurs due to hydraulic fluid returning from the

hydraulic accumulators to the hydraulic tank when the aeroplane is out of operation for a longer

time.

Operation

Air is supplied to the system from the engines compressor and at maximum overpressure of

0.54 MPa (5.5 kp/cm2). The system proper is connected to a mixer. The non-return valves (70)

- fig. 6-2 prevent air leakage from the circuit after the engines have been stopped. Silica gel filling

of air desiccator (71) absords humidity from the supplied air.

Air cleaner (72) catches mechanical impurities. Auxiliary tank (73) serves as a pressure air

reservoir for compensation of external untightness o f the system. Pressure reduction valve (74)

reduces air pressure from the engines to the operating pressure in the tank 0.08 to 0.12 Mpa

(0.8 to 1.2 kp/cm2)

Safety valve (75) protects the main hydraulic tank against overload from by-passing pressure

0.2 MPa +/- 0.02 MPa (2 kp/cm2 +/- 0.2 kp/cm2). Hand-operated valves (77) are intended for

elimination of air pressure in the main hydraulic tank (designated PRESSURE RELEASE) and for

checking function of the safety valve (designated TEST). Air pressure in the main hydraulic tank is

refilled from the ground source via the non-return valve in the left-hand engine nacelle.

FIG. 6-2 HYDRAULIC TANK PRESSURIZATION SYSTEM CIRCUIT DIAGRAM

Legend to figure - see the following page

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9. E20

50

70

71

72

- - = - - -

74

75

76

77

82

6.6. INDICATING Description

Legend to figure fig 6-3:

Hydraulic tank

Non-return valve

Desiccator

Air cleaner

Auxiliqtank - - - -

Reduction valve

Safety valve

Pressure gauge

Hand-operated valve

Ground source connection

Ser. No.

B 057 801 N

LUN 7521 -8

3976 0872

723 900-4 (8)

B 057 862 N

LUN 6651 .01-8

LUN 7548-8

MA-4

LUN 7351 -8

The hydraulic indicating system consists of:

- Pressure indicating (it indicates pressure in the hydraulic circuit of the main system and brakes - - - - - - - - - - -

~

as well as air pressure-in the hydraulic tank pressurization system),

- Temperature indicating (it indicates hydraulic fluid temperature).

Operation

Pressure in the main network circuit is indicated by 2 DIM-240 telemetric induction pressure

gauge. The telemetric induction pressure gauge system consists of a UI 2-240 K dual pressure

gauge and two ID-240 with choke D 002pressure transmitters. One ID-240 pressure transmitter is

inserted in the main system circuit and one is inserted in the brakes circuit. Air pressure in the

main hydraulic tank is indicated by the pressure gauge arranged in the left-hand engine nacelle. It

is visible through the inspection hole in the cover of the left-hand engine nacelle.

Hydraulic fluid temperature in the hydraulic pumps return circuit is checked by LUN 3192-8

hydraulic thermoswitches. If hydraulic fluid temperature exceeds 85 + 5OC, the hydraulic ~ ~ ~

- - - -thermoswitch closes the respective electrical circuit and-the HYDRAULICS signal cell lights up on

the small signal block. On temperature drop the HYDRAULICS signal cell goes out again.

Inspection whether the signal cell is in order shall be performed by means of CENTRAL

WARNING DISPLAY - ENGINE L, CENTRAL WARNING DISPLAY - ENGINE R push-buttons

(while circuit breakers CENTRAL WARNING DISPLAY - ENGINE L, CENTRAL WARNING

DISPLAY - ENGINE R are switched on).

Name Drawing No. or designation


FIG. 6-3 HYDRAULIC SYSTEM INDICATING DIAGRAM


- . - - - -


CHAPTER 7

FUEL AND OIL SYSTEM

7.1 Fuel system - general

7 2. Fuel distribution

7.3 Fuel tank ventmg

7.4. Fuel tank interconnection

7 5 Fuel dramage

7.6. Fuel indicating

7.7. 011 system - general

7 8. Oil system indicating


intentionally left blank

- - - - -


7.1. FUEL SYSTEM - GENERAL The fuel system is designed to supply the engines with fuel under all operation conditions and

temperatures which can be experienced during aircraft operation in various environments. The

fuel system is designed for supplying fuel at adequate pressure at all flight altitudes in which the

aircraft is capable of being operated.

The tuel system consists of fuel tanks, connecting piping, fuel distribution components and

instruments indicating the quantity and pressure of fuel. The system is di ided into a left hand and

right hand section. Both sections are identical. Each section contains four fuel tanks in the wing

and one wingtip fuel tank.The aeroplane can be operated either with wingtip fuel tanks or without

them. The left hand fuel system supplies the left engine, the right hand fuel system supplies the

right engine. If necessary both systems can be interconnected by solenoid valves which permit to

feed one engine from both fuel systems or both engines from one fuel system. Fuel is filled in into

each system separately through filler necks.

NOTE: On some aircraft wingtip tank need not the installed and supplied in the aircraft spare

parts set.

7.2. FUEL DISTRIBUTION Description

Fuel flows by gravity from the additional outer and middle fuel tank into the collector fuel tank and

is fed further through the pressurized section to the fuel pump. The fuel pump is installed under

the collector fuel tank. Its position guarantees that the fuel system is permanently flooded with

fuel. A cut-off cock installed between the collector fuel tank and the fuel pump enables work to be

carried out on the fuel system without draining fuel from the fuel tanks. A part of fuel is routed

from the fuel pump delivery piping trough a return piping equipped with a non-return valve back

into the collector tuel tank. Most fuel is routed through a non-return "~irve to the fuel cock and to

the cross-feed solenoid valve. The fuel cock is operated mechanically from the cockpit.

Fuel from the fuel cock is routed into a oil-to-fuel heater where fuel is heated by engine oil to the

required temperature. The process is controlled by a thermoregulator in the oil-to-fuel heater

which opens and closes the oil flow depending upon the temperature of fuel at the outlet from the

heater. From the oil-to-fuel heater fuel is routed further into a fuel filter.

From the fuel filter fuel is fed through a fire resistant high pressure hose to the engine fuel pump.

A return piping leading from the fuel filter back to the middle fuel tank serves for returning

excessive fuel mixed ,with air trapped in the filter.

Installed in the return eine is another non-return valve. The fuel system of the left hand engine is

interconnected with that of the right hand engine by a piping and a pair of solenoid valves.

The wingtip fuel tank is connected by a piping to the main fuel system. This piping delivers fuel

from the fuel pump in the wingtip tank into the additional fuel tank. A non-return valve installed in

this piping prevents fuel in the main fuel system from flowing back into the wingtip tank. The non-

return valve opens only under the pressure generated by the wingtip tank fuel pump.


In the fuel distribution system there is a possibility to pump up the fuel from the drain sump back

to the fuel bypass branch begoud fuel cleaner.

Automatic flaps are installed in the outer filler to prevent fuel sucking out.

Operation

The fuel pumps are activated by switching on the FUEL PUMP LH (RH) circuit breaker.

The fuel pump delivers substantially more fuel than consumed by the engine (to ensure safe

function of the fuel pump). Excess fuel is returned through a return piping into the collector fuel

tank. Dirt is trapped in the fuel filter where fuel is partially dearated, the air being trapped in the top - - - - - - - - -

~

space of the fuel filter.

A return branch leads from the fuel filter top side carrying excessive fuel back to the middle fuel

tank. The ID 10 non-return valve in the return branch prevents the fuel system from sucking air

(e.g. in case of fuel pump failure). If the fuel filter cartridge is heavily contaminated, fuel can flow

to the engine through the fuel filter by-pass branch after opening a pressure-relief valve.

Before entering the fuel filter fuel is pre-heated to the required temperature in an oil-to-fuel heater.

By closing the fuel cock the flow of fuel into the fire zone is completely off.

The ID 16 non-return valves in the main piping prevent fuel from being pumped from the left hand

fuel tanks into the right hand fuel tanks and vice versa when both sides of the fuel system are

interconnected by the solenoid valves. The solenoid valves are actuated by switching on the

FUEL CROSSFEED circuit breaker.

When the fuel reserve in the main fuel system drops to 400 kg or below, the level switch switches - - - - - - - - - - - - - - - - - - - -

on the fuel pump in the wingtip tanks (with the WlNGTlP TANK LH, RH circuit breaker switched

on). If the wingtip tank fuel pump circuit is not switched on automatically (due to a failure) and fuel

is further drawn from the main fuel system, the second circuit of the level switch will switch on the

yellow ACTUATE TRANSFER signalling cell when the fuel reserve has dropped to not less than

220 kg. By switching on the WlNGTlP TANK FUEL TRANSFER switch the wingtip tank fuel pump

is activated (if WlNGTlP TANK LH, RH circuit breakers are switched on).

If the wingtip tank pump circuit has been switched on automatically, the pump will be switched off

automatically after fuel has been pumped off. If wingtip fuel pumps have been actuated over the

WlNGTlP TANK FUEL TRANSFER switch, the circuit must be switched off manually after the fuel

transfer has been completed.

A type B 073 420 N delay circuit built into the fuel transfer circuit prevents the ))starts<< of fuel

- - - - - - tr-ansfer during short-period (not exceeding 5 seconds) n e g a t i v e ~ l ~ a d ~ (gusts).

7.3. FUEL TANK VENTING The collector, middle, outer and additional fuel tank are interconnected by a venting piping. The

collector and the additional fuel tanks are connected by a piping to the atmosphere. The piping is

routed to the bottom skin of the wing between ribs No. 14 and 15. The piping end is cut off

obliquely and a tear-shaped deflector is installed before the piping outlet. The deflector prevents

icing. The oblique end and the ball valves installed in the fuel tank cap necks prevent fuel from

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, E20 - -

being sucked out from the fuel tanks. Four slots arranged in the conical sealing surface of each

ball valve prevent a raise of fuel pressure inside the fuel tinks. Expansion spaces in the wingtip

fuel tank and the additional fuel tank are interconnected by a piping. A solenoid valve installed in

the piping is open only if the fuel pump in the wingtip fuel tank is operating. When the fuel pump is

off, the valve is closed and it prevents undesirable fuel overflow. The wingtip fuel tank is'vented to

the atmosphere by a separate piping with a ball valve.

7.4. FUEL TANK INTERCONNECTION The collector, middle, outer and additional fuel tank are interconnected by a large diameter piping.

The outer fuel tank is connected to the middle fuel tank and to the additional fuel tank by a twin

piping, the middle fuel tank is connected to the collector fuel tank by triple piping. This ensures a

safe flow of fuel if the aircraft is steeply descending or climbing. The middle fuel tank is connected

to the collector fuel tank by special couplings. The top coupling is a straight feed-through type, the

bottom couplings have a non-return valve installed. The non-return valve prevents fuel overflow

from flowing into the middle fuel tank when the aircraft is banking. The fuel tanks are otherwise

interconnected by tubes fitted into rubber necks and secured with clamps.

The wingtip fuel tank is connected to the additional fuel tank by a ID 10 piping.

7.5. FUEL DRAINAGE Fuel is drained through the B 066 095 N drain valves and through the 1703 A drain necks.Two

drain valves are installed symmetrically on each side. In the area of the wing-to-fuselage fairing

the valves are connected to the piping from the collector tank sump.

The drain valves are accessible after removing the correspon corresponding access hole covers.

Fuel from the wingtip tank is drained through the drain valve installed in the bottom of the wingtip

fuel tank.

NOTE: Fuel from the wingtip fuel tanks can be drained more quickly through the main fuel tank

system.

7.6. FUEL INDICATING The aircraft has two indicating systems:

- a system indicating fuel quantity,

- a system indicating fuel pressure.

- fuel flow rate measurement (if installed)

Fuel quantity indication

Fuel quantity is indicated by a system of capacitive fuel gauges with emergency fuel reserve

signalling. The aircraft has two sets of fuel gauges. One set serves the right hand group of fuel

tanks, the other set the left hand group of fuel tanks. The emergency reserve signalling system is

installed in the LUN 1635-8, LUN 1636.01-8 fuel gauge transmitters. The set of the LUN 1649-8

fuel gauge transmitter and LUN 1674-8 indicator serve for indicating the fuel quantity in wingtip

tanks.

SOLENOID VALVE

FIG. 7-1 FUEL DISTRIBUTION SYSTEM DIAGRAM

LOCATED I

NON.RETURN VALVE (16LUN 7581.02-8)

WING TIP TANK IEL TRANSFER

8' : "1 NON-RETURN VALVE (IOLUN 7581.04-0)

NON.RETURN VALVE (10LUN 7581.04-8)


The fol lowing signalling systems are installed in addition to the emergency fuel reserve signalling:

- fuel crossfeed,

- fuel transfer from wingtip tanks,

- fuel quality drop to less then 220 kg

opening of the bypass valve on the fuel filter.

Fuel pressure indication

Fuel pressure is indicated'by two pressure gauges. Each side (the left hand and the right hand

one) has an independent system. One system indicates permanently fuel pressure at the outlet of - - - - - - - - - - - -

- - - - - -

the fuel control "nit supplying fuel ti fuel nozzles. It consists of the LUN 1559-8 fuel pressure

transmitter and the LUN 1538.01 three-pointer indicator. The second system only signals fuel

pressure drop below a permitted value at the fuel filter outlet.

Fuel flow rate measurement (if installed)

Fuel flow rate measurement is performed by two measurement systems. One is for the left engine

and the other for the right one. Fuel leaded into the engines is indicated on the elevant, fuel flow

rate indicators (LH, RH) into which an electric signal from the turbine flow transmitter is brought.

The turbine flow transmitter contains a turbine rotor which rotates at a speed proportional to the

volume of fuel (passing through it). The turbine rotor produces a sinusoidal voltage in a magnetic

pickup coil installed in the body of the transmitter. This voltage is directly proportiond to the

volume of fuel flow.

- - - - - - -

Fuel indicating operation

The quantity of fuel in wing tanks in indicated by the LUN 1634.01 -8 fuel gauge indicator

(separately for each group of fuel tanks). The signal is obtained from a capacitive transmitter.

When the fuel quantity in a fuel tank group (left hand or right hand) drops to 108 kg, the fuel

gauge set will signal the minimum fuel reserve by lighting up the MINIMUM FUEL signalling cell

on the central warning display. Fuel quantity in wingtip tanks is indicated on the LUN 1674-8 fuel

gauge indicator (sep arately for each wingtip fuel tank). The signal is obtained from the

LUN 1649-8 fuel gauge transmitter.

The FUEL CROSSFEED signalling cell will light up when both fuel tank groups have been

interconnected. The cell is activated by the switching of LUN 2474.3-8 solenoid valve. Fuel

transfer from wingtip fuel tanks is signalled by the illumination of the FUEL TRANSFER signalling - - - - - - -

celh TheceH isaetivated by the 0.012-K tUN t492.01-8 pressure switch:

If the quantity of fuel in one group of wing tanks, after a failure of automatic fuel transfer from the

wingtip tanks, drops to avalue of 220 kg (but not less), the ACTUATE TRANSFER signalling cell

illuminates on the central warning display.

The cell is activated by the S-2 level switch. The opening of the fuel filter bypass valve is signalled

by the illumination of the FUEL BYPASS signalling cell. The cell is activated by a built-in fuel filter

signallizer. Fuel pressure (continuous fuel pressure indicating) is indicated on the LUN 1538.01


three-pointer indicator, for each group of fuel tanks separately. The signal is derived from the

LUN 1559-8 fuel pressure transmitter. When fuel pressure drops below the permissible value, this

is indicated by the illumination of the FUEL PRESSURE signalling cell. The signal to the cell is

transmitted by the 0.03 K LUN 1492.01-8 pressure switch. Fuel flow rate is indicated on the fuel

rate indicators (if installed). Signal is received from a turbine flow transmitter.

7.7. OIL SYSTEM - GENERAL The oil system is used to continuously supply oil at a specified temperature for the engines, to

cool and lubricate bearings, to lubricate gears of the reductor and other transmissions, to supply

oil for the propeller blade angle setting hydraulic system and for the torgue transmitter.The

complete oil system is situated at the engine compartment, in front of the fire bulkhead, and is

divided into an airframe and an engine oil systems.

Oil type: B -3V dle TU 38-101295-72

Aeroshell Turbine Oil 500 according to MIL-L23699C

Aeroshell Turbine Oil 555 according to MIL-L23699C

Mobile Jet Oil according to MIL-L23699C

The airframe oil subsystem consists of an oil cooler , a fuel heater auxiliary pump, connecting

hoses, connecting and airbleeding pipings, and indications (the oil tank is an integral part of the

engine).The oil tank can be filled and the oil level checked after tilting up the lid on the engine

nacelle cover. Oil is drained by opening the engine nacelle oil tub and unscrewing the draining

plug on the bottom side of the oil tank and reducer.The total oil tank volume is 11 I, the oil filling is

only about 7 1.

NOTE: When draining the oil completely, it is also necessary to screw off the screwed fitting on

the bottom side o the oil cooler which is the lowest point of the system.

The engine oil system is described in the DM 601 Engine Maintenance Manual((

Oil system description

Oil is conveyed from the oil tank using pumps through an oil filter, to functional parts of the engine.

The oil pressure is scanned behind the oil filter and there is also a minimum oil pressure

signalization unit. The oil is returned from the engine through a fuel heater by means of exhaust

pumps. The fuel heater function is controlled by a thermoregulator which directs the oil either to

the heating coil of the fuel heater, or directly into the cooler, bypassing the heater.. Its function is

controlled by a thermostatic overpressure valve. The oil is fed from the oil cooler back to the tank.

In the bottom part of the oil tank, an oil temperature transmitter and a low oil level trasmitter are

installed. Behind the oil cooler, at the air outler of the bottom tilting cover of the engine nacelle,

there is a ice flap adjustable in two positions, for surnmer and winter operation, respectively. The

ice flap control is described In Chapter 5. A resettable flap is built into the oil cooler air In let, which

should protect the oil cooler against fire.Two litres of oil remain in the engine oil tank for propeller

feathering using the LUN 7840-8 auxiliary feathering pump. This quantity remains in the system


even if there is a leakage of oil due to a ruptured pipe.

Oil system operation

The operation of the airframe oil subsystem consists in the oil being returned from the engine

goes through a connecting hose into the fuel heater. From there. the oil is fed to the oil cooler and,

after cooling returned to the oil tank by a connecting pipe and a connecting hose.

FIG. 7-2 INDICATING SYSTEM - SCHEMATIC GRAVITY FUELLING

7-1 0


1-.-I -.-. L. r c i t m ~ ~ WARNING OISPUY 1 i LH ENGINE AIRFRAME Rti ENGlNEl

FIG. 7-3 SCHEMATIC DRAWING OF THE OIL SYSTEM

(1) 443 512 506 705 Oil cooler, (2) TPV 1 (7) Thermostatic and overpressure

valve, (3) LUN 1558-8 011 pressure transmitter, (4) 1.25 K LUN 1469.32-8

Pressure - change - over switch a part of the engine, (5) LUN 7782.01 Tank

closure, (6) Cover of the high-pressure oil filter, (7) Drain valve of the oil tank,

(8) Oil gauge, (9) 443 958 219 001 Oil-to-fuel heater, (10) B 560 454 N

Thermoregulator. ~ - - - - ~ ~ ~ ~ ~

- - - - - - - - - - - - - - - - pressure engine oil

exhausted oil --..----- - - - - - - - air bleeding

-- - --- oil drain

a - fire bulkhead

b - drive distribut~on box of auxiliary aggregates

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9. E20

FIG. 7-4 POSITIONS OF THE FLAP BEHIND THE OIL COOLER

A - Flap position with the closed deicing system and at low ambient

temperatures

B - Flap position with the closed deicing system and at high ambient

temperatures

C - Flap position with the opened deicing system and at low ambient

temperatures

D - Flap position with the opened deicing system and at high ambient

temperatures


7.8. OIL SYSTEM INDICATING The airplane is equipped by the following indication systems:

- oil pressure measuring system

- oil temperature measuring system

- minimum oil pressure signalization unit

- minimum oil quantity signalization unit

- system signalling the presence of in chips the oil system

The oil pressure is measured by two pressure-gauge systems, one performing continuous oil

pressure measurements, the other signalling oil pressure drops below the minimum oil pressure.

The former system measur es the oil pressure behind the oil filter. It consists of the LUN 1558-8

oil pressure transmitter and the LUN 1538.01-8 three-pointer indicator.

If the oil pressure in the pressure piping drops below the minimum limit, the signal cell

(OIL PRESSURE) on the signalization unit lights up. The signal is sent out by the

1.25 K LUN 1469.32-8 pressure change-over switch.

The oil temperature is measured in the bottom part of the oil tank, i.e, after the oil has passed

through the cooler. The temperature is picked up by the LUN 1358-8 oil temperature transmitter

and its value is displayed on the LUN 1538.01-8 three-pointer indicator.

The presence of drips in the oil is detected in the drive casing and the engine reductor it is

signalled by the CHIPS signal cell on the signalization unit lighting up. The signal is sent out by

appropriate K 601-577 and P 601-634 the transmitters of the chips indication.

The minimum oil quantity signalization system allows checking the engines whether they contain

a greater than minimum amount of oil. The checking should be performed before starting the

engines.The minimum oil quantity signal transmitter IS placed ins~de the engine oil tank. During

the check, it is turned on by the OIL LEVEL pushbutton on the left-hand control panel. The

minimum oil quantity signal is received and indicated by the OIL signal lights on the left-hand and

the right-hand control panels.

Oil system operation

The oil pressure and temperature monitormg and indicat~on systern is activated by turning on the

BATTERY I, II and 36 V - I, II INVERTERS switches. T l ~ e triple Indicator then shows the oil

pressure and temperature.

The minimum pressure and chips indicator system are activated by turning on the CENTRAL

WARNING DISPLAY AIRFRAME, CENTRAL WARNING DISPLAY-ENGINE LH and CENTRAL

WARNING DISPLAY-ENGINE RH circuit breakers. The minimum oil pressure is indicated by the

OIL PRESSURE signal cell lighting up. The presence of chips in the oil tank is indicated by the

CHIPS signal cell lighting up.

The signal cells are checked for their correct function by pressing down the CENTRAL WARNING

DISPLAY (LH, RH) pushbuttons. The minimum oil quantity signalization is used for a quick check

of the amount of oil in the oil tank. The check is performed in the pilots' cabin before starting the

AIRCRAFT TRAINING MANUAL L 41 0 UVP-El E9, E20

engines. Before accomplishing the check, it is necessary to turn on the BATTERY I, II

switches on the overhead panel, and the OIL LEVEL CHECK switch on the left-hand

control panel. If the oil level has dropped down to its minimum permissible limit or lower,

the yellow OIL signal lamp of the left-or the right-hand control panels lights up. The

lighting up impulse is sent out by the minimum oil quantity signalization unit placed inside

the engine oil tank. The correct function of the OIL signal lamp is checked by pressing

down the SIGN. pushbutton on the test panel. After completing the check, the OIL LEVEL

CHECK switch must be turned off.


.-.

DETECTOR

0 1 L TEMPERATURE. TKANSMI ITER

MINIMUM OIL PRI~SSU11L!

FIG. 7-5 SCHEMATIC DRAWING OF THE INDICATION SYSTEM


CENTRAL WARNING DISPLAY i:NC;INI< 111 ENGINE RH

CHIPS

INVERTERS .

1 36-1 I I1

RELAY




CHAPTER 8 FIRE PROTECTION

8.1 General

8 2 Detect~on

8 3 Ext~ngu~sh~ng

8 4 Fire protection of AC generator cooling alr intake




8.1. GENERAL The aircraft fire-fighting equipment consists of:

- a fire detection system in engine nacelles and in front and rear baggage compartments

- a fire extinguishing system in engine nacelles, in front and rear baggage compartments and in

aircraft cabin

- an engine nacelle drainage system, a fuel drainage system of cornbustion chambers and fuel

system components and an oil tank drainage system.

As a means of fire protection, the engine nacelle is divided into separe fire zories and separated

from the airframe structure by means of firewalls.

8.2. DETECTION The engine nacelle fire detection system consists of fire detectors, and actuating unit and light and

acoustic signalling devices. The fire detectors are situated in the engine nacelle. They are

connected in series, 3 detectors in each of the three signalling circuits. Fire in the engine nacelle

actuates the red FlRE flashing signal accompanied by ringing of the signalling bell.

The front baggage compartment fire (smoke) detection system consists of a smoke detector and

a light signalling device. The smoke detector is located in the upper part of the front baggage \

compartment near the bulkhead No. 4 and protected from mechanical damage by wire-welded

guard. Fire in the front baggage compartment actuates the red BAG. COMP. FlRE signal.

The aft baggage compartment fire (smoke) detection system consist of a smoke detector and a

light signalling device. The smoke detector is located in the upper part of the aft baggage

compartment and is protected by wire-welded cover against mechanical damage. In case of fire in

the aft baggage compartment the red signal cell FlRE AFT BAG. COMP is actuated.

On the aircraft is installed as possible of the rear baggage compartment fire extinguishing system

and fire detection system and fire detection system (if installed rear baggage compartment).

Detection system operation

In case of temperature rise in the engine nacelle the DPS fire detectors send pulses which are

processed in the BI-2A, Series 2 actuating unit. Having processed the pulse, the operating unit

actuates a signalling bell and the FlRE signal (of left-hand or right-hand engine, provided the

CENTRAL WARNING DISPLAY - LH ENGINE,RH ENGINE circuit breaker is switched ON).

Check of correct functioning is carried out with help of the switch CENTRAL WARNING DISPLAY

- ENGINE (LH), CENTRAL WARNING DISPLAY - ENGINE (RH) pushbuttons. Fire detectors are

checked for correct functioning with help of ENG. FlRE SIGN. I, 11, Ill pushbuttons. When

depressing the above-mentioned pushbuttons, the signalling bell and the FlRE signal (of right-

hand or left-hand engine) must be actuated if the fire detectors function correctly.

If smoke in the front (rear - if installed) baggage compartment is generated, it passes through the

DS-3M2 smoke detector. The passing smoke shades the light beam inside the smoke detector


BELL

FIG. 8-1 ENGINE NACELLE FIRE DETECTION SYSTEM


FIRE ZONES

\ SIMPLlf-I: D WIRING


which is generated by a filament lamp and directed onto a photocell. The signal from the photocell

is amplified and led onto the BAG. COMP. FlRE or FlRE AFT BAG. COMP (if installed) signal

lamp which lights up. Signalling cell correct functioning of the signal lamp is checked with help of

the CENTRAL WARNING DISPLAY - AIRFRAME pushbutton. The smoke detector is checked

for correct functioning with help of the FlRE DET. CHECK (front baggage compartment) or

DET. CHECK (rear baggage compartment - if installed) pushbutton.

CENTRAL WARNING DISPLAY BAG. COMP. FIRE DET. AIRFRAME FRONT AFT.

WARNING DISPLAY AIRFRAME

. - - - - - -

SMOKE D E T E C T O R

FIG. 8-2 FRONT AND AFT BAGGAGE COMPARTMENT FlRE DETECTION SYSTEM


8.3. EXTINGUISHING The engine fire extinguishing system consists of fire extinguishing rings, fire extinguishers,

distribution piping and electrical control and check system. The fire extinguishers placed on the

back side of the left-hand and right-hand rear fire walls of engine nacelles are interconnected by

two pipes. By means of one of them the distribution piping of the left-hand engine nacelle is fed,

by the second one the distribution piping of the right-hand engine nacelle is fed. The distribution

pipings and fire extinguishing rings are of different cross sections depending on the rate of flow of

the fire extinguishing agent.

Each fire extinguisher is fitted with two discharge valves (for the left-hand and right-hand engine

nacelle) controlling the fire extinguishing agent inlet into the distribution piping. The pressure in the

fire extinguishers is checked with help of pressure gauges after opening the lids on the rear

engine nacelle cowling. Connected to the distribution piping in front of the fire wall is a branch

pipeline leading to the cylinder controlling the fire flap in the AC generator cooling air intake.

The flap in the oil cooler air inlet channel is fitted with a fuse which closes air inlet into the cooler in

case of fire.

The fire extinguishing system of the front baggage compartment consists of one fire extinguishing

piping, one fire extinguisher and of a control and check system.

The fire extinguisher is provided with, a mechanical valve controlled by a handle located on the

right-hand control panel and a flexible pull rod. The fire extinguisher is situated in front of the

bulkhead No. 4 under the front baggage compartment floor near the nose wheel bay.

The fire extinguishing agent is led through the distribution piping into the fire extinguishing tube

located under the ceiling of the front baggage compartment. The fire extinguisher is fitted with a

safety valve set to a pressure of 1.45 to 1.55 MPa and connected by a vent pipe with the

atmosphere. Pressure in the fire extinguisher is checked with help of a pressure gauge situated

on the instrument panel.

The fire extinguishing system of the rear baggage (if installed) compartment consist of fire

extinguishing rings, fire extinguisher and system of electrical control. The fire extinguisher is

located behind the aft wall of the aft baggage compartment and is connected by piping with the

extinguishing ring. The pressure in the extinguisher can be checked by a gause located on the

extinguisher. Fire extinguishing in the pilot's cabin is secured by two FH 15N-364 fire

extinguishers one of which is situated on the floor behind the co pilots seat and the second one on

the left-hand side near the main door in the passenger compartment. Both fire extinguishers are

controlled manually. Fire in the engine nacelle of an parking aeroplane can be extinguished

through an access hole in the engine nacelle whose lid breaks upon impact of the nozzle of

airfield fire fighting equipment.

m: AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, E20

BAT TERY

FIG. 8-3 ENGINE FIRE EXTINGUISHING SYSTEM


FlRE ZONES

OPERATING CYLINDER /

1

FlRE EXT BOTTLE

OPERATING CYLINDER 'I

NON - RilXJRh: VALVE.


PRESSURE GAUGE

FlRE EXTINGUISHER

FIG. 8-4 FRONT AND AFT BAGGAGE COMPARTMENT FlRE EXTINGUISHING

SYSTEM


Fire extinguishing system operation

The engine fire extinguishing system is controlled electrically by means of four pushbuttons: FlRE

EXTING. PRIM. (L.H. and R.H.) and FlRE EXTING. SEC. (L.H. and R.H.). The FlRE EXTING.

PRIM. pushbuttons are intended for switching the first (primary) fire extinguishing circuit and the

FlRE EXTING. SEC. pushbuttons for switching the second (secbndary) fire extinguishing circuit of

the left-hand and right-hand fire engine extinguishing system.

By depressing the pushbutton , the corresponding pyrocartridges are activated. The pressure of

gas releases the discharge valve which opens by the internal pressure in the fire extinguisher and

lets the fire extinguishing agent flow into the distribution piping and through the holes of the fire

extinguishing rings to the place of fire.

The FlRE EXTING. PRIM. (L.H., R.H.) and the FlRE EXTING. SEC. (L.H., R.H.) pushbuttons are

fitted with caps protecting them from unintended switching. The caps are sealed with easily

removable seals. The engine fire extinguishing circuits are connected to the battery terminals by

the FlRE EXT. L.H. and FlRE EXT. R.H. circuit breakers located in the battery bay. Such a

connection enables fire extinguishing system activation even if all other electrical power sources

are disconnected from the aeroplane electrical system.

The fire extinguishing system in the front baggage compartment is controlled mechanically. The

system is actuated by a handle located on the right-hand control panel. The handle must be pulled

in. By releasing the handle, fire extinguishing may be stopped or interrupted.

The aft baggage compartment (if installed) extinguishing system is controlled electrically. It is

controlled by FlRE EXT. AFT BAG. COMP. pushbutton. After pushing the pushbutton the

pyrocartridge explodes and releases the fire extinguisher valve. The valve is opened by the

overpressure from the extinguisher and extinguishing agent flows trough the extinguishing ring

holes to the fire area. The aft baggage compartment fire extinguishing circuid is connected to the

battery terminal. For extinguishing fire the FH 15-364 hand-held fire . inguishers are used.

8.4. FlRE PROTECTION OF AC GENERATOR COOLING AIR INTAKE The fire protection of the AC generator cooling air intake consists of a flap closing the AC

generator cooling air duct in case of fire.

The fire protection system of the AC generator cooling air intake is assembled of a pipeline, an

operating cylinder and a closing flap.

When the engine fire extinguishing system is actuated, the pressure of the fire extinguishing agent

shifts the operating cylinder piston which closes the flap in the AC generator cooling air intake.


FIG 8-5 FIRE PROTECTION OF AC GENERATOR COOLING AIR INTAKE

(1) Pipeline, (2) Cylinder, (3) Flap, (4) Nut, (5) AC generator coollmg air ~ntake,

(6) Cylinder body, (7) Piston, (8) Cap nut, (9) Spring, (10) Ring, (1 1) Packing

ring.

Cutaway view of cylinder, ~ t e m (2)


CHAPTER 9

ELECTRICAL POWER

9.1. General

9.2. AC generation

9.3. DC current sources

9.4. External power

9.5. Electrical load distribution

9.6 Electric system of airplane systems

9.7. Lighting system

9.8. Flight compartment lighting

9.9. Exterior lighting

9 10 Emergency lighting




9.1. GENERAL The main sources of the electrical power for the aeroplane are two starter-generators having

voltage of 28 V, each having the power output of 5.6 kW. As an emergency sources are exploated

two NiCd batteries having voltage of 24 V and each capacity of 25 Ah.

Electrical appliances for alternating current are supplied with the power from two LUN 2450 static

inverters having voltage 3x36 V1400 Hz, two PC-250 or LUN 2460 static iverters having voltage

11 5 V1400 Hz and 26 V1400 Hz (one of which is a stand-by z inverter) and two LUN 21 02 or

LUN 2102.01 alternators having voltage 3x1 15 V1200 Hz. Inverters are supplied from the direct

current board network of 28 V.During ground servicing, the aeroplane can be connected to a DC

external power source of 27 to 29 V.

9.2. AC GENERATION The aeroplane source of alternating power supply are two LUN 2450 static inverters having

3x36VI400 Hz, two PC-250 or LUN 2460 static inverters having 115 Vl400 Hz and 26 V/ 400 Hz

(one of them serving as a stand-by), two LUN 2102 or LUN 2102.01 alternators of voltage

3x1 15/200 V, 300 to 507 HZ AGR-74-5.

The suplying of devices from LUN 2450 inverters

At normal state the inverter I supplies the following devices (applies to the airplanes in which the

second inverter installed is Model PC 250):

- engine devices of the LI i engine

- fuel gauge LH

- pressure gauge LH

- wing tip tank gauge LH

- gyro compass II

- icing indicator

Inverter 11 suppllies the following devices:

- engine devices of the RH engine

- fuel gauge RH

- pressure gauge RH

- wing tip tank gauge RH

- hydraulic pressure gauge

- turn indicator l

The supplying of devices from LUN 2450 inverters

At normal state the inverter I supplies the following devices (applies to the airplanes in which the

second inverter is Model LUN 2460):

- horizon LH (through the transformer)

- engine devices of the L t i engine

- gyro compass I I

- turn indicator II

- icing indicator

Inverter II supplies the follow~ng devices:

- engine devices of the RH engine

- fuel gauge RH

- turn indicator I

The supplying of devices from the PC-250 inverters (if installed)

At the normal state the inverter 1 supplies the following devices:

- turn indicators ( I I 5 V--)

- horizon RH (1 15 V-)

- gyro compass 1 , 1 1 (1 15 V-)

- NAV 1 (36 V-)

- ADF I, 11 (26 V-)

- weather radar (1 15 V--) - if installed

- vertical gyro (1 15 V-) - if installed

In case of damage of inverter I these devices are supplied by inverter II

The suplying of devices from the LUN 2460 inverters (if installed)

At the normal state the inverter I supplies the following devices

- turn indicators

- horizon RH

- gyro compass I

In case of damage of inverter I these devices are supplied by inverter II

The suplying of devices from the LUN 2456.02-8

This inverter supply standby horizont only

The suplying of instrument from the LUN 2120 or LUN 2102.01

At the normal state the windows heating circuit is supplied by the RH engine alternator and the

propellers deicing circuit by the LH engine alternator

In case of damage one of alternators the windows heating and the propellers de-icing circu~t are

automatically connected to the working alternator

AC generation system operation

LUN 2450 inverters switchmg on ( ~ f ~nstalled)

After the switches BATTERY I, II and the c~rcuit breaker CENTRAL WARNING DISPLAY -

-ELECTRO are switched on, then, among other, s~gnalling cells INVERTER 1 36 V and


-ELECTRO are switched on, then, among other, signalling cells INVERTER 136 V and

INVERTER 11 36 V on the signal panel are also lighted on. By switching on the switch

INVERTER 36 V I at the overhead panel the signal cell INVERTER 1 36 V will extinguish and all

electric devices 3x36 V/400 H z are supplied by the inverter I.

By switching on the switch INVERTER 36 V II at the overhead panel the signal cell

INVERTER 11 36 V on the signal panel will extinguish and all electric devices connected to the

inverter II are supplied by voltage of value 3x36 V/400 Hz from this inverter.

In case of damage one of inverters the corresponding cell and consumpers on the central warning

display lights on and the electric devices are automatically connected to second working inverter.

The individual electrical consumpers are connected to the inverters voltage through the

distributing box RS-6. In the distributing box are mounted circuits for the inverters damage

signalling and for automatic change-over switching of electric appliances to the working inverter in

case of damage one of them.

Check of the individual inverters phases voltage 3x36 V/400 H z is provided by means of the

change-over switch of the inverters 36 V - voltmeter on the right control panel and by the voltmeter

VF 0.4-45.

NOTE: There must be switched on circuit breakers INVERTER I, 11 36 V between 7 and 8 frame

on the left side, to secure the inverters supplying.

PC-250 inverters switchin on (if installed)

When the switches BATTERY I, II and the circuit breaker CENTRAL WARNING DISPLAY-

ELECTRO are switched on, then among other, signal cells INVERTER 1 115 V and

INVERTER 11 115 V on the central warning display are also lighted on. By switching on the

switches INVERTERS I, 11 115 V at the overhead panel the signal cells INVERTER 1 115 V and

INVERTER 11 115 V on the central warning display will extinguish; the devices are supplied by

inverter I. Inverter II then is shut off and stand-by (out of function). In case of a breakdown of

inverter I the signal cell INVERTER 1 115 V on the signal panel will light on and inverter II will be

automatically switched on and supplies all electric consumpers 11 5 V1400 H z (26 V/400 Hz).

Voltage 115 V/400 Hz from inverter is supplied to a relay change-over switch RP-5 (8 ) , with built-in

signall circuits serving to breakdown signalling and automatic switch over from oue inverter under

breakdown to the functioning one. From the relay change-over switch the volt age is supplied

through the distribution box to the individual electric consumpers. Voltage 26 Vl400 Hz from

invertors is supplied to the A 131 relay which is switched over as per operating invertor.

Consumpers 26 V/400 Hz are connected to this relay. The change-over switch INVERTER

SELECT 115 V in position AUT. serves to select inverters switching on.

- in posit~on AUT - inverter I is functional, in case of its breakdown, inverter II is made functional

and the devices are automatically switched over to inverter II. Inverter

breakdown is signalled on the central warning display,

- in position

- in position


- inverter I is functional, in case of its breakdown is signalled on the central warning

display.

I - inverter II is functional, in case of its breakdown the AC devices are out of function.

The breakdown is signalled on the central warning display.

NOTE: The normal position of the switch is AUT. In case of breakdown in the automatics, the

pilot selects another position of the switch after pulling up the fuse.

Voltage checking of inverters 11 5 V/400 Hz is performed by means of inverters 1 15 V and

alternators voltmeter switch (on the right control panel) and by the voltmeter VF 0.4-150 on the

right side of instrument panel.

NOTE: To secure the inverters supplying, there must be switched on the circuit breakers of

INVERTER 115 V I, II between the 7 and 8 frame.

LUN 2102 or LUN 2102.01 alternators switching on

When the switches BATTERY I, II and the circuit breaker CENTRAL WARNING DISPLAY-

ELECTRO are switched on and the engines are running, then, among other, signal cells

ALTERNATOR-LH and ALTERNATOR-RH on the central warning display are also lighted on. By

switching on the switch ALTERNATOR-LH the signal cell ALTERNATOR-LH on the signal panel

will extinguish and the windows heating circuit is supplied by voltage 1 15/200 V. By switching on

the switch ALTERNATOR-RH the signal cell ALTERNATOR-RH on the central warning display

will extinguish and the propellers de-icing circuit is supplied by voltage 11 51200 V.

The voltage from alternators is supplied to the de-icing systems through the distribution box RS-4.

In the distribution box are mounted the alternators breakdown signalling circuit and the circuit of

the automatic switch-over the de-icing system to working alternator.

Voltage check of the alternators phases is performed by the voltmeter change-over switch of the

alternator and of the inverters 115 V (on the right control panel and by the voltmeter VF 0.4-1 50

on the right side of instrument panel).

9.3. DC CURRENT SOURCES

The sources of DC voltage in the aeroplane are two startergenerators.

As emergency source of DC current serve 4 four alcaline NiCd accumulator batteries.

Accumulator battery space temperature is measured and indicated by temperature pointer on the

left control panel.

The starter-generators can be used either as a generator of as a starter.

left starter-generator circuit

right starter-generator circuit

battery circuit and external sources

Source of DC voltage system operation


Operation of generators circuit

When the left (right) engine is running, generator A1 (A2) is connected to the aeroplanes network

by switching on of the breaker GENERATOR LH (RH) A7 (A8). By switching on the breaker,

voltage is supplied from the generator through circuit breaker A 63 (A 64) to the contactor coil of

differential relay A 5 (A 6) which will connect the generator to the aeroplane network in case when

the voltage in the electrical circuit of the generator is about 0.3 to 0.7 V higher than voltage in the

aeroplane network (battery).

Derived from the moveable contact of differential relay contactor is signalling of generator shut-off

from the aeroplane network using the relay A 21 (A 22) and signal cell GENERATOR LH (RH) on

the warning display. After connecting a generator to aeroplane electrical network, signal cell

GENERATOR LH (RH) will extinguish.

Generator voltage is maintained within the required tolerances by a semicounductor voltage

regulator A 3 (A 4) under various speeds of the engine and under different loads in the aeroplane

electrical network.

If voltage exceeds the value of 31 V (in case of defective voltage regulation), the generator will be

automatically shut-off from the aeroplane electrical network by a protective circuit located in the

voltage regulator A 3 (A 4), which will activate the switch GENERATOR LH (RH) A 7 (A 8). This

will cause a disconnection of differential relay A 5 (A 6) contactor.

When voltage of the generator A 1 (A 2) will drop below the limit of aeroplane electrical network

voltage the current coil of the differential relay A 7 (A 8) will disconnect the generator from the

aeroplane electrical network with inverse current magnitude between 25 - 35 A.

Magnitude of voltage and current supplied by the generator can be checked by the voltammeter

A 29 (A 28) with the shunt A 41 (A 42). Measuring circuits of the voltammeter are protected by

cut-off fuses A 45, A 47 (A 46, A 48). Voltammeter A 28 is also used to check the voltage and

current supplied from aeroplane batteries. Selection of the measured mode is performed by a

switch-over VA METER A 27. When measuring voltage and current of the right side generator,

the switch-over A 27 is in position GEN RH.

The left and right generator circuits under normal operation conditions are independent of each

other, and are separated by means of a contactor A1 3 and circuit breaker A 12. When only one

generator operates, the contactor A 13 interconnects both generator circuits. Contactor A 13 is

controlled automatically, the control being derived from position of relay A 21 (A 22) contacts,

which are controlled from the moveable contact of differential relay A 5 (A 5) contactor. The

contactor is always supplied from battery of that generator circuit, from which the generator is cut-

off by differential relay A 5 (A 6). Dur~ng the time bus-bars are interconnected by contactor A 13,

functioning are circuits for parallel cooperation of generators, which are located in the voltage

regulator A 3 (A 4). The regulators for parallel cooperation (terminals 5 of both regulators) are

connected through the relay A 62 controlled by the con;actor A 13.

NOTE: Parallel cooperation of generators is functioning in case of second generator connection

to network only.


Operation of the batteries circuit

Each generator circuit under normal functional conditions incorporates one battery A 9 (A 10).

When both generator circuits are interconnected through the contactor A 13, both batteries can

operate parallely, provided one of the generators is operating.

The battery is connected to the aeroplane network by switches BATTERY I, I I , A 25 (A 26). By

switching on the contactor A 15 (A 16) its coil circuit will be closed and it will connect battery A 9

(A 10) through fuses A 73 (A 74), shunts A 4 0 (A 20) and circuit breakers A 67 (A 68), A 43

(A 44) to the aeroplane electrical network of generator A 1 (A 2).

Batteries signalling circuit signals in case when the battery is not connected to the aeroplane

electrical network. Signalling is derived from the contactor A 15 (A 16) and from the controlling

relay A 35 (A36), which connects by its contacts the supply circuit of signal cell BATTERY on the

central warning display.

The signal cell is common to both batteries, and the distinction as to which of the two batteries is

cut-off must be made by ammeter A 28 (deviation of pointer during consumption of charging).

Supply of signal cell is protected by cut-off fuse A 79.

Voltage and charging of discharging current can be checked by the voltammeter A 28 of a switch-

over VA METER A 27 is in position BAT I or BAT II. The checked parameters are recorded from

the shunt A 40 (A 20). The measuring circuit is protected by cut-off fuses A 53, A 69 (A 54, A 70).

Charging current of the battery is measured on the ammeter scale from ,,Ol< to the left and the

discharging current from jjO(< to the right.

In the regime of the engine starting-up from aeroplane batteries, both batteries will be parallely

interconnected by means of contactors A 77, A 78. These contactors are controlled by contacts of

relay A 65, A 66, which are activated during the engine start up, namely relay A 65 when the port

engine is starting up, and relay A 66 when the starboard eingine is being started up. Contactor

control circuits A 77 are protected by cut-off fuse A 80, contactor A 78 and by cut-off fuse A 79.

Operation of battery temperature measurement circuit (if installed)

On the aircraft is installed a accumulator battery temperature measuring device. The device is put

into operation by switching on the switching on the switches BATTERY I, II and circuit breaker

BAT. TEMP. on the oveihead panel.

The temperature measurement circuit inclued the circuit breaker AZRGK-2, BTI 600-2A

temperature indicator, push button KNR and P 600-48 temperature probe.

9.4. EXTERNAL POWER

The aeroplane can be connected to an external direct current source of voltage 28 V, which

serves to starting of drive units to supply and checking of the individual electric circuits when the

aeroplane is on the ground.

The external source is connected to the aeroplane through the external source connecting plug

located under a lid under the battery bay on the left hand side of the aeroplane nose.


Technical data: For devices checking

Voltage: 27 V - 29 V

Other requirements are indentical to standard GOST 19705-81

For engines starting

Maximum short-time overload: 700 - 800 A

The voltage must not fall under 14 V at any phase of starting.

(When the button ENGINE START UP is pressed the voltage drop to 14 V is

admissible max. 2. seconds)

NOTE: External source voltage is measured in position of V-A meter change-over switch NS I or

NS II.

Operation

The external source is connected to the aeroplane network through the external power plug A1 1.

After the plug is inserted the relay A 33, A 34 will change-over switch and will disconnect from the

aeroplane network:

- battery A 9, A 10 by contactors A 15 and A 16,

- generators A 2 (2 pieces) by means of differential relay A 5 , A 6 contactors.

After switches BATTERY I, I1 on the overhead panel have been put in the ON position, contactor

A 19 will connect the external source to aeroplane network. Connecting of external source is

signalled by signal cell EXTERNAL POWER SUPPLY in the warning display in the cockpit.

Control circuits of the automatic switching are secured by cut-off fuses A 30, A 39.The power

circuit of the external source is secured by a cut-off fuse A 51.

9.5. ELECTRICAL LOAD DISTRIBUTION The direct current aeroplane network is designed as being of a one conductor type. The positive

pole is distributed by insulated conductors, while the aeroplane frame represents the negative

pole. From the generator electric power is supplied through a differential relay to the individual

bus-bars of the left (right) generator circuit.

The circuit of the left generator A 1 is securing the supply to the following bus bars:

S 1 A secured by circuit breaker A 59

S 1 B secured by c i rc~~ i t breaker A 57


N S l secured by c~rcuit breakers A 43, A 67 and A 37

In the event of a supply voltage loss the bus-bars are automatically switched over:

S 1 A to battery A 10 (NS 2) circuit contactor A 24 through circuit breaker A 38 and through circuit

breakers A 68, A 44 to the circuit of generator A 2.


S 1 B to generator A 2 circuit through contactor A 49 and breaker A 52

The circuit of the right generator A 2 is securing the supply to the following bus bars

S 2 B secured by circuit breaker A 58

S 3 B secured circu~t breaker A 61


N S 2 secured by circuit breakers A 44, A 68 and A 38

In the event of a supply voltage loss the bus bars-are automatically switched-over as follows:

S 2 B a part of electric devices (supplied through contactor A 23) to the battery A 9 (NS 1) circuit

by contactor A 23 through circuit breaker A 37 and through circuit breaker A 67, A 43 to the

generator A 1 circuit, and a part of devices supplied via contactor A 32 and circuit breaker

A 38 to the battery A 10 circuit.

S 3 B to the generator A 1 circuit through contactor A 50 and circuit breaker A 55

Directly supplied from the battery A 9 bus-bar are:

- crash recorder circuit secured by circuit breaker K 16

- circuit of the left engine fireproof system secured by circuit breaker M 201

- position lighting circuit secured by fuse C 34

- stand-by horizon secured by circuit breaker L 52

NOTE: During the supplying from external power source is via fuse C 34 supplied IELU LH, KH

through circuit breakers M 189 and M 190, board intercommunication device through

circuit breaker FA 13 and transceiver VHF I through circuit breaker FN 1

Supplied directly from the battery bus bar A 10 are:

- circuit of the right engine fire system secured by circuit breaker M 202

- -circuit of the rear baggage compartment extinguishing secured by c~rcuit breaker M 207

- lighting of p~lot s cockp~t secured by circurt breaker M 207

9.6. ELECTRIC SYSTEM OF AIRPLANE SYSTEMS

All airplane systems are connected mostly to the a~rplane bus - bars and classifred ~n to the

followmg circuits

-circuit A - sources of electrrc energy

- circuit B - starting

- circuit C - lightmg

- circuit D - de-icing


- circuit E - controls

- circuit M - extinguishing and fire signallization, engine instruments

- circuit K - recorder

- circuit F,L - navigation, communication

Operation of the particular systems inclusive of their light signallization is mentioned in relevant

chapters of this L 410 UVP-E, E9, E20 Airplane training manual. Only sources and lighting are

mentioned in this chapter.

9.7. LIGHTING SYSTEM The lighting system consists of the internal and exteinal lighting of the aircraft, the central warning

dispay and the emergency lighting.The internal lighting consists of the flight compartment and rear

lighting, the passenger compartment lighting and the front baggage compartment lighting.

The external lighting consists of a search light position lights, anticollision beacons, searchlights

for lighting of the fin, the static ice detector lighting and vertical tail. (if installed)

No.

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

16.

17.

18.

19.

20.

21.

Name

-ighting fixture B 082 840 N

Compass lighting I LUN 2880.3-8

Type or drawing No.

Lighting fixture

Navigator's table lamp

I SM 16

B 082 614 P, L

LNS (8)

Lamp (Type)

Portable lamp

Outside air thermometer lighting

15

Lighting fixture B 082 401 N

Power (W)

1

PL (8)

-

Rear panel lamp

Instrument panel lighting

QtY 1

1 SM 16

Emergency instrument panel lighting

Light assembly

Lighting fixture

SM 14

SM 37

LZP (8) (B 580 593 N)

Lighting fixture

Lighting fixture

Lighting fixture '08-9340.86

Posit~on light BAN0 57 CHS 57

15

B 091 225 L, P

5

1.4

1

62 051

A-7 1 5- 1

B 590 670 N

B 091 345 N

LFSM 28 1 200+300 1 3 1

1

1

63 0030

SM 16 SM 14

Searchlight

Anticollision beacon

Static ice detector lighting

External emergency lighting

3

0.5

SM 30

SM 38-0.05-1

MSL-3

B 091 420 N

1

30

3

15 5

Emergency lighting panel

5

18 9

5

1.4

1

4

I B 091 462 N 67 038 2 4


No.

22.

9.8. FLIGHT COMPARTMENT LIGHTING - Flightcompartment - - - - - - -

23.

24.

25.

26.

The flight compartment lighting is a part of the internal aeroplane lighting. It consists of the cockpit

lighting, the left hand and right hand control panel lighting, the central console lighting, the

compass lighting, the outside air thermometer lighting, the chart holder lighting (if installed),

navigator' s table lamp and of the central warning display. The instruments installed on the

instrument panel either have a built-in illumination or they are lighted by lamps installed above the

instruments. Besides, an emergency instrument panel lighting is installed in the flight

compartment.

For the activation of the different lighting fixtures the BATTERY I, II switches on the overhead

panel must be switched on. Switch on the LIGHTING circuit breaker to supply voltage to the

LNS (8 ) navigator' s table lamp and to the LZP (8) rear panel lamp (B 580 593 N) and the lamp for

the chart holder lighting (if installed). The lamps are turned on and off with switches on eachlamp. - - - - - -

- - - - - - - - - - - - - - - -

- - - - - - -

- - -

switch on the COCKPIT circuit breaker to turn on the B 082 840 N cockpit lighting fixture.

Switch on the INSTRUMENT PANEUCIRCUIT I circuit breaker to turn on the instrument and

compass lighting. The lighting intensity can be controlled with the control knob on the

LUN 2412.01 instrument lighting control panel and with the RL-10 rheostat on the instrument

panel.

Switch on the INSTRUMENT PANEUCIRCUIT II circuit breaker to turn on the B 082 401 N

lighting panel illuminating the central console (front and central control panels) and the overhead

panel, the outside air thermometer lighting and the B 082 614 P/L lighting fixtures illuminating the

LH and the RH control panel. The illumination intensity of the LH, RH, central and front control

panels can be controlled with the RL-10 rheostat installed on the instrurnent panel. The

INSTRUMENT PANEUCIRCUIT II circuit breaker also actuates the lighting of instruments (supply

voltage of 28 V). The-illumination intensity can b-e controlled withlheRLlQrheostat installed OR - - - - - - - - - - - - - - -

the instrument panel.

Switch on the INSTRUMENT PANEL - STBY CIRCUIT circuit breaker to turn on the secondary

lighting of instruments. Switch on CENTRAL WARNING DISPLAY LH ENGINE,

AIRFRAME, ELECTRO and RH ENGINE circuit breakers to activate the four blocks of the central

warning display.

r

Name

Lamp for chart holder lighting .

Lighting fixture

Searchlight for lighting of the fin

Lighting fixture (if installed)

Lighting fixture (if installed)

Type or drawing No.

L 41 0.8292-04

B 590 741 N

B 571 527 N

B 091 579 N

B 091 581 N

Lamp (Type)

62051

BA-7s

SM 28-70

67 038

67 038

Power (W)

3

Q ~ Y

2

2

70

2

2

2

2

1

1


LIGHTING

FIG. 9-1 FLIGHT COMPARTMENT LIGHTING - SCHEMATIC

AIRCRAFT TRAINING MANUAL L 41 0 U V P - E , E9, E20

Passenger compartment

The passenger compartment lighting system includes the lighting of the passenger compartment

proper, the lighting of the passenger compartment rear section (the toilet) and the lighting of

transparencies.

Before switching on individual lighting fixtures switch on the BATTERY I, II switches on the

overhead panel. When the CABIN 113 circuit breaker is switched on, a single bulb

will light up in each light assembly in the passenger compartment except in the first light assembly

on t he left-hand side and in the last light assembly on the right-hand side of the passenger

compartment. Two bulbs in each light assembly will light up when the PASSENGER CABIN 213

circuit breaker is switched on. The toilet is lighted by switching on the switch on frame No. 19a

provided 'the LIGHTING circuit breaker on the overhead panel is switched on.

When the main door is opened, the convenience lighting will light up automatically (one bulb in the

first light assembly on the left-hand side and one bulb in the last light assembly on the right-hand

side of the passenger compartment provided the CONVENIENCE LIGHTING switch is switched

on. This switch permits the convenience lighting to be switch off with the main door open.

The FASTEN SEAT BELTS circuit breaker on the overhead panel lights up lamps in the

FASTEN SAFETY BELTS and RETURN TO YOUR PLACE transparencies on the vertical control

channel and on the passenger compartment ceiling near the toilet frame No. 19 respectively or

toilet - door (if installed). # the ,toilet door is opened and the LIGHTING circuit breaker switched

on, the WC OCCUIPED transparency on the wall of the rear baggage compartment will light up

automatically (if installed).

Cargo and service compartments

The rear baggage compartment (between frames 19th to 21st or 18th and 19a) has not lighting.

It is lighted by the light assemblies in the passenger compartment and by the light in rear part

passenger compartment. The rear baggage compartment (between frames No. 21st. and 22nd -

- if installed)

The front baggage compartment is lighted by two lighting fixtures installed on No. 4 bulkhead.

To switch on the lamps in the front and rear baggage compartments switch on the BATTERY I, II

switches and the LIGHTING circuit breaker on the overhead panel.

Switching on and off of lamps in the front baggage compartment is carried out by the switch

situated by the left lamp on the 4th bulkhead. Switching on and off of a lamp in the rear baggage

compartment is carried out by the switch situated near the door of the rear baggage compartment.


PASSENGER CABIN FASTEN SEAT

6 FASTEN SAFETY BELTS F

RETURN TO

YOUR PLACE

WC OCCUPIED + [ (if installed)

FIG. 9-2 PASSENGER COMPARTMENT LIGHTING - SCHEMATIC

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9. E20

FIG. 9-3 FRONT BAGGAGE COMPARTMENT LIGHTING - SCHEMATIC

9.9. EXTERIOR LIGHTING The external lighting of the aeroplane consists of anticollision beacons, position lights,

searchlights, searchlights for lighting of the fin, lamp illuminating the static ice detector, and

vertical tail plane lighting.

The external lighting is operative only if the BATTERY I, II switches on the overhead panel are

turned on. When the SEARCHLIGHTS I change-over switch on the overhead panel is turned on,

the central searchlight lights up. When the SEARCHLIGHTS II change-over switch is turned on

the two side searchlights light up. The searchlights power can be changed by setting the

SEARCHLIGHTS I, SEARCHLIGHTS II switches on the overhead panel into the TAXIING or

LANDING position.

The switching of search lights is signalled by the illumination of the SEARCHLIGHTS cell in the

AIRFRAME section of CWD. (except for the SEARCHLIGHTS I - TAXIING mode). The

anticollision beacons are switched on and off with the ANTICOLL. BEACON circuit breaker on the

overhead panel. The position lights are switched on and off with the POSITION LIGHTS c~rcuit

breaker on the overhead panel. The static ice detector lghting is switched on and off with the

POSITION LIGHTS circuit breaker and by pressing the ICE DETECTOR LIGHTING push button

between frames No. 6 and No. 7 on the left-hand side window frame.

The Searchlights for lighting of the fin are switched on and off with the FLOOD LIGHTS circuit

breaker on the overhead panel.

STATIC ICE DETECTOR

IN A ~ T I C O L L FLOOD BEACON LIGHTS

FIG. 9-4 EXTERNAL LIGHTING - SCHEMATIC


9.10. EMERGENCY LIGHTING The aeroplane emergency lighting serves for the illumination of the EXIT transparencies at the

main door and at the emergency door as well as for lighting the space just ahead of the aeroplane

exits. The emergency lighting is independent of the main power supplies.

The emergency lighting is normally supplied from the main power bus. If the power supply from

the main power bus breaks down, the emergency lighting bulbs are supplied, for a period of 10

minutes, from four NiCd batteries intended as a standby power source.

When supplied from the main bus, the emergency lighting is actuated by switching on the

BATTERY I, II switches and the CENTRAL WARNING DISPLAY - AIRFRAME circuit breaker on

the overhead panel and by throwing the toggle of the EMERGENCY LIGHTING switch on the left-

hand control panel into its upper position.

If the main power supply fails, the emergency lighting is automatically switched to its own power

supply. When the BATTERY I, II switches and the CENTRAL WARNING DISPLAY - AIRFRAME

circuit breaker are switched on and the toggle of the EMERGENCY LIGHTING switch on the left-

hand control panel is in its lower position, the emergency lighting batteries are being charged.

RELAY P---

( i f inslallcd)

FIG. 9-5 EMERGENCY LIGHTING - SCHEMATIC


CHAPTER 10

COMMUNICATION AND NAVIGATION

10.1. Communication - general

10.2. HF transceiver

10.3. VHF transceiver

10.4. Intercommunication

10.5. Navigation - general

10.6. Course system

10.7. Gyro horizons and turn indicators

10.8. Compass

10.9. Automatic director finders

10.10. Radio altimeters

10.1 1. Navigation system

10.12. Transponder

10.1 3. Flight data recorder

10.14. Weather radar

10.15. Static discharging



10.1. COMMUNICATION - GENERAL The communication system consist of two VHF transceivers, one SW transceiver KHF 950 for

radio connection outside the aircraft. In addition an HF transceiver may be installed in same

aircrafts. An interphone system providing intercommunication of crew. A system of static

discharges is used to protect transceiver operation against electrostatic discharges.

7 Ilo static discharges are located on trailirly d y e s of the wing, tail unit, and fuselage tail cone. A

conductive connection is done betwen all parts of the aircraft. A voice recorder is installed in the

aircraft.

In addition, pilot's s voice recording device may be installed as part of the communication system.

10.2. HF TRANSCEIVER KHF 950 HF radio station is installed on the aircraft. It is used for communication between aircraft

and ground station. The radio station is set in operation by switching on BATTERY I, II switches

and INTERCOM or INTERCOM I, II circuit breakers on the overhead panel and turning

OFFNOLUME knob on the control box of the radio station (KCU 951) on the rear control panel

from the left checked position to the right.

N0TE:l. If the radio station is not set in readines (about 1 minute after switching on), the

selected frequency is not displayed and transmission will be blocked.

2. If the FREQICHAN switch is not depressed, the frequency which was used during the

previous transmission is displayed on the screen.

3. If the FREQICHAN switch is depressed the channel with corresponding frequency

which was used during the previous transmission is displayed on the screen.

Listening in headphones is switched on by HF lever switch on the audio selector box.

Selection of the short-wave radio station for transmission is made by the turnable switch on the

audio selector box. By switching it on to the HF position the signal from the microphone is led to

the SW radio station. By turning the SQELCH noise squelch knob in the anti-clockwise direction

until the noise is very low or disappears.

Control and using of the SW radio station during flight is described in the Flight manual.

The SW radio station is switched off by switching off the above mentioned switches and circuit

breakers.


AUDIO

FIG. 10-1 SCHEME OF KHF 950 RADIO STATION (the variant with two circuit breakers

INTERCOM I, II)


INTERCOM

I"' l!? 1 A U D I O

FIG. 10-2 SCHEME OF KHF 950 RADIO STATION (the variant with two circuit breakers

INTERCOM I, 11)


10.3. VHF COMMUNICATION 10.3.1. Installation VHF transceivers LUN 3524.13 (fig. 10-3)

The airplane is equipped with two VHF transceivers LUN 3524.1 3. Both are designated for

communication with ground stations or with other aircrafts. The transceiver are located on the rear

control panel. Connected to the transceivers are power and control wires and coaxial cables from

antennas. The LH VHF I transceiver antenna is located up on the fuselage between frames No. 7

and 8. The RH VHF II transceiver antenna is located down on the fuselage between frames No. 6

and 7. The installation also includes a circuit-breaker and a fuse through witch power is supplied

to the transceiver, and a rocker-type switch on the control wheel by means of which the

transceiver is switched to the transmition mode. The VHF I and INTERCOM I are fed directly from

the storage battery if connected to on external power source. This is to prevent them from being

damaged if the exteernal source characteristic are not suitable.

The transceiver stand by mode is turned on by switching on the BATTERY I, II switches, VHF I, II,

and INTERCOM 1, I1 circuit-breakers on the overhead panel.

To turn-on, rotate the OFF, PULL TEST knob clockwise. This knob serves moreover to set the

listening level (rotating the knob left and right) and for audio test (pul "out" position of the volume

control knob). Audio for headphones is selected by COM 1 and COM 2 push-buttons PHONE, or

for 2 loudspeakers placed near overhead control panel, by push-button SPEAKER. The rotary

switch on the right side of the audio selector box selects the desired transmitter which will be used

for the transmission (position 1 or 2). The transceivers work permanently in a reception mode.

The transmission mode is turned on by depressing the INT.-VHF switch on the control wheel on

the sinusoid-marked side. A lighted "T" will appear between the "USE" and "STBY" displays on the

transceiver in transmission mode. Under normal operating conditions, the reception and

transmission modes are independent for both pilots.


FIG. 10-3 VHF COMMUNICATION (LUN 3524.13), PASSENGER ADRESS AND

ENTERTAINMENR, INTERCOMMUNICATION SYSTEMS- SCHEMATIC


10.3.2. Installation VHF transceivers KY 196A, M 165A (fig. 10-4, 10-5)

The airplane is equipped with two VHF transceivers (KY 196 A and KX 165 A). Both are

designated for communication with ground stations or with other aircrafts. A navigation system

forms part of KX 165 A transceiver. The transceiver are located on the central instrument panel.

Connected to the transceivers are power and control wires and coaxial cables from antennas.

The LH VHF I transceiver antenna is located up on the fuselage between frames No. 7 and 8. The

RH VHF II transceiver antenna is located down on the fuselage between frames No. 6 and 7.

The installation also includes a circuit-breaker and a fuse through which power is supplied to the - - - - - ~ ~

- - - - transceiver,-and a rochec-typeswitch on the control wheel by means of which the transceiver is -

switched to the transmition mode. The VHF I and INTERCOM or INTERCOM I are red directly

from the storage battery if connected to an external power source. This is to prevent them from

being damaged if the external source characteristics are not suitable.

The transceiver stand by mode is turned on by switching on the BATTERY I, II switches, VHF 11,

NAV I VHF I and INTERCOM or INTERCOM I, II circuit-breakers on the overhead panel.

To turn-on, rotate the OFF, PULL TEST knob clockwise. This knob serves moreover to set the

listening level ) rotating the knob left and right (and for audio test) pull "out" position of the volume

control knob).

Audio for headphones is selected by COM 1 and COM 2 push-buttons PHONE, or for 2

loudspeakers placed near overhead control panel, by push-button SPEAKER. The rotary switch

on the right side of the audio selector box selects the desired transmitter which will be used for the - - - - - -

- - - ~ - - - transmission (position 1 or2I2 - - - - - - - - - - - - - - - - - - - - - - -

The transceivers work permanently in a reception mode. The transmission mode is turned on by

depressing the INT.-VHF switch on the control wheel on the sinusoid-marked side.

A lighted "T" will appear between the "USE" and "STBY" displays on the transceiver in

transmission mode. Under normal operating conditions, the reception and transmission modes

are indipendent for both pilots.


Fig. 10-4 VHF COMMUNICATION (KY 196A, KX 165A), PASSENGER ADRESS AND


(the variant with one circuit breaker INTERCOM)


Fig.10-5 Vt-tF COMMUNICAIION (KY 196A, KX 165A), PASSENGER ADRESS AND


(the variant with twoo circuit breakers INTERCOM I, II)


10.4. INTERCOMMUNICATION, PASSENGER ADDRESS AND ENTERTAINMEN

10.4.1. lnstallation of intercommucation with audio selector box LUN 3591.11 (LUN 3591.13), connection box LUN 3959.22 (LUN 3591.23) and installation passenger address and entertainmen (fig. 10-3)

The intercommunication system is used for intercommunication of the aircraft crew members.

It allows also connection of the passenger adress system.

The system is put into operation by switching on the BATTERY I, II switches, VHF 1, II circuit

breakers, and INTERCOM 1, II in the overhead panel. The transceiver VHF I is on the left side,

VHF II is on the right side. The switching on of the transceiver is perforrned by rotating the knob

INTERCOM in the transceiver panel from the left caged position lo the right. In the receiving mode

audio signal is fed via the audio selector box, amplifier and connector box to the headphones.

Both of the audio selector box are fed with audio signal parallely and a choice of signals is

performed by the VIiF I, II switch in the audio selector box.

On customer's request, a connection box located between frames No. 8 and 9 on the LH side can

be connected in parallel to the pilot's connection box to provide for comrnun~cation between a

third member with the pilot's.

The transmitting mode is energized by pushing the key switch over on the pilot's (copilot's) control

wheel on the sinusoid-marked side. The intercom mode is energized by pushlng the key switch

over on the pilot's (copilot's) control wheel on the telephone-marked side. The signal from the

microphone is fed via Ihe audio selector boxes to the headphones.

If one of the audio selector boxes fails, it is possible to switch over the failed audio selector box to

operating audio selector box by the EMERG. switch in the failed box. In suctt case both parties

listen to common selected audio signals and their microphone circu~ts are parallel-connected

during transmission.

The intercom functions in a normal way.

The passenger adress system is put into operation by switching on the "PA" circuit-breaker on the

overhead panel (at the same thine the BATTERY 1, II switches and INTERCOM I, II circuit

breakers must be switched on). The selection of "PA" with the microphone switch and keying in

the microphone rocker push-button on the pilot's (copilot's) control wheel on the sinusoid-marked

side, amplifies the voice received by a microphone and permits to adress cabin occupants over

the cabin speakers.

10.4.1. lnstallation of intercommucation with audio selector box KMA 24H-70, connection panel 6 581 731 N and installation passenger address and entertainmen (fig. 10-4,10-5)

The intercommunication system is used for intercommunication of the aircraft crew members.

It allows also connection of the passenger adress syslem.

The cabel leading from the headphone set with the centre lead of protection from damage is

connected to the panel with connectors. There are two sockets on the panel - one for headphones

and another for the microphone.

Intercom is put into operation by sw~tching on the switches BATTERY I, II and circuit breaker

INTERCOM or INTERCOM I, II switch which is located on the overhead panel.


Switching on the switch INTERCOM in the steering wheel is switched on by a realy in the audio

selector box which results in connecting through the microphone supply circuit. The signal from

the microphone goes via amplifiers, located in audio selector boxes, and from there to the

headphones. A pilot adjust desired volume of the intercome himself by means of the smaller

concentric knob INTERCOM at the audio selector box itself. In case that either of audio selector

boxes fails there is possibility of switching over the switch for microphone choice so that the

microphone and headphones will be connected direct to the transceiver VHF I. The intercome

operates in a usual way. Basically, it means that the intercome is coupled including feeding of two

indipendent bus bars. The passenger address system withsystem of entertainment broadcasting - - - - - - - - - - - - - - -

for passengers is put into operation by switching on the circuit breaker PA on the overhead panel

(at the same time the switches BATTERY I, I1 and circuit breaker INTERCOM must be switched

on).

By switching the switch for microphone choice into the position PA and by pushing the push button

marked with a telephone symbol on the steering wheel of the left (right) pilot, sound collected by

microphone is amplified and is transmitted over loadspeakers in the passenger compartment.

Note: On some aircraft two INTERCOM I, II circuit breakers can be installed instead of one

INTERCOM circuit breaker.

10.5. INTERPHONE Amog instruments informing the crew about the attitude and direction of flight belong:

--instalation-ofGMK; lGEcourse system (it installed) - - - - - - - - - - - - -

- installation of LUN 1205.31 -8, AGR-74-5 (if installed), gyro horizons and LUN 121 5.XX turn

indicators

- installation of LUN 1221 .01-8 magnetic compass

- installation of KCS 55A, directional gyro compas system (if instdled)

- installation of KI 254,510-22F, AIM 520-1A (if installed)

- aircraft position relative to ground stations

- approach maneouvers

- height above the ground

- other aircraft

An electronic flight information system, DMI, ATC transponder, autopilot and flight recorder are

included in this chapter - - - - - - -


[ FLUX VALUl

FIG. 10-6 BLOCK DIAGRAM OF GMK-1GE COURSE SYSTEM


'l"

FIG. 10-7 BLOCK DIAGGRAM OF KCS 55 A DIRECTIONAL GYRO COMPASS

AIRCRAFT TRAINING MANUAL 1-410 UVP-E, E9, E20

10.6. GMK-1GE COURSE SYSTEM

10.6.1. Installation of GMK-1GE course system (fig. 10-6)

The compass system is actuated by switching the GMK switch on the overhead panel. This

supplies soltage to the TKE 554 PODG relay in the RS-5 distribution box which connect

alternating voltage of 3x36V/400 Hz.

After switching on the GMK switch the red lighting of the control box on the rear control panel will

come on. The following controls are located on the control box:

- geographic latitude change-over switch (1) - see Fig. 10-6

- geographic latitude selector (2)

- gyroscope change-over switch (3)

- function selector (4)

- synchronization change-over switch (5)

- test selector (6)

- control lamps (7)

10.6.2. Installation directional gyro compass system (fig. 10-7)

Two sets of KCS 55 A compass system lor determination of the heading are installed on the

aircraft.

The gyro compass systems are activated by switching on the switches BATTERY I, II

INVERTERS 36 V AC I, II and INVERTERS 11 5 V AC I, II and circuit breakers

GYRO COMPASS I, II on the overhead panel.

The data of gyro magnetic compass are displayed on two KNI 282 radio-magnetic indicators and

two KI 525A navigation indicators, one each being located on the LH and RH instrument panel.

The gyro compass are switched off by circuit breakers GYRO COMPASS I, II and switches on

overhead panel.

10.7. GYRO HORIZONS AND INDICATORS

10.7.1. Installation of LUN 1205.31-8, AGR-74-5 gyro horizons and LUN 1215.XX turn indicators (fig. 10-8)

Gyroscopis instruments e.g. the gyro horizon and the turn indicator serve for determination of the

aircraft attitude in respect to the horizon during straight or in turns. The aircraft pitch and yaw

angles are evaluated by the gyro horizon, bank angles during turns are evaluated by the turn

indicator.

The LUN 1205.31 -8 gyro horizons are switched on by switches BATTERY I , II

INVERTERS 36 V AC I, II and INVERTERS 115 V AC I, II and circuit breaker CENTRAL

WARNING DISPLAY-ELECTRO on the overhead panel and switches GYRO HORIZON LH, RH

on left and right instrument panels. When the horizon is up to speed, the red warning flag in the

view field disappears.


The LUN 1205.31-8 gyro horizons also provide warning signals if the bank angle limits of:

+ 30°during cruise with flaps retracted, and - + 15' when in 18" and 42" flaps conflgurat~ons are exceeded -

Warning levels are switched over automatcally as the flaps are extended or retracted. Information

that the bank limit has been reached rasses from a contact system of the gyro horizon to the

signalling cells KPEH JlEB. BEnMK and KPEH IlPAB. BEnMK located below the gyro horizons.

The spare AGR-74-5 is switched on by switches BATTERY I, II, INVERTERS I, II,

INVERTERS 11 5 V AC I, II and circuit breaker STBY-GYRO HORGON o-n the overhead panel. - - - - - - - - - - - - - - - - -

When the horizon is up to speed, the red warning flag in the view field disappears.

The turn indicators are switched on by switches BATTERY I, II, INVERTERS 36 V AC I, II and

INVERTERS 115 V AC I, II on the overhead panel and switches TURNIBANK IND. on the left and

right instrument panels (even for operation of only one turn-and-bank indicator both switches must

be switched on). When the turn-and-bank indicators are up to speed, the red warning flag in the

view field disappears.

FIG. 10-8 BLOCK DIAGRAM OF LUN 1205.31-8, AGR-74-5 GYRO HORIZONS AND

LUN 1215.XX-8 TURN INDICATORS

10.7.2. Installation of LUN 1205.31-8, AIM 520 gyro horizon and LUN 1215.XX turn indicators (fig. 10-9)

Gyroscopis instruments e.g. the gyro horizon and the turn indicator serve for determination of the

aircraft attitude in respect to the horizon durlng straight flight or in turns. The aircraft pitch and yaw

angles are evaluated by the gyro horizon, bank angles during turns are evaluated by the turn

indicator.

The LUN 1205.31-8 gyro horizons are switched on by switches BATTERY I, 11,

INVERTERS 36 V AC I, I1 and INVERTERS 115 V AC I, II on the overhead panel and switches

GYRO HORIZON on left and right instrument panels. When the horizon is up to speed, the red

warning flag in the view field disappears. The spare AIM 520 is switehed on by switches

BATTERY I, I1 INVERTERS 36 V AC I, 11, INVERTERS 115 V AC I, II and circuit breaker

STBY GYRO HORIZONT on the overhead panel. When the horizont is up to speed, the red

warnlriy flag in the view freld disappears.

The turn indicators are switched on by switches BATTERY I, II (I), INVERTERS 36 V AC I, 11 (4)

and INVERTERS 115 V AC I, 11 (5) On the overliead panel and switches TURNIBANK IND. (3) on

the left and right instrument panels (even for operation of only one turn-and-bank indicator both

switches must be switched on-power 36 V1400 I-lz is made In crose) When the turn-and-bank

indicators are up to speed, the red warning flag in the view field disappears.

10.7.3. Installation of Kl 254, 510-22F, AIM 520 gyro hotlzon and LUN 1215 XX-8 turn indicators (fig. 10-10)

KI 254, 510-22F artificial horizons, AIM 520 standby gyro horizon and LUN 1215 XX-8 turn and

bank indicators are installed in the aircraft. Gyro horrzons and turn and bank indicators serves for

determination of aircraft position to the horizor~ in straight fhght and in turns. Gyro horizon evaluate

aircraft inclination round lateral and longitudinal axes, turn and bank indrcator evaluates bank

angle in turns.

Main gyro horizont are set in operation by switching on switches BATTERY I, II,

INVERTERS 36V AC I, 11, INVERTERS 115 V AC I, II, on overhead panel and switches HORIZON

on left and right instrument panel. After run to operatronal state of gyro horizons the red warning

flag, situated in view field of gyro horizons is slided down.

Standby horizon is set in operation by switching on switches BATTERY I, II,

INVERTERS 36V AC I, II, INVERTERS 1 15V AC I, II and STBY HORIZON on the overhead

panel. After run to operational state of gyro horizon is slided down red warning flag, situated in

view field of gyro horizon.

Turn and bank indicators are set in operation by switching on switches BATTERY I, 11,

INVERTERS 36V AC I, II and INVERTERS 115 V AC I, II on overhead panel and switches

TURNlUAtJK INUICA I OF? oli left arid right ~nstrutnet~t panel (also for operation of only one turn

and bank indicator must be switched both switches - power 36V1400 Hz is made in crose). After

run to operational state the red warning flag situated in view field of turn and bank indicators si

slided down.

FIG. 10-9 BLOCK DIAGRAM OF LUN 1205.31-8, AIM 5520 GYRO HORIZON AND

LUN 1215.XX-8 TURN INDICATORS


10.8. COMPASS The magnetic compass serves the crew to determine the flight course, to check the flight course

and to maintain the aircraft on course. The compass can be employed without modification in

geographical latitude with a vertical intensity component of the earth magnetic field between

- 0.05 to + 0.5 Oe.

10.9. AUTOMATIC DIRECTOR FINDERS 10.9.1. Installation of ARK-15M automatic director finder (fig. 10-1 1)

Two sets of ARK-1 5M automatic direction finders are installed in the aircraft that provide bearing

information to ground stations and commercial radio stations.

The ADF is switched to the operation by switching on the ADF I or ADF II circuit breakers and by

switching the selector switelches on the control units to the respective mode position.

The operation modes of the ADF are as follows:

- automatic direction finding of the selected radio stations

- reception of signals from modulated and unmodulated mode radio beacons

For the automatic direction finding the selector switch on the control unit should be switched to the

COM position. The radio signals from both sense and loop antennas are directed to the input of

the receiver.

lnput to the goniometr of the direction finder receiver comes from the loop antenna. The

goniometr forms an electromagnetic field of the same direction and magnitude as acts on the coils

of the loop antenna. lnput from goniometr, which corresponds to the position of the airplane

relative to radio station is processed and applied to the balanced modulator and compared with

the processed input from sense antenna. The resulting processed signal allows the UGR-4U

combined indicator and IKU-IA radiomagnetic indicator to indicate the radio course.

At the hand controlled direction finding of the radio stations the LOOP press-button switch is used.

In such case the sense antenna is disconnected. The input from the loop antenna is processed in

a normal way as an audio signal. When pressing the press-button switch the rotor of the

goniometr revolves clockwise and the respective input is transmitted to the indicator. The pointer

of the indicator revolves also clockwise.

Through hearing and monitoring the audio signal magnitude in the headphones it is possible to

determine the radio course. It is not distingvished the FROM direction from the TO direction.

Through the use of the LOOP press-button switch it is possible to rotate the pointer and so to

check the automatic direction finder function in the compass mode. When receiving the

modulated and unmodulated signals from beacons, the selector switch on the control unit is

switched to the ANT position and the loop antenna is disconnected. So the input to receiver

comes from the sense antenna.

The output signal from the receiver is presented by an audio signal. The magnitude of this signal

for all directions is the same.

BATTERY I I

FIG. 10-10 BLOCK DIAGRAM OF AIM 520, KI 254, 510-22F GYRO HORIZONS AND

LUN 1215.XX-8 TURN AND BANK INDICATORS SYSTEM


A I M 520

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9, EZO

FIG. 10-1 1 BLOCK DIAGRAM OF ARK-15M AUTOMATIC DIRECTOR FINDERS

AIRCRAFT TRAINING MANUAL L 410 UVP-E. E9, E20

10.9.2. Installation of KR 87 automatic direction finder (fig. 10-12)

One set of KR 87 ADF system are installed in the aircraft. They indicate the bearing to the

received ground beacons or radio stations. The system of ADF is switched to the standby

operation by switching on the BAT rERY I and BATTERY II switches.

The LUN .I205 31-8 gyro horizons are switched on by switches BATTERY I, II,

INVERTERS 36 V AC I, II and INVERTERS 115 V AC I, II on the ovehead panel and switches

GYRO HORIZON on left and right instrument panels.

When the horizon is up to speed, the red warning flag in the diew field disappears. The turn

indicators are switched on by switched BATTERY 1, II ( I ) , INVERTERS 36 V AC I, I1 (4)and

INVERTEliS 115 V AC 1. I1 (5)on the overhead panel and switches TURNIBANK IND. (3) on the

left and right instrument panels (even for operation of only one turn-and-bank indicator both

switched on - power 3614000 Hz is made in cross). When the turn-and -bank indicators are up to

speed, t h ~ red warning flag in the view field disappears.

10.9.3, Installalion of automatic direction finder KR 87 (fig. 10-13)

Two sets nf KR 87 ADF system are installed in the aircraft. They indicate the bearing to the

received ground beacons or radio stations.

The system of ADF is switched to the standby operation by switching on the BATTERY I and

BATTERY II switches, INTERCOM I, II and ADF, RMI circuit breakers on the overhead panel and

by rotating the knob on the KR 87 receiver from OFF position.

The system is switched off by rotating the knob to the OFF position and by switching off the circuit

breakers IN-TERCOM I, 11, ADF, RMI and switches BATTERY I, I1 on the overhead panel. The

system of RMI is switched to the standby operation by switching on the BATTERY I and

BATTERY I1 switches, INTERCOM I, 1 1 (2) and RMI circuit breakers on the overhead panel.

Operation and c~sing in flight are described in Flight Manual.

CAUTION: IT IS SUITABLE SO THAT THE ADF WILL BE OFF DURING ENGINE STARTING


FIG 10-12 BLOCK DIAGRAM OF KR 87 ADF


FIG. 10-13 BLOCK DIAGRAM OF KR 87 ADF

61aq AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E20

10.10. RADIO ALTIMETER 10.10.1. lnstallation of A-037 radio altimeter (fig. 10-14)

The aircraft is equipped with the A-037 radio altimeter whith provides information to the crew of

the actual altitude above ground level as follows:

- the needle of thhe indicator gives continuous indication of the altitude above ground level of

750 m and below.

- it provides acoustic and visual warning of a flight altitude bellow a preset height above ground

level

- it enables proper funtion to be checked

- in case of a failure or loss of ground contact a warning flag comes into view

The radio altimeter is switched on by means of the radio altimeter circuit breaker. A pushbutton of

the face of the altitude indicator serves for checking proper function of the system.

By turning the cap of the warning lamp a decision height index is to be set to 0 - 10 m. When

depressing the TEST button the needle must move tot the altitude of 13 to 17 m. When the TEST

button is released the needle must come back to the zero position and when passing through the

decision height index a warning tone must be heard in the head receivers for 3 to 9 second and a

yellow warning lamp in the button must light up.

10.10.2. lnstallation of KRA 405 radio-altimeter (fig. 10-15)

One set of KRA 405 radioaltimeter is installed in the aircraft. It provides the crew about actual

flight altitude.

The system is switched into standby operation by switching RATTERY I, II and

RADIOALTIMETER circuit breaker on the overhead panel.

Geometric height readonts are indicated on the KNI 415 indicators on the left and right section of

the main instrument panel. The system indudes also annunciation of decision height.

It is switched off by switching of the RADIOALTIMETER circuit breaker and switches

BATTERY I, I1 on the overhead panel. Operationand using in flight are described in the Flight

Manual.

10.10.3. lnstallation of KRA 405 radio altimeter (fig. 10-16)

Two KRA 405 radio altimeters are installed in the aircraft. They provide information for the crew

about geometrical flight altitude.

The radar altimeters are set in operation by switching on the switches BATTERY I, II and circuit

breakers RV I, II on the overhead panel. Geometric height readouts are indicated on the KNI 415

indicators on the left and right section of the main instrument panel. The system includes also

annunciation of decision height. It is switched off by switching of the RADIOALTIMETER I, II

circuit breaker and switches BATTERY 1, I1 on the overhead panel. Operation and using in flight

are described in the Flight Manual.


INTERCOM 1 I

R A D I O ALTIMETER

FIG. 10-14 BLOCK DIAGRAM OF A-037 RADAR ALTIMETERS


FIG. 10-15 BLOCK DIAGRAM OF KRA 405 ALTIMETER


BATTERY RADIOALTIHETER I 11 I II

FIG. 10-16 BLOCK DAGRAM OF KRA 405 RADAR ALTIMETERS


10.1 1. NAVIGATION SYSTEM 10.11.1. Installation of SP5O-ILS navigation system (fig. 10-1 7)

This system includes SP 50-ILS navigation system set which is an airborne radio equipment

sewing for receiving signals of ground beacons for precision approach to the runway, either

according to the international ILS system or according to the SP 50 system used in the RUSIAN.

The set also includes a marker receiver.

The SP 50lILS system is switched on by switching on the SP-50 circuit breaker on the overhead

panel and turning on the volume contoller on the navigation receiver. When the.intrcom and NAV I ,

switch on the switching box are switched on noise will heard in the head phones. The system

required is selected by the ILS-SP 50 selector, and the frequency of the localizer beacon is set by

means of two selectors (frequency of the glide-slope beacon will be set up automatically). The

glide-slope and localizer beacon signals, having been processed in the navigation receiver and a

converter will deflect both the horizontal and vertical D-hass according to the aircraft position

relative to the glide path on both indicators. A failure of one of the systems is indicated by a red

GS or LOC flag.

The marker receiver is switched on by MKR circuit breaker, and by setting the function switch on

the LH indicator from the OFF position to the LO or H I position. When passing over a marker the

appropriate light on the upper side of both indicators with illuminate, and the appropriate acoustic

signal will be audible in the headphones.

Proper function of the marker receiver is chacked by depressing the SIGN. pushbutton on the text

panel located on the LH control panel. The signal lights on both indicators must illuminate.

10.1 1.2. Installation of KX 165 and KNS 81 VORIILSIMKR navigation system (fig. 10-18, 10-19)

NAVICOMM sets of KX 165 and KNS 81 VORIILSIMKR are installed in the aircraft control the

aircraften route (VOR), at landing (ILS) and receiving the signals (MKR).

The NAVICOMM system is turned into standby operation by switching on the following switches

BATTERY I, II, INVERTERS 36 V AC 1, II, INVERTERS 11 5 V AC I, II and circuit breakers

INTERCOM, GYRO COMPASS I, II which are located on the overhead panel and for activation of;

NAV I -

NAV II -

MKR -

NAV 1, VHFl circuit breaker on the overhead panel and by turning the knob from the

its OFF position on the receiver KX 165

NAV 11 circuit breaker on the overhead panel and by turning the knob from its OFF

position on the receiver KNS 81

MKR circuit breaker placed on the overhead panel and switching the switch from its

OFF position on the receiver KR 21

Turning OFF of the system is carried out by turning the knobs to the OFF position on the receiver

KX 165, KNS 81, KR 21 and by switching OFF all circuit breakers NAVI-VHFI, NAVII, MKR,

GYRO COMPAS I, II and switches INVERTERS 36 V AC I, II INVERTERS 1 15V AC I, II and

BATTERY I, I1 placed on the overhead panel.


CAUTION: IT IS RECCOMENDED TO TURN ON THE RECEIVER ONLY AFTER ENGINE

START-UP.

- L . 2

FIG. 10-1 7 BLOCK DIAGRAM OF SP-SOIILS


FIG. 10-19 BLOCK DIAGRAM OF KX 165 AND KNS 81 VORALSIMKR SYSTEM (IT IS

VALID FOR AIRCRAFTS, WHEN I.B.NO. L 410 UVP-E1040B IS

PERFORMED)


10.11.3. Installation of GARMIN 155 GPS navigation system (fig. 10-20)

The GARMIN 1555 Global Positioning system (GPS) is a satellite based navigation system that

provides high precision of data acquisition, high speed of the first aircraft position finding, and

continnous updating of navigation data. The GPS receiver contains the Jeppesen database that

contains a large amount of information on airports, navigation system, etc.

NOTE: See the GPS GARMIN 155 Personal Navigator Owner's Manual and the Aircraft Flight

Manual for more details.

CAUTION: - THE GPS CAN ONLY BE OPERATED IN CONJUCTION WITH ANOTHER - - - - - - - ~ ~

- - - -

NAVlGATION SYSTEM

-THEJEPPESENDATABASEMUSTBEUPDATEDEVERY28DAYS

- THE GPS CAN BE USED FOR VFR OPERATION ONLY

The GPS is set in operation by switching on the BATTERY I, II switches and the GPS circuit

breaker on the overhead panel. The system is then operated according to the instructions given in

the Owner's Manual by means of controls on the GPS receiver control panel.

10.1 2 REPLIER 10.12.1. Installation of SRO-2 replier (fig. 10-21)

This part includes an equipment used for automatic aircraft identification as a response to inquiry

signals from ground identification searchers or inquiers from other aircraft. This equipment

consists of a SRO - 2 replier. In the following table all instruments and main parts of a replier - - - - - system are listed.

By switching on the INVERTES 115V AC I, II and BATTERY I, II switches and the SRO circuit

I breakers located on the overhead panel voltage is supplied to contacts of the switch on the control

I panel. By switching over of this to switch position POWER the replier is put into operation. Then I the CODE ON lamp lights on the control panel and 1 - 2.5 min. after switching on the POWER I I lamp lights up.

I During the flight the replier response to searchers inquiry (ground or airborne) is manifested by I blicking of the REPLY neon lamp on the control panel. The system can be destroyed if necessary I I by pressing the explosion push button.


BATTERY

b@ 0

I GPS

' i AUDIO SELECTOR i AUDIO SELECTOR i BOX BOX

FIG. 10-20 BLOCK DIAGRAM OF GARMIN 155 GPS SYSTEM


BATTERY I If

FIG. 10-21 BLOCK DIAGRAM OF SRO-2 REPLIER SYSTEM

10.12.2. Installation of SO-69 transponder (fig. 10-22)

The active - response radio equipment consists of instruments and equipment which provide

information on the aircraft attitude, its identification, and its flight parameters in cooperation with

ground stations. This equipment comprises an SO-69 transponder.

When the BATTERY I, II and INVERTER 115V AC I, II switches are switched on, supply voltage is

fed to the contacts of a switch located on the transponder control panel. The transponder is

activated. By setting this switch to the BKJl position and by setting the mode selector to the

rOTOB position. When the transponder is switched on, the lighting of the control panel will come

on. Set the QFE pressure on the encoding altimeter.

WARNNING: THE REQUIRED PRESSURE CAN BY ONLY BE SET ON THE ALTIMETER IF

THE ALTIMETER HAS BEEN CONNECTED TO 27V SUPPLY VOLTAGE AND

SIMULTANEOUSLY TO 115Vl400 HZ VOLTAGE. NON-COMPLIANCE WITH

THESE CONDITIONS MAY RESULT IN A DAMAGE TO THE ALTIMETER.

In flight the transponder automatically transmits reply to the interrogation of the ground ATC

secondary surveillance radar.

NOTE: On the ground the mode selector may be set to the YBA position for testing only.

10.13. FLIGHT DATA RECORDER 10.13.1. lnstallation of BUR-1-2G flight data recorder (fig. 10-23)

The BUR-1-2G flight recording equipment is specified for acquisition and recording of flight

parameter information during the flight and for stormg this information even in a case of an air

crash. The parameters are recorded onto a magnetic tape. The equipment provides records of 25

analog parameters and of 48 discrete commands. Special ground equipment type SNUO-1 is

used for processing of records deoosited in shock-resistant and fireproof containers.

The BUR-1-2G flight recording eqbiprnent consists of'the following functional units:

- flight data gathering unit BSPI-4-2 interconnected with two.encoders,

- encoders for input signal and for interrogation frequency programming.

The analog input signals of the individual transmitters are converted to a numeric code in binary

system. Signals modified in the described way are recorded onto a magnetic tape by means of the

ZBN-1-1 recorder. The recording equipment is deposited in a metal container with a high

resistance against shocks and fire.

The PU-25 control panel consists of two parts, one of these parts is intended for putting the

system into operation and for operational data loading and the other one serving for functional

supervision issuing signals for the control panel display section.

The isolating box is used to protect the recorder against spurious short-circuits in the aircraft

system. In addition, the BUR-1-2G recorder system comprises a system of receivers and

transmitters and a system of sensors the signals of which are recorded even when their sources

belong to another systems (e.g. to the radar altimeter).

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9. EZO

SRO 0

I I

FIG. 10-22 BLOCK DIAGRAM OF SO-69 TRANSPONDER SYSTEM

AIRCRAFT TRAINING MANUAL L 41 0 UVP-EI E9. E20

List of registered parameters of the BUR-1-2G flight recorde-

Analog (continuous) record:

1. Altitude

2. Geometric height

3. Instrument speed

4. Overload in the x-axis

5. Overload in the y-axis

6 Overload in the z-axis

7. Elevator displacement

8. Displacement of the ailerons

9. Rudder displacement

10. Displacement of the rudder-trim tab

1 1. Angular speed y,

12. Angular speed Y,

13. Angular speed C I ~

14. Angle of bank

15. Angle of pitch

16. Heading

17. L.H. engine ECL position

18. R.H. engine ECL position

19. L.H. engine torque value

20. R.H. engine torque value

21. L.H. engine generator r.p.m.

22. R.H. engine generator r.p.m.

23. L.H. engine propeller r.p.m.

24. R.H. engine propeller r.p.m.

25. Voltage in the 28 Votts distribution system

Recording of one-shot commands:

1. L.H, engine fire signal on

2. R.H. engine fire signal on

3.

4 .

5.

6.

7. VIL transceiver keying

8. 21P transceiver keying

9. Port engine propeller control lever in position FEATHER


10. Startboard engine propeller control lever in position FEATHER

11. Airframe de-icing system turned on by the air crew

12.

13. Signal "LANDING GEAR DOWN"

14. Signal confirming the presence of the 36 V voltage in the aircraft distribution system

15. Signal indicating dangerous altitude reading of the radar altimeter

16. Signal LIFT SPOILERS EXTENDED

17. Signal indicating minimum oil pressure in the L.H. engine

18. Signal indicating minimum oil pressure in the R.H. engine

19. Signal indicating minimum fuel pressure in the L.H. engine

20. Signal indicating minimum fuel pressure in the R.H. engine

21. Signal L.H. ENGINE AUXILIARY PUMP ON

22. Signal R.H. ENGINE AUXILIARY PUMP ON

23. Port engine IELU on (function limiting)

24. Standboard engine IECU on (function limiting)

25. L.H. engine BETA CONTROL signalling on

26. R.H. engine BETA CONTROL signalling on

27. Signal confirming the water injection system activation

28. Signal MINIMUM FUEL REMAINDER - left side

29. Signal MINIMUM FUEL REMAINDER - right side

30. Signal WlNG FLAPS IN O0 POSITION

31. L.H. engine automatic feathering circuit on

32. R.H. engine automatic feathering circuit on

33. L.H. engine generator malfunction

34. R.H. engine generator malfunction

35. Rime signalling active (rime indication)

36. Pneumatic de-icing system function signal

37. Free

38. Free

39. Free

40. Voltage in the 115 Vl400 Hz distribution systen

41. Stall speed signalling system on and in function

42. Signal WING FLAPS IN 1 8°POSITION

43. Course system SLAVE mode

44. Voltage at the S 28 busbar

45. Voltage at the S 39 busbar

46. Automatic pitching active - port side

47. Automatic pitching active - starboard side

48. Signal WlNG FLAPS IN 42O POSITION


10.13.2. Installation of F 1000 flight data recorder

The model F 1000 is a Fairchild Flight Data Recorder with Solid State memory. It is specified for

acquisition and recording of Flight parameter information durring the flight and for staring this

information even in a case of an air crash. There are 12 independent synchro signal inputs for

conventional synchro signals 11.6 V with 26 V AC excitation.

There are no controls or switches associated with the FDR and its operation is completely

automatic. To operate, first disengange the aircraft circuit breaker for the recorder while the rest of

the aircraft electric system is on. The Recorder Fault light should turn on. Next, engage the circuit

breaker to apply proper aircraft power to the recorder. The recorder fault light should turn off in

approximately five seconds. Should the recorder fault light come on after seven seconds, the

recorder may not be functioning properly (see Section 4.0) or the input data to the recorder is

incorrect.

Should any parameter expected by the recorder (ARING 542 mode) be missing, the recorder fault

light will turn on. If no tested input parameter to the recorder 's missing and the self test BITE

circuitry does not find a defect within the FDR, the recorder will operate automatically until power

is removed. The recorder will convert analog data into digita! data and record the information in its

memory. The recorder continuously records and retains flight data as presented to it.

List of registered parameters of the flight data recorder

1. Time

2. Altitude

3. Airspeed

4. Heading

5. Vertical acceleration

6. Pitch attitude

7 Roll attitude

8. Radio transmission keying

9. Engine RPM-L

10. Engine RPM-R

11. Engine Torque - L

12. Engine Torque - R

13. Trailing edge flaps

14. Thrust reverser position - L

15. Thrust reverser position - R

16. Spoiler

17 Marker beacon

18. Autopilot engagement

19 Longitudinal acceleration

20. Lateral acceleration

21. Total air temperature

AIRCRAFT TRAIN ING MANUAL L 410 UVP-E, E9, E20

22. Control surface - Alleron

23. - Rudder

24. - Elevator

25. - Pitch trim

10.14. WEATER RADAR

10.14.1. Installation of 3A 81 3 weather radar (fig. 10-23)

The radio device for strom finding - 3A 813 weather radar serves for investigation of weather

configurations with possibility to determinate of the stage of their dangerous to flight and serves

for terrain mapping

Weather radar system is set in operat~on by switching on the circuit breakers BATTERY I, II

INVERTERS 115V I, II and the circuit breakers RADAR, the overhead panel and switch

HORIZON on the right instrument panel.

Proper using of weather radar in fliqht is described in Flight Manual. Switching off the weather

radar is made by switching the above mentioned circuit breakers and switches.

10.14.2. Installation of RDR 2000 weather radar (fig. 10-24)

The radio device for strom finding - RDR 2000

- serves for investigation of weather configurations with possibility to determinate of the stage of

their dangerous to flight and serves for terrain mapping.

Weather radar system is set in operation by switching on the circuit breakers BATTERY 1, I1

INVERTERS 115V I, II and the circuit breakers RADAR, on the overhead panel and switch

HORIZON on the right instrument panel. The antenna scans in the sector 90d and is longitudinally

and cross-wise stabilized withm the range of slopes and banks up to + 25+ if stabilization is

switched on. If the navigation is realized according to the equipment. Proper using of weather

radar in fllight is described in Flight Manual.

Switchmg off the weather radar is made by switching the above mentioned circuit breakers and

switches.

10.14.3. Installation of RDS 81 weather radar (fig. 10-25)

The radio device for storm finding - RDS 81 with the GC 381 A radar graphic unit serves for

investigation of weather configurations with possibility to determinate of the stage of their

dangerous for flight and serves for terrain mapping.

Weather radar system it set in operation by switching on the circuit breakers BATTERY I, 11,

INVERTERS 115V AC I, II and the circuit breakers RADAR, INTERCOM I, NAV I1 and GYRO

COMPASS I, II on the overhead panel and GYRO HORIZON switch on the left instrument panel.

Proper using of weather radar in flight is described in Flight Manual.

Switching off the weather radar is made by switching the above mentioned circuit breakers and

switches.


FIG. 10-23 BLOCK DIAGRAM OF 3A 813 WEATHER RADAR

AIRCRAFT TRAINING MANUAL L 41 0 UVP-El E9. EZO

FIG. 10-24 BLOCK DIAGRAM OF RDS 2000 WEATHER RADAR


VERTICAL iL7

FIG. 10-25 BLOCK DIAGRAM OF RDS 81 WEATHER RADAR SYSTEM

10.15. STATIC DISCHARGING

Normal operation of the radiocommunication and navigation equipment of the aircraft is

endanered by interferences originating in various electric power appliances and devices, such as

electric motors, relays, contactors etc., as well as by interferences due to accumulated static

electricity. In order to minimize these interferences, the aircraft is protected by:

- elektrical bonding of aircraft assemblies (see section 024.60.00)

- installation of static dischargers.

The static dischargers provide for discharging static electricity charges from the aircraft surface

into the atmosphere.

--- ---.

FIG. 10-26 LOCATION OF STATIC DISCHARGERS

(1) Rudder - 2 off, (2 ) Elevator - 2 off, ( 2 ) Ailerons - 4 off, (4) Stabilizer tips -

- 2 off, (5) Fuselage tail cone - adjacent to position light - 1 off, (6) Wing tip

fuel tanks - 2 off


CHAPTER 11

PITOT - STATIC SYSTEM

1 1.1. General




11 .l. GENERAL The pilot and static system consists principially of three circuiis:

- pitot pressure circuit

- static pressure circuit

- ram-air pressure head circuit

Pitot and static system feeds the airspeed indicators, altimeters, rate-of-climb indicators flight-

- recorder and air speed signalization.

The airspeed indicator with stall speed signalizer is connected to the ram air pressure head

system.Pitot presure is derived from Pitot tubes situated between frames No.4 and NOS. Pitot

pressure from both tubes is fed to the upper selector cock which enables total pressure to pass to

the airspeed indicators and the stall speed signalizer. If the selector cock is set to PITOT I position

the airspeed indicator and the stall speed signalizer on the LH instrument panel are supplied from

the left Pitot tube. The airspeed indicator and speed signalizer on the RH instrument panel are

supplied from the right Pitot tube.

In position PITOT II the above stated instruments are supplied from the right Pitot tube.

The bottom selector cock makes operative the emergency static port. Moisture traps are installed

at the lowest parts of the tubing. Static pressure is sensed by two static pressure heads. Both

static pressure heads are interconnected and pressure is fed further to nipples on airspeed

indicators, altimeters, rate-of-climb indicators, air-speed signalizers and flight recorder sensors.

Moisture traps are installed at the lowest parts of the tubing. Pressure from the ram-air pressure

head is led through the wing and through the fuselage to the instrument panel and to the

corresponding nipple on the stall signalizer.


FIG. 11-1 PlTOT - STATIC SYSTEM


CHAPTER 12

EQUIPMENT1 FURNISHINGS

12.1 General

12.2. Flight compartment and pilots ' seats

12.3. Instrument and control panels

12.4. Passenger compartment and passengers seats

12.5. Portable oxygen equipment

12 6. Toilet

12 7. Baggage compartments

12.8. Emergency equipment



- - - - - -


12.1. GENERAL In the aircraft fuselage appropriately furnished compartments are provided for the crew,

passengers and cargo (baggage), as well as for a toilet. Internal equipment includes also

emergency and rescue equipment:

first - aid kits, an axe, a crew call signalling (if installed), life jackets, life rafts emergency

packages (if installed) and Emergency Locator Transmitter.

12.2. FLIGHT COMPARTMENT AND PILOTS' SEATS Flight compartment

The flight compartment is located in the fuselage nose section between the frames No. 4 and 7. It

is separated from the passenger compartment by a fixed partition with folding wings. The flight

compartment is equipped with pilots seats, a navigators table, map spring bracket (if installed)

a spare fuse box, a flight documentation case and a portable lamp. Greater part of the flight

compartment is provided with upholstery and acoustic insulation for noise reduction.Located on

the right-hand side of the flight compartment is an emergency exit door.

FIG. 12-1 FLIGHT COMPARTMENT FURNISHINGS

(1) Pilots seats, (2) Navigator's table, (3) Spare fuse box, (4) Portable lamp,


The flight compartment is equipped with a electric fan. The DV 3 fan is attached with two screws-

- to the windshield central column. The fan is put into operation by switching the FAN (AZRGK 5)

circuit breaker situated on the clock panel.

Pilot' s seats

The seats of both pilots are fitted with seat adjusting mechanism and are attached to the floor.The

seat adjusting mechanism is placed between the upper part of the front strut and the bottom part

of the rear seat strut. The adjusting mechanism enables the seat to be shifted upwards-

downwards (60 mm travel) and fnrwards-backwards (1 30 mm travel). Longitudinal and vertical

seat positions are mutually dependent.

The seat proper consists of a metallic skeleton and foamy filler material which are coated with

fabric. Fixed on the rear backrest sides are the pockets for life jackets.

Each seat is provided with 2 shoulder and 2 thigh straps fitted with a ring lock and fixed to the seat

skeleton. The straps are inserted into the lock in the position marked with a dot. Then the lock

control knob is turned into the position marked with a cross in which.the safety belts are secured.

By turning the control knob into the position marked DO", the safety belts get loosened. The length

of individual straps may be changed by the buckle fixed on each strap.Both the pilot seats have

removable elbow rests.

At custome's reguest the pilot's seat can be fitted with arm rests. The arm rests can be installed

on the LH side of the captain's and on the RH side of the co-pilot's seat. The arm rests can be set

five positions in the up-down direction. To adjust a suitable arm rest position, the safety pin is to

be removed, the arm rest raised or lowered, and the safety pin insarted back into an appropriate

hole in the arm rest frame.


FIG. 12-2 PILOT' S SEAT

(1) Pilot's seat, (2) Front strut, (3) Rear strut, (4) Front bracket, (5 ) Rear

bracket, (6) Adjusting mechanism, (7) Front bracket, (8) Rear bracket,

(9) Central bracket, (1 0) Control wheel, (1 1) Cable, (12) Safety belts, (1 3) Lock

knob, (14) Pin, (15) Lock, (16) Buckle.

Lock knob positions:

- insertion of individual straps into the lock

x - belts fastened

o - belts loosened


FIG. 12-3 ELBOW REST ON PILOTS ' SEATS

(installed against the customer's order)

(1) Rest, (2) Pin, (3) Safety pin, (4) Spring, (5) Plate, (6) Screw with washer.

NOTE: The elbow rest is placed on the left hand side of the 1st pilot seat and

on the right hand side of the 2nd pilot seat.

ARCRAFT TRAlNlNG MANUAL L 410 WVP-E, E9, UO

12.3. INSTRUMENT AND CON'TROL PANELS In the pilot' s cabin, there are concentrated instruments and controls on the instrument panel

necessary for the engines controll~ng and for the aircraft control.Beside those, some instruments

and controls are mounted on the overhead panel and on the glareshield, a clock are installed on

the control column of the windshield and two stop watches are located in the glareshield.

FIG. 12-4 INSTRUMENT PANELS - LEFT SECTION

NOTE: Location of controls (circuit - breakers, switches), their models and

layout on the left instrument panel as shown in the figure are given

for information only. The left instrument panel can be modified for

different airplane versions and their variants.

AIRCRAFT TRAINING MANUAL L 410 UVP-E, €9, E2O

FIG. 12-5 INSTRUMENT PANELS - CENTRE SECTION


layout on the centre instrument panel as shown in the figure are given for

information only. The centre instrument panel can be modified for different

airplane versions and their variants.

FIG. 12-6 INSTRUMENT PANELS - RIGHT SECTION


layout on the right instrument panel as shown in the figure are given for

information only. The right instrument panel can be modified for different



FIG. 12-7 LOCATION OF INSTRUMENT AND CONTROLS ON THE CENTRAL

CONTROL PANEL


layout on the front control panel as shown in the figure are given for

information only. Thefront control panel can be modified for different airplane

versions and their variants.

AIRCRAFT TRAINING MANUAL L 41 0 UVP-Ey E9. E20

FIG. 12-8 LOCATION OF' CONTROL ON THE CENTRAL CONTROL PANEL


layout on the central control panel as shown in the figure are given for

information only. The central control panel can be modified for different


FIG. 12-9 LOCATION OF INSTRUMENT AND CONTROLS ON THE REAR CONTROL

PANEL


layout on the central control panel as shown in the figure are given for

informatim only. The central control can be modified for different



FIG. 12-10 LOCATION OF INSTRUMENT AND CONTROLS ON THE LEFT CONTROL

PANEL


layout on the left control panel as shown in the figure are given for

information only. The left central control panel can be modified for

different airplane versions and their variants


FIG. 12-11 LOCATION OF INSTRUMENT AND CObrTROLS ON THE RIGHT

CONTROL PANEL


layout on the right control panel as shown in the figure are given for

information only. The right control panel can be modified for different

airplane versions and their variants

I RADIO - NAVIGATION

1 ENGINES

DEICING ! i i l l l l i l 1M

W ! l m IIlM 4 I,

p & p $ - p g 4 L 2 - ! - 3 . . I

FIG. 12-12 LOCATION OF CONTROL ON THE OVERHEAD PANEL


layout on the overhead panel as shown in the figure are given for

information only. The overhead panel can be modified for different

airplane versions and their variants


FIG. 12-13 FUSE PANEL

NOTE: The fuse values and the text of the text of the label differ according to

the equipment of a particular airplane version a label on the fuse

panel snows the system protected (e. g. OIL PRESSURE) and the

value of the fuse (e.g. 0.63)


FIG. 12-44 GLARESHIELD OF THE INSTRUMENT PANEL

FIG. 12-1 5 GLARESHIELD OF THE INSTRUMENT PANEL


layout on the glareshield of the instrument panel as shown in the

figure are given for information only. The glareshield of the instrument

panel can be modified for different airplane versions and their variants

AIRCRAFT 'TRAINING MANUAL L 410 UVP-E, E9, E20

12.4. PASSENGER COMPARTMENT AND PASSENGER S SEATS Passenger compartment

The passenger compartment is situated behind the flight compartment as far as the bulkhead

No. 21.

The passenger compartment is upholstered and protected a~ainst noise by soud insulation.

Passenger compartment floor is covered by a carpet. Situated on the left-hand side in the rear

part of the passenger compartment is the entrance door. Near the entrance door there is a shelf

with hooks where headwear and clothes may be put off. The passenger compartment is provided

with foldable serving plates and a newspaper magazine holder (if installed).

The airplane can be modified and delivered in the following versions:

Passenger transport version - 19 seats (basic version)

The passenger compartment is provided with 5 single seats on the LH side and 7 double seats on

the RH side

FIG. 12-16 PASSENGER COMPARTMENT FURNNISHINGS

(1) Single passenger seat, (2) Double passenger seat, (3) Rear baggage

compartment, (4) Shelf for headwear and rack, (5) Toilet, (6) Foldable serving

plate (if installed), (7) Newspaper and magazine holder (if installed),

(8) Refresh, bar (if installed).

AIRCRAFT TRAINING MANUAL L 41 0 UVP-El E9; E20 -

Passenger transport version - 17 seats

The passenger compartment is provided with 5 single seats on the LH side and 6 double seats on

the RH side.

FIG. 12-1 7 PASSENGER COMPARTMENT (COMMUTER AIRCRAFT)

(1) Single passenger seat, (2) Double passenger seat, (3) Airstar locker,

(4) Toilet, (5) Folding table, (6) Newspaper and magazine holder, (7) Rear

baggage compartment, (8) Pilot emergency exit door, (9) Passenger emergency

exit door, (10) Rack, (11) Room for baggage, (12) Signalling panel (Signalisation

to call in the crew), (1 3) Curtain, (14) Fan

Passenger transport version - 11 seats

In the front section of passenger compartment, three rows of seats are installed (i.e. 7 seats).

The seats are provided with foldable serving plates. In the rear section (between frames No. 13

and 18), two double seats with armrests are installed facing a conference table with foldable

leaves.

FIG.12-18 PASSENGER COMPARTMENT LAYOUT (EXECUTIVE VERSION)

(1) Passenger seats, (2) Front double seat with armrests, (3) Conference

table, (4) Rear double seat with armrests, (5) Box for storing airstairs,

(6) Buffet, (7) Toilet, (8) Baggage compartment, (9) Shelf for neadwer and

rack, (lO)Curtains, (1 1) Newspaper holder (12) Emergency exit.

r- - ---L. 4 4 F J T i +=-, - i no- 11 , 3 1q iv 4 ,k- A

1 1 1


Passenger executive version - 8 seats

If the fronnt and rear section of passenger compartment, two double seats with armrests are

installed facing a conference table with foldable leaves.

FIG. 12-19 PASSENGER COMPARTMENT LAYOUT (EXECUTIVE VERSION)

(1) Newspaper rack, (2) Forward double-seat, (3) Table, (4) Rear double seat,

(5) Toilet, (6) Curtain, (7) Refrigerator (if installed), (8) Galley (if installed),

(9) Airstar locker, (10) Fan, (1 1) Attendant seat

Flying ambulance

There is a possibility of installing medical equipment. The medical equipment serve for

transportation of wounded (ill) persons on seats and a nurse.

FIG 12-20 ARRANGEMENT OF ILL PERSONS TRANSPORT EQUIPMENT

( I ) , (2) Stretcher, (3), (4) Seat, (5) Table for a nurse, (7) Litter bin, 8) Canister

with drinking water


Parachute jumping

The parachute jumping consists of the following basic items:

- seats

- anchor cables

FIG. 12-21 LOCATION OF SEATS AND AVAILAB!-E SPACES IN THE AIRPLANE

CABIN - PARACHUTE VERSION

(1) Seats of parachutist, (2) Holes and covers in the ceilling for PARA ropes,

(3) Rear baggage compartment, (4) Toilet, (5) Paratroop guide seat.

CARGO VER!YON

The cargo version of the aicraft is intended for quick transprt of different cargo es (goods and

equipment) of limited dimensions with the total mass up to 1250 kg, namely:

- 1000 kg in a container,

- 100 kg in the front baggage compartment,

- 150 kg in the rear baggage compartment.

The cargo version is characterized by a stationary transport container placed in the passenger

compartment of the aircraft.

Passengers' seats

The passengers seats are designed as single-seat units situated on the L.H. side and as double-

seat sets situated on the R.H. side. The single-seat units and double-seat sets differ from each

other only by the legs by means of which the seats are fixed to the passenger compartment floor.

The seat structure consists of a metallic skeleton filled with foamy material and coated with fabric.

The seats are fitted with inflammable removable washable covers buttoned to them. Each

passenger's sea t is provided with a safety belt.

The safety belt consists of two straps,two anchoring buckles and a lock. The length of the safety

belt is adjusted on the lock.

The safety belt is fastened by coupling both lock parts. By depressing the red pushbutton on the

top lock side the safety belt is loosened. The pockets for life jackets are attached to the underside


of the seats (in the middle).

The seats may be fitted with foldable serving plates against the customer's order.

FIG. 12-22 CARGO VERSION EQUUIPMENT

(1) Container with cargo (sections A, B), (2) Seat for the person

accompanying the cargo, B 094 125 N (aircraft technician), (3) Double

passenger seat,

(4) Rear baggage compartment, (5) Toilet, (6) Headwear shelf with rack.


FIG. 12-23 PASSENGERS' SEATS

(1) Single-seat unit, (2) Double-seat set, (3) (4) (5) Legs, (6) Flange, (7) Safety

belt, (8) Lock, (9) Removable cover.


12.5. PORTABLE OXYGEN EQUIPMENT Portable oxygen equipment is determined for supplaying of the crew and passengers with oxygen.

The oxygen equipment for the crew consists of two LUN 1807 oxygen masks and two

BKP-2-2-210 oxygen breathing apparatuses or two KM-1-4 oxygen masks and two BKP-2-2-210

oxygen breathing apparatuses or two DUO-SEAL oxygen masks and four SCOTT 5500 oxygen

breathing apparatuses.

The oxygen masks and apparatuses are located behind backrests of the pilot seats.

The oxygen equipment consists of two BKP-3-2-210 oxygen breathing apparatuses located in the

passenger compartment in a holder in the forward bulkhead of the rear baggage compartment,

each of the apparatuses inclused two MKP-IT or two MARK 1 oxygen masks each of the

apparatuses includes oxygen masks.

The handling of portable oxygen equipment is described in the L 410 UVP-E, E9. E20 Flighht

Manual.

FIG. 12-24 SCHEME OF LOCATION OF BKP-2-2-210, BKP-3-2-210 OXYGEN

EQUIPMENT (if installed)


FIG. 12-25 SCHEMA OF LOCATION OF SCOTT OXYGEN EQUIPMENT (if installed)


12.6. TOILET

One of three following toilets can be installed on aircraft:

- toilet with collecting vessel (placed at the 21st frame)

- toilet with collecting bag (placed between frames No. 19th to 21st)

- discharging toilet (placed between frames No. 19th to 21 st)

Toilet with collecting vessel - versior A

The aircraft is equipped with a toilet. The toilet is equipped with a collecting vessel for foeces and

it is installed inside the toilet compartment (see chapter 025) in the aircraft. The toilet with the

collecting vessel is intended for manual removal of foeces.

The toilet consists of a vessel with a dish and of a seat with a lid. The seat is attached to the

vessel by a pair of spring loaded hinges installed on both sides of the vessel. The seat lid is

attached by two hinges. The toilet vessel bottom carries a pair of feet. One of them (that without

hole) runs in a circular groove and the other one (with a hole) is engaged by a pivot. The pivot

features a locking snap. This arrangement permite to pull out the entire toilet vessel to the front.

A hole in the seat serves for venting the toilet intterior space. The venting tube runs under the

passenger cabin floor into the right hand landing gear nacelle.

Toilet with collecting bag

The aircraft is equipped with a toilet with a collecting bag of fecal solids located in a toilet space of

the aircraft. The toilet with the collecting bag is determined for operating conditions with a manual

removal of fecal solids.

The toilet consists of a jacket, a sitting desk with a bowl and a cover, a bucket and a collecting

bag. At the toilet jacket bottom footings in which a bucket is set are riveted. The collecting bag is

inserted in the bucket with its upper border overleaping the bucket border.

The bucket lid is fastened from inside to the toilet jacket. The sitting desk is secured with springs

at its both sides. The jacket is provided with two outlets. The first one is determined for the waste

connection from the wash-basin wh113t the other one is determined for dearing of the toilet internal

space.


FIG. 12-26 TOILET WITH COLLECTING VESSEL - VERSION A

(I) Toilet vessel, (2) Collecting vessel dish, (3) Seat, (4) Clamp, (5) Bellows,

(6) Locking lamp, (7) Spring, (8) Handle, (9) Cover, (10) Snap locks-

-13 pieces, (1 1) Box fox tissues, (12) Box for toilet paper, (13) Belt


FIG. 12-27 TOILET WITH COLLECTING BAG

(1) Toilet, (2) Sitting desk with a bowl, (3) Lid, (4) Bucket, (5) Footings,

(6) Bucket lid, (7) Spring, (8) Collecting bag, (9) Wash-basin discharging hose,

(10) Internal toilet space deairing hose, (1 1) Passage, (12) Wing nut,

(13) Tube, (14) Screw with a washer, (15) Wash-basin panel, (1 6) Hand rail,

(17) Toilet paper box, (18) Paper towel box, (19) Ring.


Discharging toilet

The discharging toilet with rinsing device consist of a jacket, collecting vessel for fecal solids, seat

part, collection vessel dish, rinsing device and control mechanism. The seat part is equipped with

a closing lid. The both part are tipped out of the seat position to discharge position by means of

the pedal and other mechanisms. The back of the dish leads into the discharge neck which in the

seat position leans against the spring washer on the jacket and in this way it separates the dish

from the space with fecal solids. The dish neck is connected to the collecting vessel neck by

textile collar. The rinsing device is equipped with distributing pipes in the dish. These pipes are

connected to two belows (tanks) with cleaning liquid.

The belows are provided with suction valves for sucking liquid from the tank and with delivery

valves for pressing liquid up to the distributing pipes. The bellows are controlled with mechanism

from the tipped dish. A reserve tank is located on the jacket. it contains 4 1.

In the lower part of the collecting vessel is a neck for connecting the waste hose of a fecal truck.

The waste neck is provided with a closure and throttle controlled with the hand lever.

Further, in the toilet is fixed inlet piping for rinsing pressure water supply from a fecal truck. The

piping by means of two hoses leads into the dish. The piping is equuipped with closures. Both the

closures are accessible after opening lid on the fuselage. At opening the throttle with the hand

lever, the dish is tipped into the discharge position simultaneously.

By lifting the cover the toilet is put into operation. After using close the cover and stop on the

pedal. The dish will turn and contents will tip out into the collecting vessel. Simultaneously the

bellows are stretched and in these the liquid for the toilet is sucked at open suction vales. At return

movement of the dish the suction valves will close and by opening the delivery valves of the liquid

for the toilet is pushed into the distribution pipings and the inside of the dish is rinsed. The wash

basin is installed in the toilet room. Water from water mains (andlor with admixture of deodorant)

for washing is in the tank with content of 5 1 at the wash basin case. Water from tank is led into the

discharge valve which leads into the basin. Water discharge is carried out by turning the handle of

the valve.

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E 9 E20

FIG 12-28 DISCHARGING TOILET

(1) Dish, (2) Toilet desk, (3) Tiltable cover, (4) Jacket, (5) Collecting vessel,

(6) Tank, (7) Rinsing device. (8) Pedal with control mechanism, (9) Outlet

neck, (10) Rinsing neck, (1 t) Outlet control lever.

12.7. CARGO VERSION The front and the rear baggage compartments are used as cargo compartments

The front baggage compartment is sltuated between the bulkheads No 1 and 4 The rear

baggage compartment is located between the frame No 19 and bulkhead No 21 and is divided

by means of a sandwich panel into two sectlons Two detachable sandwich covers are provided to

close both sections When installed, they can secured in place with a lock

When stowing the baggage, remove and store the covers beh~nd the passenger's seats so that

cover damage may be prevenfed

The technical compartments in which the different major components of the alrcraft systems are

installed have no spec~al design and are a constituent part of the aircraft structure

Aircraft can operate also In cargo verslon - see paragraph 12 4 of this manual


12.8. EMERGENCY EQUIPMENT Aircraft emergency equipment consists of the following items:

- first-aid kits

- emergency axe

- crew call signalling

- inflatable life jackets (ASZh-63P) (if installed)

- inflatable life rafts (SP-12) (if installed) and emergency packages (NAZ-7) (if installed)

- emergency exits

- - -emergemy kcatop transmitter ELT-10 oc POLNTER 3QOOlif installed) - - - - - - - - - -

- inscriptions explaining the emergency exit opening sequences and markings of the places where

the emergency axe and first-aid kits. are located, as well as of the cut-out area for an emergency

manhole

- hand fire extinguishers

i f installed

FIG. 12-29 EMERGENCY EQUIPMENT

AIRCRAFT TRAINING MANUAL L 410 UVP-El E9. €20


- - - - - - - - - - -


CHAPTER 13

POWER PLANT

13.1. General

13.2. Fixing of the engine

13.3. Fireseals

13.4. Drains _

13.5. Engine controls

13.6. Emergency shutdown

13.7. Indicating

13.8. Engine starting



AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9. EZO

13.1. GENERAL The power plant of the L 410 UVP-E airplane is the M 601 E engine with the V 510 propeller and

instruments necessary for their operation. There are two engines on the aeroplane.

NOTE: The aircraft can be also equipped with M 601 E-21 engine. This engine has increased

power higher ambient temperatures. The installation of this engine is carried out in

accordance with the INFORMATION BULLETIN No. L 410 UVP-E1055b.

The engine is of two shaft, free turbine reverse flow design. The engine ist composed of two basic

parts- gas generator and power free turbine driving the propeller through the reduction gearbox. - - - - - - - - - - - - - - - - -

Water injection system is available, intermediate contingency mode can be used, the engine can

be started up on the g round or up to 4 km altitude.

The engine is suspended by means of three spring pins on the engine bed which is fitted into the

wing hinges. The engine can be started up by an electrical starter and torch igniters. For M 601 E

engine protection an additional device is built that protect the engine against exceeding the critical

parameters. This device is double level system of combined limiters. ,

Another additional feature is an adjustable stop for maximum take-off rating.

The engine fuel system is a low pressure system and consist of an fuel control system whose

main part is a gear pump and a fuel controller.

The engine oil system is a pressure circulation system with gear pumps and an integrally-arranged

oil tank inside the gear box. Oil turned back from the engine is cooled in an oil cooler that is

located in_theenginenacelle. Oil tempe~ature regulation is-controlled b y - an - thermostatic - - valve in - - - -

the oil cooler. To achieve optimal oil temperature at very low air ambient temperature th e oil

cooler is equipped with an adjustable valve and engine cowlings are equipped by slots with

adjustable screens.

Chip detection in oil system and minimal quantity of oil signalling are installed at the engine.

Technical data

Engine power ratings

Maximum take - off

Equivalent power 595 kW (809 HP)

Shaft power ' 560 kW

Basgemrator speed - - - - - - - - - - - - - max98A%

Propeller speed 2 080 RPM + 20 RPM

Torque max. 100%

Max. fuel consumption 39591 eq. kW1h 12909 Ieq. HPIh

I TT max. 735%

Single uninterrupted run max. 5 minutes

Total run (including max. take-off rating with

water injection and intermediate continguency rnax. 4% out of operating time

NOTE: *Is maintained up to temperature of +27OC and pressure 97,325 kPa for this rating with

water injection

Intermediate Continaencv a~~ l icab le in case of enaine failure


Shaft power min 560 kW (min 760 HP)

Gas generator speed rnax. 98.6%

Propeller speed 2 080 RPM + 20 RPM

Torque rnax. 100%

Max. fuel consumption 3959 Ieq. kW1h (290 Ieq. HPIh)

ITT max 710°C

Single uninterrupted run 1 hour max 3 times during operating time

Total run at this rating (including at max. take-off

and at max. take-off with water injection) max. 4% out of the operating times

Under non - standart conditions (at 97.325 kPa + 30°C)


Shaft power 515 kW (700 HP)

Gas generator speed

ITT

Max. take-off with water iniection

Under non-standard conditions (at 97.325 kPa + 33OC)

Equivalent power

Shaft power

Gas generator speed

Propeller speed

Torque

ITT

max. 100%

max. 750°C

560 kW (760 HP)

min. 530 kW (min 720 HP)

max. 100%

2 080 RPM + 20 RPM

max. 100%

max. 735OC


Atmospheric temperature

Single uninterupted run

Total run at this rating (including max.

take-off and Intermediate Contingency)

Maximum Continuous

Equivalent power

Shaft power

Gas generator speed

Propeller,speed

Torque

Max. fuel consumption

Uninterrupted run

Maximum Continaencv (without or with water iniection)

Under non-standard conditions (at 97.325 kPa + 35OC)

Equivalent power

Shaft power

Propeller spsed

Max. torque (for information)

Gas generator speed

Acceleration time from initial take-off rating to Maximum

Contingency

Single uninterrupted run

Information ITT value

Idle run

Gas generator speed

above 23% on high altitude airports

from 150C at altitude 800-1 300 m

from 10% at altitude over or equal

to 1300 m

max. 1 minute

max. 4% out of operating time

521 kW (708 HP)

min. 490 kW (min 666 HP)

max. 96.5%

1800 up to 1900 RPM

100%

4109 Ieq. kW/h / 3019 /eq. HP/h

unlimited

max. 690°C

630 kW (857 HP)

595 kW (809 HP)

2080 RPM + 20 RPM

106.5%

102%

max 1 sec.

max 2 minutes

60 + 3% on ground for altitude not excceding 2.5 km ISA

60% - for altitude above 2.5 km ISA

rnax 76% for altrtude 4.2 km ISF?

max. 78.2% for altitude 6.0 km ISA

Accelerations

Minimum initial gas generator

Speed

Accelerations time from initial rating to 95% of max. take-

off rating (99% of gas generator speed)

- when advancing lever at 1 sec

- when advancing lever within 3 sec

Gas generator speed overshoot

(rnax. 3 times before stabil~zation)

Torque overshoot

(rnax. 3 times before stabilization)

Propeller speed overshoot

Reverse mode

Shaft power

Gas generator speed

Propeller speed

70% for altitude 0 - I km ISA

75% for altitude 1 - 4km ISA

5 sec for altitude 0 - 4km ISA

6 sec for altitude 4 - 6km ISA

rnax 73S°C

rnax 101 .OOh

rnax 106%

max. 2 140 RPM (3 times

before stabilization)

max. 2 220 RPM at aborted

landing (stabilized within 6

sec.)

min 337 kW (458 HP)

max. 97%

rnax 1 900 RPM

Max. inter - turbine temperature max 710°C

Single uninterrupted run max 1 minute

NOTE: Above mentined data corespond to power values of the engine without engine cowlings

and without customer bleed and customer loads.

Oil quantity

in the oil tank

in the oil system

Oil temperatures

mln 5.5 litresmax 7 litres

I 1 litres

- at start up min -20°C

- for run up and normal engine run min -20°Cmax +85OC

- at Intermediate Contingency Rating max +95OC (for a period of

1 hour

13.2. FIXING OF THE ENGINE

The fixing of the engine is with the help of the engine bed, which is fixed in the front and rear

suspensions of the engine bed placed on the wing. The engine bed is assembled from the

supporting ring of the engine bed and from the engine bed's struts.

The supporting ring, which is a part of the engine - see Fig.13-1 is manufactured from a steel pipe.

On the upper side and on the flanks, in the place of the fixing of the engine, the supporting ring is

strenghtened by boxes and is provided with suspensions for the gripping of the engine bed's

struts and with bushing for the gripping of the engme. Due to the manufacturing inaccuracies of

the engine bed and the necessity of setting the engine into the correct plane with the airframe

there are bearings for pins of the engine bed's struts in the suspensions.

The struts of the engine bed are manufactured from steel pipes, having the forks welded at their

ends.

The upper and side struts have a fixed length, the lower ones are adjustable.

The engine is fitted to the supporting ring of the engine bed in three points by means of absorbers.

The elastic fixing of the engine secures reliably vibrations absorbing and prevents their transition

on the aeroplane's construction.

The setting of the engine on the aeroplane in all planes is carried out by adjustable lower struts.

13.3. FIRESEALS

To protect against fire propagation to the whole engine the area of the engine nacelles is divided

by fire bulkheads into 4 zones. The construction of the fire bulkheads and their sealing prevents

the propagatim of the fire between the individual sections.

The fire bulkheads are manufactured from metal sheet of fire resistant steel. All openings in the

bulkheads are sealed by bushings or by sealing from, fibre-glass cloth. Bes~des this the bulkheads

and walls are fastened to the engine na celle by a sealing pennon (Firesol Arreohead), which is a

part of the engine nacelle. The front and rear fire (engine) bulkhead is a part of the engine.

13.4. DRAINS

The oil and f ~ e l system of the power plant is equipped with drainage piping for the fuel and oil

waste which originates during the operation and servicing of M 601 E engine. This waste is led

under the engine cowling beyond the aeroplane contour. Here is terminated the venting of the oil

tank, the oil waste from the filling neck of the oil tank, the fuel waste from the combustion

chamber, tht waste from the fuel pump and waste from the fuel governor, oil waste from the

auxiliary pump, fuel w aste from the pressure switch, oil waste from the torque transmitters.

FIG. 13-1 THE ENGINE BED

(1) Engine bed's ring, (2) Upper suspension, (3),(4) Side suspension,

(5),(6) Bushing, (7) Right support, (8) Left support, (9) Side plate, (10) Side

plate, (1 1) Lower strut, (12) Side strut, (13) Side strut, (14),(15) Upper strut,

(16) Pin, (17) Nut, (18) Shim, (19) Split pin, (20) Bridging, (21) Screw,

(22) Shim, (25) Front suspension of the engine bed

I - Detail of the mounting of the engine on the engine bed

(26) Absorber, (27) Shim, (28) Nut, (29) Split pin, (30) Screw, (31) Absorber's

box, (32) (33) (34) Screw, (35) (36) Shim, (37) Lid of the absorber's box

AiRCRAFT TRAINING MANUAL L 410 UVP-E, E9, E20

FIG. 13-2 THE DIVISION OF THE AREA OF THE ENGINE NACELLE INTO THE FIRE

PREVENTION ZONES

I, 11, Ill, IV - the fire prevention zones

(1) Front fire bulkhead (on the engine), (2) Rear fire bulkhead (on the engine),

(3) Front fire bulkhead, (4) Rear fire bulkhead, (1) - (9) Bulkheads of the

engine nacelle.


13.5. ENGINE CONTROLS The engine controls consist of the following systems:

- power control system,

- emergency control system.

The former consists of engine control and propeller control The normal engine control system is

based on a fuel control unit and a system of limiters which prevent the engines from exceeding

their principal operating parameters. The emergency engine control system is essentially a direct

control of the flow of fuel into the engines by means of a fuel shut-off valve excluding the fuel

control unit and appropriat limiters.

The engine controls are mechanical. The control levers are situated on the control panel. Control

movements of the control levers are transmitted by tie rods and a cable system runnmg over an

arrangement of pulleys in the fuselage and wings. The cable system represents a closed circuit

starting at the end pulleys on an indepenaent countershaft below the control panel in the pilots'

cabin and ending at the end pulleys in the engine nacelles.

The diameter of the cables is 2 mm. Each cable circuit is connected by means of turnbuckles. The

turnbuckles on all the cab1 es are situated In the vert~cal control channel.

The cables are prestressed to 196 +I- 10 N (20 +I- 1 kp).

The engines and fuel shut-off valves are controlled by end pulleys on vertical countershafts via tie

rods and auxiliary shafts, the appropriate control cables then run to the controls proper located on

the engine components.

The control of propeller governor is provided by end pulleys on veritical countershafts via flexible

tie rods on the engines connected to propeller governor control levers.

The control of fuel cocks is provided by tie rods connecting the end pulleys and fuel cock control

levers.


FIG. 13-3 SCHEMATIC DilAWING OF THE CABLE SYSTEM OF THE ENGINE

CONTROLS - - - (7) Propetler governor controtcaHe system, (2) Engine powereonkol cable -

system, (3) Fuel cock control cable system, (4) Fuel shut-off valve control cable

system

a - control panel as far as vertical control channel,

b - fuselage roof between frames No.7 and 11 and as far as the third wing rib,

c - third to eighth wing ribs as far as engine nacelle

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9. E2O

FIG. 13-4 ENGINE CONTROLS

(1) Engine control, (2) Propeller control, (3) Fuel shut-off valve control, (4) Fuel

cock control.

I - Engine control levers with tie rods and countershaft

I1 - Countershafts

Ill - Countershafts

IV - Bracket with pulleys and guiding elements

V - Engine controls in the RH engine nacelle

AlRCRAFT TRAINING MANUAL L 410 UVP-E, E9. E20

Engine control

The engine control levers are situated on the front control panel sharing their countershaft with the

propeller control levers. The engine is controlled by the two levers labelled POWER and situated

to the left. There is a friction lever located beside which is used to arrest the power control levers

in a given position. The power control levers are provided with guiding pins coupled to the handles

by means of tie rods. The guiding pins travel in the gate slots: the gate is fixed to the fron t control

panel. The slots in the gate are designed to correspond to the basic engine ratings. The handles

with the tie rods allow the control levers to be shifted over the stops in the gate slots.

The engine (power) c~ntrolLlevers a-re~ntendgd for - - continuous control of the engines from the - - - - - - - - - - - - - - - - -

ground idle run to the take-off power. If their action is combined with that of the propeller governor

(to which the power control levers are connected/, they allow for changing the engine power from

the ground idle run to the maximum reverse thrust. If one of the engines fai!s daring take-off, it is

possible to set the maximum contingency power.

The slot in the control lever gate has three sections: the middle one is intended for controlling the

engine in the 100% n~ (control lever shifted forward) -to-idle run (control lever at the idle run

suop) range.

The lever can be moved into the rear part of the slot no sooner than after unlocking the reverse

thrust stops and using the handle to lift up the gu iding pin. As the lever is being shifted farther

backward, the BETA regulation system sets the propeller blades at a smaller angle so that the

minimum thrust value is achieved. As the lever continues its backward travel, the propeller blades

are-belng gradually set to rregative angles, the-power being inweased up to themaximum reverse

thrust at which the lever rests against the rear stop.

The BETA regulation system is described in detail in the V 510 Propeller Maintenance ManuaLThe

control le ver can be shifted into the front section of the slot using the handle which lifts up the

guiding pin. Shifting the lever forward results in the engines operating at the maximum

contingency power. The max. contingency power can be set by forcing the lever against a spring

stop to its extreme forward position. Forcing the kver beyond the spring stop results in the limiter

being turned off automatically. Consequently, it is necessary to watch the instruments and prevent

the engine from exceeding 106. 5 % torque, irrespective of other engine parameters

(see L 410 UVP-E, E9, E20 Flight Manual).

CAUTION: FORCING THE CONTROL LEVER BEYOND THE SPRING STOP RESULTS IN THE

ELIMINATION OF THE SPRING STOP RESISTANCE: THE CONTROL LEVER CAN - - - - - -

- - - - - - - - - - - - - -

THUS BE SHIFTED FREELY FROM IDLE RUN TO MAXIMUM CONTINGENCY -

POWER POSITION. THE STOP FUNCTION CAN BE RESTORED BY A

REPRESENTATIVE OF THE MANUFACTURER BY MEANS OF A SPECIAL

FIXTURE AND AFTER THE ENGINE HAS BEEN CHECKED.


Engine controls adjustment

The design and adjustment of the engine controls have been focused on achieving the maximum

control sensitivity in the maximum forward thrust range (control lever shifted to its extreme forward

position, or near to it, approximate speed 90 to 100% n ~ ) .

A carefull selection of radii of the control levers, lengths of the tie rods in the engine nacelles and

under the front control panel and a suitable setting of the end positions of the engine control

leverage with respect to the tie rods res ulted in achieving an angular shift in the maximum

forward thrust range which is twice than that in the idle run region for the same speed

increase.This setting must be observed every time the engine control system is being adjusted.

The airplane is equipped with an adjustable maximum take-off power stop. Instructions

concerning the use of the adjustable maximum take-off power stop (including its adjustment chart)

are presented in the L 410 UVP-E, E9, E20 Flight Manual.

FIG. 13-5 ENGINE CONTROL

I - Engine (power) control with tie rods and countershaft

Ill - Countershafts

I l - Countershafts

IV - Bracket with pulleys and guide elements

V - Engine control in the right - hand engine nacelle


Propeller control

The propeller control levers are situated on the front control panel sharing their countershaft with

the engine control levers. The propellers are controlled by the two levers situated to the right.

There is a friction lever located beside which is used to arrest the propeller control levers in a

given position.The control levers are provided with guiding pins coupled to the handles by means

of tie rods. The guiding pins travel in gate slots: the gates are fixed to the front control panel . The

slots in the gates are designed to correspond to the basic propeller settings. The handles with the

tie rods allow the control levers to be shifted over stops in the gate slots. The propeller control

levers are intended for setting the propeller governor lever to a position corresponding to the

propeller RPM selected. The propeller governor maintains the propeller speed setting by adjusting

the angle of setting of the propeller blades to make up for changing flight conditions.The propeller

control lever has three basic positions secured by the shape of the gate slot. The control lever at

the forward end in the lower section of the slot corresponds to the propeller speed equal to

2.080 rpm, which is employed during the take-off. The propeller speed can be reduced to

1,700 rpm by shifting the lever backward (rear end of the lower slot).The upper (rear) slot can be

used to set the control lever into the FEATHERED position in which the propeller blades are set

correspondingly.

Fuel shut-off valve control

The control levers of the fuel shut-off valves are situated on the front control panel sharing their

countershaft witn the levers controlling the fuel cocks. The levers controlling the fuel shut-off

valves are the two inner oneswith the control levers in the rear position, the fuel supply to the

engine is cut off whereas in the middle position (on the gate step) it is open. The control levers

can be locked in the rear position. By tilting the control levers to the right and shifting them

forward an emergency mode is established. The emergency engine control mode can be

employed in case of fuel control unit failure after switching on the ISOL. VALVE LH (RH) circuit

breakers on the overhead panel. In such case, the ISOLATION VALVE signalling cell on the

central warning display lights up.The emergency engine control allows for throttling the fuel supply

to the engines directly by means of a shaped needle in the fuel shut-off valve (part of FCU). The

pilot must carefully check essen tial operation parameters of the engines, i.e. torque, inter-turbine

temperature, generator and propeller speed, and prevent them from being exceeded, as the

limiters are put out of operation when the emergency control mode is used.The initial position of

the control levers for starting the engines in the emergency control mode is marked by a red line

on the control levers and the gate.

NOTE: A schematic drawing of the fuel shut-off valve control is shown in fig. 13-6.


27 V

LH ENGINE ELECT'RO RH ENGINE

MAX. TAKE-OFF MAX. REVERSE MAX. CONTINGENCY

I D L E E

VALVE

/ - FUEL CONTROL UNIT

FIG. 13-6 SCHEMATIC DRAWING OF ENGINE CONTROL


@@B OIL OIL yp


FIG. 13-7 PROPELLER CONTROL


Il - Countershafts

Ill - Countershafts

IV - Bracket pulleys and guide elements

V - Engine controls in the right-hand engine nacelle


FIG. 13-8 STOP OF PROPELLER CONTROL LEVERS

(1) Propeller control lever, (2) Propeller control lever guiding pin, (3) Gate with

guiding pin slot, (4) Forward stop of propeller control lever, (5) Rear stop of

propeller control lever,

I - Low propeller blade setting angle, maximum controllable speed 2,080 rpm,

II - High propeller blade setting angle, minimum controllable propeller speed,

1 1 1 - Start of propeller blade resetting into feathered position,

IV - FEATHERED position stop on the propeller governor,

V - Arrestment of propeller in feathered position.


FIG. 13-9 FUEL SHUT-OFF VALVE CONTROL


Il - Countershafts

Ill - Countershafts



FIG. 13-10 LOCKING OF FUEL SHUT-OFF VALVE CONTROL LEVERS

(1) Lock - modified, (2) Lever, (3) Bushing, (4) Coupling, (5) Bolt with nut,

washer and split pin, (6) Pin with washer and split pin.

Limiters

The purpose of the limiters installed on the engines is to prevent the engines from exceeding their

maximum parameters thus ensuring their reliable operation.

The system covers gas generator speed, propeller speed, inter-turbine temperature, propeller

torque and inter-turbine temperature gradient limitation. The signals of the parameters listed

above are transmitted to the LUN 5224.05 lntegrated Electronic Limiter Unit by the corresponding

sensors.

Except for the torque limiter transmitter, the remaining transmittersare cammon for-both the - - - - - - - - - - - - - -

- - - - - - - - - - - - - - - - -

- - - - - - - - -

integrated electronic limiter unit and the indicators in the pilots' cabin.

If the maximum permitted value of any of the parameters being monitored is exceeded, the

appropriate transmitter sends a signal to the lntegrated Electronic Limiter Unit. The IELU

processes the signal and sends an output signal the magnitude of which depends on to what

extent the particular parameter has been exceeded and how many parameters have been

exceeded. The out put signal is fed to an electrohydraulic converter which, together with the

LUN 6590.05-8 fuel control unit, throttles the fuel supplied to the engine.

Such situation and, hence, an intervention of the lntegrated Electronic Limiter Unit, are indicated

by the IELU INTERVENT signal cell lighting up on the central warning display. The lighted up IELU

signal cell indicates that the lntegrated Electronic Limiter Unit is out of operation.The IELU TEST

pushbutton located on the left-hand control panel can be used to check the lntegrated Electronic - - - - - - - - - - -

- - - - - - - - - - - - - - -

Limiter Unit (at a reduced ITT). - - - - - - - - -

- - - - - - -

- - - - - - - - - - -

The operational readiness of the IELU is signalled by means of two signal lamps located on the

right-hand control panel.

Operation limiters

The system of limiters is fed by 27 V DC current via the IELU LH (RH) circuit breakers on the

overhead panel. If some of the parameters monitored by the IELU exceeds the maximum

permitted value, the signal of the appropriate transmitter is processed by the IELU which sends a

control current signal into the electrohydraulic converter (EHG). The value of the current is

proportional to the extent of parameter exceeding. The EHC converts the electric signal into an

intervention into the fuel con trol unit: the fuel metering needle is throttled. The throttling is again

proportional to the magnitude of the transmitter signal, or the sum of signals sent by several

transmitter at a time if more parameters are-exceeded: The fuel supply-throttling results in a drop- - - - - - - - - - -

- - - - - - - - - -

of engine power and, consequently, in a reduction, of the paraneters to an acceptable level.

In order to prevent engine power oscillation due to IELU intervention, higher parameter overruns,

the drop of gas generator speed at the IEL U intervention is controlled by the LUN 5223-8

generator speed derivator.

As soon as the drop rate attains an undesirable value, the derivator interrupts the control signal for

a moment and the electrohydraulic converter stops throttling the fuel supply.

This means that the speed drop is momentarily stopped or reduced. Repeated actions of the

derivator guarantee that the required parameter values are achieved without oscillations and are

quickly stabilized.

A sirnilar principle is used to control the gradient of inter-turbine temperature (dlTTldt) and to

maintain it at a value lower than 140° - 1800Clsec. If the temperature gradient reaches these

values, the IELU throttles the fuel supply, which results in a gradual increase of the temperature

and gas generator speed thus preventing the engines from exceeding temperature limits when

being started.

After the engine is started and the torque reaches the value of approx. 17%, the LUN 3280-8

aatomatic feathering pressure switch turns off th e function of this channel not to impair the engine

acceleration.

There are two IELU intervention levels, one consisting in reducing the torque to a value not lower

than 70 +I- 5%, while the other can reduce the torque down to the flight idle run value ( n ~ = 60%).

The first level guarantees that there will be no excessive power drop should the limiter system fail

and send a false signal. This is why the second level limitation function is blocked by a signal from

the KRA 405 or A 037-1A radioaltimeter receive r and a signal from the limit switches of the

mechanical locks of the extended position of the main landing gear at height below 700 m.To

provide for the required compressor stall margin at low temperatures (less than -1 5OC), the IELU

reduces the maximum gas generator speed in compliance with the ambient air temperature. This

limitation occurs most frequently at low temperatures and high altitudes where the torque is not a

limiting value. The ambient air temperature is measured and sent to the IE LU by the electric

resistance ambient air temperature transmitter, P5 (7), located in the engine nacelles.

If reveised thrust is employed, the propeller blades may be unloaded during their resetting (in the

region of the zero angle setting). To prevent the propellers from overspeeding even if the

lELU LH (RH) circuit breakers are not switched on the ))Vc microswitch on the LUN 7816.8

propeller governor activates the IELU automatically as soon as the propeller blades start resetting

below minimun flight angles.

In case of automatic feathering the LUN 3280-8 automatic feathering pressure switch, which

transmits signals for the extension of the ABC tab and for the switching of the automatic

feathering circuit, also transmits a signal to the TKE 54 PODG relay of the other engine to turn off

the IELU power supply. This action is indispensable if the maximum contingency power is to be

employed, as the torque and inter-turbine temperature involved are already beyond the IELU-

controlled limits.

A shortened test of the IELU serviceability (only the temperature channel) is performed by

pressing down the IELU TEST pushbutton on the test panel of the left-hand control panel. The

IELLJ is thus activated and the ITT limit is reduced so that the IELU intervention capability can be

checked at a lower engine power.

The remaining functions ( n ~ , np, torque) of the unit can be checked by resetting the channel test

selector on the IELU body to an appropriate position.


FIG. 13-1 1 SCHEMATIC DRAWING OF THE SYSTEM OF LIMITERS

(1) LUN 1333.1 2-8 gas generator speed transmitter, (2) LUN 1333.12-8

propeller speed transmitter, (3) LUN 1377-8 inter-turbine temperature

transmitter, (4) LUN 1476-8 torque pressure switch, (5) KRA 405 radioaltimeter

or A 037-IA, (6) KPS (7) U limit switch

(7) LUN 5223-8 generator speed derivator, (8) LUN 5260.05 integrated

electronic limiter unit, (9) LUN 2476-8 electrohydraulic converter,

(10) LUN3280-8 automatic feathering pressure switch, (1 1) NVU microswitch

on the LUN 7816-8 propeller governor, (12) P5 (7) ambient air temperature

transmitter, (13) KNR pushbutton IELU TEST, (14) IELU LH, RH circuit

breakers, (15) BETA RANGE signal cell, (16) IELU signal cell, (17) SLC 51

signal lamp IELU OPERATIVE, (18) IELU INTERVENT signal cell,

(19) Channel test selector on the LUN 5260.05 IELU body or LUN 5260.04,

(20) TKE 54 PODG relay.


@ LH ENGINE

I IELU INTERVENT

RH ENGINE 0

IELU I

IELU OPERATIVE I I

FIG. 13-1 2 SCHEMATIC DRAWING OF IELU IN-AND-OF-OPERATION SIGNALLING

13.6. EMERGENCY SHUTDOWN

The emergency engine shutdown system of the arplanes consists of two mechanically controlled

subsystems, the shutdown system proper, which is comnon for normal and emergency

shutdowns, and a system cutting off the supply of fuel in front of the fire wall, which is necessary

in case of engine fire.

The engine shutdown is accomplished by shutting of the fuel supply to the engine by a fuel shut-

off valve needle: its control lever is moved against its rear stop in the gate.

The control levers are situated in the pilots' cabin, in the central part of the front control panel.The

shutting off of the fuel supply to the area in front of the fire wall is prov~ded by fuel (fire) cocks.

Their control levers are located on the front control panel, on the same countershaft as the fuel

shutt-off valve control levers, offside of the latter. In the extreme rear position FUEL - SHUT, the

fuel cock is shut, in the front position FUEL - OPEN, it is open. A torsional spring of the fuel cock

two-position plate valve presses and fixes the control lever in its extreme position, of the latter is

moved behing the central ))dead(( position. The movement of the control lever is transmitted to the

fuel cocks by means of levers and tie rods in the pilots' compartment (in a way similar to the fuel

shut-off valve control levers) and pulleys and cables from the pilots' cabin to the wing. The last

part of the transmission train consists of a lever and a tie rod.

13.7. INDICATING

The engine indications include those systems the ~ndicators of which inform the pilots on the

engine function. Especially if the maximum engine settings are exceeded as a result of a failure.

The indication systems warn the pilots that the engine is working in an inadmissible mode. In such

cases, the pilot makes an intervention into the engine setting bringing its working parameters back

below the maximum tolerable limits.

Power indicating

The term ))power output indication (check)(( comprises measurements of the following

parameters:

- propeller shaft torque

- gas generator speed

- propeller speed

The torque measurement system measures the propeller shaft toraqe. It consists of the

LUN 1540.02-8 torque transmitter, the LUN 1539.02-8 torque indicator and connecting lines.The

system of gas generator speed indication measures the gas generator speed. It consists of the

LUN 1333.12-8 tachogenerator, the LUN 1347.XX-8 dual (gas generator) speed indicator and

connecting lines.The propeller speed indication system measures the propeller speed. It consists

of the LUN 1333.12-8 tachogenera tor, the LUN 1333.12-8 speed indicator, the LUN 1348.XX-8

dual (propeller) speed indicator and connecting lines.


FIG. 13-1 3 FUEL COCK CONTROL

I - Control levers of the engine regulating system with tie rods and counter shaft

Il - Countershafts

Ill - Countershafts




I CONVERTERS I

I I - TORQUE INDICATOR DUAL /GAS GENERATORI SPEED INDICATOR

,-,--1 L,,,,

DUAL /PROPELLER/ SPEED I x D l c A T o R 1 i I I

TORQUE TACHOGENERATOR

I I FIG. 13-14 SCHEMATIC DRAWING OF THE POWER INDICATING SYSTEM I

I I

I I I 13-28


Temperature indicating

The temperature indication is intended to measure the inter-turbine temperature. The system of

inter-turbine temperature measuring measures the temperature between the gas generator

turbine and the propeller turbine. The measurement takes place in front of the propeller turbine

stator. The system conslsts of the LUN 1377-8 inter-turbine temperature transmitter the

LUN 1370.02-8 inter-turbine temperature indicator and connectmg lines.

I LUN 1377- 8

FIG. 13-15 SCHEMATIC DRAWING OF THE TEMPERATURE INDICATION SYSTEM


13.8. ENGINE STARTING

The engine is started by means of an starter-generator and a semiconductor ignition unit. The

starter-generator and the ignition unit with spark plugs are mounted on the engine and are its

integral part. The spark plugs ignite fuel in the main flow path. The devices listed above are

connected to the electrical system of the aircraft. The starting system provides for fault less

starting down to -20°C. If the ambient temperature drops below -20% heating is required.The

starting operation c an be made by using an external source of electricity or the on-board storage

battery. Starting conditions and procedures are described in the M 601 E Engine Maintenance - - - - - - - - -

- - - - - - - - - - -

- - Manual-and the L-410 UVP-'t,E3,-E20 Airplane Flight Manual.

Starting system operation

The circuit-breaker ENGINE STARTING LH, RH turns on and safeguards the starting circuit of the

left (right) engine. The ignition is checked by the IGNITION I,'II (LH, RH) change-over switch. If

the change-over switch is in position I, the ignition system No. I is in function and sparking can be

heard from the left (right) engine. If the change-over switch is moved to Position II, the ignition

system No. II is in function and sparking can also be heard from the left (right) engine. There is n

o fuel injection during the ignition system check.

On pressing down the ENGINE STARTING LH (RH) pushbutton, the following operations are

carried out automatically:

- the starter-generator of the engine left (right) is switched on, - - - - - - - - -

- - - - -

- - - - - - - - - - - - - -

; the-fwo ignitkn systems of the left (right) engine are switched on,

- the interrupted fuel injection into the torch igniter LH, RH is switched on,

- the time relay is activated and terminates the starting cycle automatically after 22 seconds

The starting is signalled by the ENGINE STARTING signal cell. Individual ignition circuits are

safeguarded by 4 A fuses. the signalization circuit is safeguarded by 0.5 A fuses.

When starting the other engine using the airplane batteries and the running starter-generator the

current of the starter-ge?srator is limited to 350 A. The cirsuit limiting the current of the starter-

generator is built into the voltage regulator. By pressing down the DRY MOTORING RUN LH (RH)

pushbutton, only the st arter-generator of the appropriate engine is turned on for 22 seconds, this

situation being signalled by the ENGINE STARTING signal cell.


TIME RELAY - - -

- - 1 INTERRCPTER

0 . . . I&--- I 1 : : . . 11 : :

IGNITION SYSTEM I I I -

----

i ' $ SPARK PLUG I I / I , * / , , \> -

STARTER GENERATOR

FIG. 13-16 SCHEMATIC DRAWING OF THE CRANKING-UP SYSTEM




CHAPTER 14

PROPELLERS

14. I. General

- - -7 4.2.-Propellers controlling

14.3. Indication



14.1. GENERAL

The propeller is thrust, five-blade, variable pitch with feathering and reverse positions, clockwise

rotation with exchangeable duralumin blades.The propeller blades pitch is controlled by hydraulic

servo which is controlled by speed governor that maintains the constant speed within normal flight

modes. The propeller is equipped with a device allowing for normal control at small and negative

pitch angles ))so called Beta regulation((.

Technical data

Propeller external diameter 2 300 mm

Number of blades 5

Way of speed regulation speed governor within 1 700 - 2 080 RPM

Feathering both automatic and manual

Propeller speed of reverse mode, ISA, V = 0 max. 1 900 RPM

Pitch range 103O 30

14.2. PROPELLERS CONTROLLING

The propellers are controlled by a system of cables and rods, which are connected to the propeller

control levers in the cockpit.

The setting of the propeller blades is carried out by a servo system, controlled by a speed

governor, which maintains constant propeller speed in the range of normal flight conditions.

The propeller is equipped with a mechanism for manual operation during small and negative

angles of setting of the propeller blades which is called the ))beta control((.

Into the hydraulic circuit of the electrohydraulic governor the propeller speed limiter is inserted,

which in case of a fault of the speed governor prevents the increase of the propeller speed over

the rated limit. In the propeller hub a secondary stop is placed, which will prevent accidental re-

setting of the propeller blades into negative angles in case of a fault or the disconnection of the

speed governor rod in the system ))beta controkc.

The propeller is equipped with a system of automatic, manual, e mergency and checking

feathering.

Operation

If a serious fault of the engine occurs, which could affect the safety of the flight, it is necessary to

put this unit out of operation and to re-set the propeller into the feathering position. For this

purpose the propeller is equipped with gear for feathering, which is according to mode either

automatic or manual, emergency and checking.

AIRCRAFT TRAINING MANUAL i 41 0 UVP-E, E9, E20

Automatic feathering

Automatic feathering is intended for flight conditions, during which the engines power setting is set

above the rated limit (89% ng or 93% ng in dependence on the adjustment of the terminal

switches for emergency of the automatic feathering circuit D 701 (M 304) for summer or winter

conditions).

The automatic feathering c~rcuit is switched on by two circuit breakers FEATHERING

/AUTOMATIC BANKING on the overhead panel and the switches AUTOMATIC FEATHERING on

the middle control panel. The sy stem will be put into emergency after moving both engines

control levers over the value corresponding to 89+/-1% resp. 93+/-1% ng, which will connect both

D 701 (0 613-3A-U2) (M 304) terminal switches. By connecting the terminal switches the

ZO-4s (8) (M 313, M 314) delay circuits will be put into operation which will connect after

5 - 7 seconds the TKE 52 PODG (M 321, M 322) relay.

Through the connected contacts of these relays voltage will be brought to terminal connector 20 of

the V 139 termina I board and to terminal connector 20 of the V 140 terminal board. From the

terminal connector 20 of the V 139 terminal board the voltage will be brought across the

connected AUTOMATIC FEATHERING switch, the disconnecting contacts 1 and 2 of

TKE 54 PODG (M 305, M 306) relay to terminal A of TKE 54 PODG (M 305, M 306) relay coil.At

the same time voltage will brought into the AUTOM. FEATHERING signal cells and both green

coloured signal cells on the CWD will light up.

From terminal connector 20 of the V 140 terminal board voltage will be brought on the movable

contact of the LUN 3280-8 (M 310) automatic feathering pressure switch -jack bush No. 6. For

example at the failure of the left (right) engine the resulting drop of oil pressure for the measuring

of Mk, which is led into the LUN 3280 (M 310) automatic feathering pressure switch, under the

value 0.222 MPa (cca 24% Mk), causes the 1st degree of this pressure switch will be connected,

which will cause the closing of TKE 54 PODG relay (M 305, M 306), which will carry out:

- disconnection of the right engine's automatic feathering circuit (it will cancel readiness).

- switching off of the AUTOM. FEATHERING on the CWD ENGINE R.H. green coloured signal

cell.

- the supply of voltage on terminal connectors 9,6 A of TKE 54 PODG (M 307) relay and on

terminal connector V of the LUN 7880.01-8 (M 312) left solenoid valve.

- the bringing of the signal into the ABC circuit

If the oil pressure will drop under the value of 0.162 MPa (about 18% Mk) the second degree of

the automatic feathering's pressure switch will connect (LUN 3280-8 (M 31 O)), which will cause

closing of TKE 54 PODG (M 307) relay, which will carry out:

AIRCRAFT TRAINING MANUAL L 410 UVP-E, E9. EZO

- the switching off of the AUTOM. FEATHERING green coloured signal cell on the CWD ENGINE

L.H.

- the opening of the electrical circuit of IELU feeding of the right engine, which is signalled by

lighting up of the IELU yellow coloured signal cell on the CWD ENGINE RH.

- closing of the electrical circuit of the insulating valve (M 27) of the left engine, which is signalled

by the lighting of the ISOL. VALVE yellow coolured signal cell on the CWD

ENGINE LH.

- closing of TKE 54 PODG (M 339) relay, which will put into operation the LUN 2601-8 (M 337)

time relay for the time of 13 - 15sec. This timerelay together - - with - - - - the closed contacts of - - - - - - -

TKE 54 PODG (M 339) relay will carry out closing of TKD 501 DOD (M 343) contactor and this

will put into operation the LUN 7840-8 (M 345) auxiliary pump, which is fed from the AZRGK 40

(M 341) circuit breaker.

During the time of operation of the auxiliary pump (13 - 15 sec.) the 0.7s LUN 1492.04 (M 347)

pressure switch will close, which will cause the lighting of the FEATHERING PUMP yellow

coloured signal cell on the CWD ENGINE LH.

- closing of the of LUN 7880.01-8 (M 312) (left) solenoid valve actuator electromagnet, which will

cause the setting of the propeller blades into the feathering position.

NOTE: The TKE 54 PODG (M 339) relay remains closed after finishing the feathering cycle and

with its connected contacts 11, 12 it feeds the automatic feathering circuit, for the - - - - - - - - - -

- securing of feeding ofthe L-UN7880.01-8 (M 312) solenoid valve actuator left and for - - - - -

preventing the repetition of the feathering cycle again. This state may be cancelled only by

switching off the FEATHERING IAUT. BANK overhead circuit breakers on the overhead

panel.

The automatic feathering circuit may be put into a state of readiness again by switching

off and by switching on again the FEATHERINGIAUT. BANK circuit breakers on the

overhead panel - it is carried out after successful starting of the engine or after carrying

out checks of this circuit.

The exact values of adjustment of D 701 (B 613-3A-U2) (M 303, M 304) terminal switches

of automatic feathering readiness is carried out according to the Maintenance Manual for

the M 601 E engine.

The exact values of switching the 1st and second stages of the au tomatic feathering - - - - - - - - - - -

LUN 3280-8-(M3@9, M 31 0) switches are mentioned in thece~tifisate of the instrument - - - -

and the checking of the adjustment is carried out according to the Maintenance Manual

for the M 601 E engine.

The checking of the electrical circuit of automatic feathering is carried out with the help of

the FEATHERINGIAUT. (M 327) checking push button on the testing panel on the left

control panel and it is described in the Flight Manual of the L 410 UVP-E aircraft.

Manual feathering

Manual feathering can be carried out in any flight condition during the failure of the engine if a

defect occurred in the automatic feathering circuit, or if the engines control levers are set outside

of the sector, which is determined for the automatic feathering function.

It is carried out by pressing the MANUAL FEATHERING push button of the respective engine on

the front control panel.

The electrical circuit is connected by two FEATHERINGIAUT. BANK circuit breakers on the

overhead panel. - - - - - - - - - - - - - - - - - - - - - -

- - - - - - - - - - - - - - - - - -

Bybepres&g the MANUAL FEATHERING (M 333, M 334) push button the TKE 54 PODG

(M 319, M 320) relay will switch on and at the same time it will give a current pulse for the

LUN 2601-8 (M 337, M 338) time relay, which will be put into operation for the time of 13 - 15 sec.

The time relay will carry out the switching on of TKD 501 DOD (M 343) contactor and this will put

into operation the LUN 7840-8 (M 345, M 346) auxiliary pump, which is supplied by power from

the AZRGK 40 (M 341, M 342) circuit breaker.

By switching on the TKE 54 PODG (M 319, M 320) relay current will be brought also to the

electromagnet of the LUN 7880.01-8 solenoid valve actuator. By its switching on the passage will

be opened to the small piston of the working valve, which will shift into the edge position

(feathering). By shifting the valve the outlet of oil from both output channels of the speed governor

will close, the direct connection of the pressure branch for adjusting the propeller will open to a

large angle (into feathering) and also the waste channel will apen-for We small angle. The- - - -

- - - - - - - - - - - - - - -

propeller will be thus set into the feathering position and is in this position blocked, as long as the

slide-valve remains in the position reached.

During the operation of the auxiliary pump (about 13 - 15 sec.) the 0.7s LUN 1492.03 (M 347,

M 348) pressure switch is switched on, which will cause the lighting up of the FEATHER. PUMP

yellow signal cell on the CWD ENGINE LHIRH.

By closing the relay (M 319, M 320), voltage and ground potential is supplied to the relay (M 317,

M 318). The relay (M 317. M 318) will close and voltage will be supplied through is contacts to the

isolation valve (M 27, M 28). The corresponding yellow ISOL. VALVE.

Feathering check

Checking of feathering is carried out by the simultaneous depressing of FEATHERING MANUAL

(M 323, M 325, M 324, M 326) push buttons for the left(right1 engine,-whichare-pla~ed on-the -

- - - - - - - - - - - - - - - - - - - -

- - -

testing panel on the left control panel.

These push buttons put into operation the TKD 501 DOD (M 343, M 344) contactor and the

LUN 7840-8 (M 345, M 343) auxiliary pump except the circuit of the time relay and the auxiliary

pump is in operation only during the time when these push buttons are depressed. During this

checking of the feathering it is necessary to follow the procedures, which are mentioned in the

Maintenance Manual for the M 601 E engine.

Emeraencv featherinq


Emergency feathering is carried out in case of failure of the automatic and manual feathering, by

shifting the propeller control lever of the defective engine beyond the stop, into the extreme rear

position. By shifting the PROP control lever into this position an oil channel enlarging the pitch

angle will be opened in the speed govern0 of the propeller.

In case the engine works at least in the idle power setting and its oil system is under pressure, the

blades are set by this pressure int o the feathering position.

In case the engine does not work, or there is no pressure in the oil system, the blades of the

turning propeller during the flight are adjusted into feathering position by moments of the

centrifugal force of small weights attached to the propeller blades.The time of feathering may in

some cases reach up to 25 sec (at zero pressure of oil).

NOTE: This way of feathering is used also on the ground after parking longer than two hours for

the venting of the propeller oil system and for the flooding of this system with warm oil.

14.3. INDICATION

The control instruments of V 510 propeller serve the screw for information about the immediate

function of both propellers. They signal the propeller, operation heating, P control and feathering.

The immediate revolutions of the left and right propellers are signalled through LUN 1348.01

propeller speed indicators which are on the instrument panel.

The propeller revolutions are scanned by combined LUN 1333.12-8 transmitters which are placed

on the reducer of the left and right engine.

The switching on and funct~on of heating of the propeller blades are signalled by signal cells on

the CWD.

The switching on and function of the BP - control(( is signalled by signal cells on the CWD.

The switching on and operation of feathering of the propellers are signalled by signal cells on the

CWD.


nv CENTRAL WARNING DISPLAY PROP. FEATHERING

ENGINE LH ENGINE RH AUT. BANK CONTROL

PROP. FEATHERING MANUAL AUTO MANUAL

@ @ @ @ @ I ---- i i r,---

I I MINIMUM I

MAXIMUM - 3 - -- ' I FLIGHT 1 l SETTING REVERSE SERVOMECHANISM

FIG. 14-1 PROPELLER CONTROL SYSTEM




AIRCRAFT TRAINING kl~NUAL L 410 UVP-E, E9, €20

CHAPTER 15

WATER INJECTION

15.1. General

15.2. Water distribution

15.3. Dumping and purging

15.4. Indicating

c& AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9. €20


AIRCRAFT TRAINING hlANUAL L 410 UVP-E, E9, E20

15.1. GENERAL The water injection system in the engine compresscr is intended for maintaining the maximum

power output of the engines during starting or at hicrer standardless temperatures of the

environment.Demineralized or distilled water is storsd in a tank built into the right-hand

undercarriage nacelle. The tank is made of light allcys. Its top part is provided with a stopper and

an air-bleeding valve, its bottom part with a filler. A lsvel gauge is fitted on the tank stopper. The

tank volume is 1 1 litres. its useful capacity is 10.5 litres.

15.2. WATER DISTRIBUTION - - - - - - - - - - - -

Description

The water distribution svstem comprises aluminum-dloy pipes the inner diameters of which are 6 .

8, 10 and 12 mm. The ?;ping system is attached to me airframe by means of sleeves. 3 e water

from the tank is delivered by the LUN 5155-8 water injection pump via the LUN 7377-8 three-way

cock 8 to the injection nozzle rings of both engines.Tne water injection pump is adjustable into

three working positions. each with a different water slipply amount, by means of a hamoperated

by-pass valve accessible after removing the oval lid under the leading edge of the right-hand

undercarriage nacelle.

Perm~ssible area of the use of tnjected coolant to maintain the englne power output

5 10 15 20 25 30 35 LO 45

Temperature o f the enwonment. OC

FIG. 15-1 WATER INJECTION PUMP WORKING STAGE SETTING VS.

ATMOSPHERIC PRESSURE AND AMBIENT AIR TEMPERATURE CHART

I. il. Ill. working stages of the water injection pump

AIRCRAFT TRAINING MANUAL L 410 UVP-E. E9. E2O

FIG 15-2 SCHEMATIC DRAWING OF THE 'NATER INJECTION DISTRIBUTION 5'. $ - - ?,I

AIRCRAFT TRAINING hlANUAL L 41 0 UVP-E, E9, E20

1

PRESSURE .WITCH

\ 1


The draining branch, leading from the water injection pump to the draining solenoid valve.

incorporates a nozzle which prevents larger amounts of water from returning to the tank when the

injection pump is running, and allows the discharge branch upstream of the water injection pump

and remnants of water in the tank to be drained.

NOTE: The lever of the LUN 7377-8 cock 8 must always be in such a position so that the WATER

INJECTION inscription is visible. If moved to the COMPRESSOR WASH position, the

system is ready for connecting a ground compressor-washing source.

The water ~nlection pump involves two connections or branches.

- to the 0 03 K LUN 1492.01-8 pressure switch

- to the LUN 2474 4-8 solenoid valve.

Operation

The stanby operation of the water injection system is actuated by turning on the WATER

INJECTION circuit-breaker on the over-head panel. Upon pressing down the WATER INJECTION

- ON pushbutton on the front control panel, the LUN 5155-8 water mjection pump starts runnmg.

NOTE: The LUN 5155-8 water injection pump starts running only when at least one of the engine

control levers is in the maximum take-off power ou t~u t position.

As soon as there IS pressure in the p~ping, the WATER INJECTION slgnal cell on the central

warning display lights up. The pushbutton must be kept pressed down until the signal cell lights

up. The maximum time during which the water injection system can be used is 1 minute. As soon

as the pressure drops (i.e. after pumplng out the water) in the piping, the pressure switch

a~scmfiected the circuits of the water injection pump and the WATER INJECTION signal cell. The

water injection system can also be turned off by pressing down the WATER INJECTION - OFF

pushbutton change-over switch on the front control panel.

The water injection system must be turned off immediately after the airplane has taken off, and

surplus water drained from the tank. The water injectior~ system can be used at ambient air

temperature equal to or higher than +23OC.

NOTE: When switching from the maximum take-off power output mode to a reduced power

output mode, the water ~njection system is turned off automatically.

15.3. DUMPING AND PURGING

The water Injection system is drained by the LUN 2474.4-8 solenoid valve. If the water injection

system is not employed during the take-off phase, the water must be drained in the course of flight

to prevent it from freezing inside the system. The solenoid valve is built into a bracket located

inside the right-hand undercarriage nacelle, under the water tank.The LUN 2474.4-8 solenoid

valve is turned on by the WATER DRAIN switch. The WATER INJECTION circuit-breaker must

be turned on as we1 I.


F!G 15-3 SCYEMATIC DRAiNING OF THE DUMPING AND PURGING


15.4. INDICATING The water injection svs:am signalization informs t k t user pilot that the water injection system is

working.lf there is a pressure in the water injection system piping, the 0.03 K LUN 1492.01-8

pressure switch turns 2n the WATER INJECTION signal cell on the central warning disp1ay.A~

soon as the pressure drops to zero (after water exhaust, the pressure signalization device

disconnects the water injection pump and signal ce!l circuit.

C E ~ R A L WARNING D I S P ~ Y AIRFRAME

I "! @ , I

u

" G 15-4 SCHEM?T!C ERAWING OF THE INDICATING

CHAPTER 16

SERVICING

16.1. Lifting and shoring

16.2. Levelling

16.3. Weighing

16.4. Minimum turning radius when

aircraft taxiing on the ground

16.5. Towing

16.6. Parking and mooring

16.7. Exterior marking

16.8. Interior placards and markings

16.9. Airfield servicing

16.10. Servicing in emergency situations

16.1 1. Ground equipment and tools

16.12. Airplane maintenance



- ~ - - ~ - ~ ~ ~


16.1. LIFTING AND SHORING 3 hydraulic lifting jacks and a tail prop strut XL 410.9620 are supplied together with the aircraft for

its lifting due to aircraft maintenance or levelling. On-board hydraulic lifting jack is supplied for

emergency cases during aircraft operation-tyre blowing, etc. For lifting the aircraft with one

hydraulic lifting jack in case of defect of the nose wheel, one short tail prop B 596 340 N is

supplied.

16. I. 1. Aircraft lifting - - - - - - - - - - - -

CAUTION: AIRCRAFT LIFTINGMAYBY CARRIED OUT ONLY m THE HANGAR, A - -

ROOFED HALL, ETC. IN THE OPEN SPACE, THE LIFTING MAY BE CARRIED

OUT ONLY WHEN NO WIND IS BLOWING.

(a) Remove covers enabling the acces to jacking points on landing gear nacelle - - see fig. 16-1

NOTE: Jacking points for hydraulic lifting jacks are marked with inscription PROP

LOCATION.

(b) Position the hydraulic lifting jacks under the respective aircraft jack point

(c) Check on, whether the overflow valve of lifting jacks is closed

(d) Open the vent valve - - - - - - - - - - - - - - - - - - - - - - - - - - -

(e) Shift the control bar on the pump and start pumping. If possible, do use thewhole stroke-of

the pump for pumping.

NOTE: Hydraulic lifting jack is not equipped with safety relief valve, therefore the maximum

lifting jack extension is limited by fixed stops. If you feel increased resistance during

the pumping, it means the jack has reached the maximum of hydraulic stroke. Stop

pumping. When using the lifting jack at temperatures of -45OC up to -50°C,

approximately 60 strokes per minute must be performed, so as to assure the lifting

jack full efficiency and preventing its aeration.

(f) After lifting the aircraft to the necessary height secure the lifting jack by the security strut.

WARNING: WITHOUT USING THE SECURITY STRUT, NO PERSON SHALL STAY - - - - IJNDERWEAIRCRAF-T.- - - - - - - - - - - - - - - - - -

(g) Position the tail prop strut into the hole in the lower fin of fuselage rear part near the frame

No. 24 and adjust the lenght of the prop L 410.9620 so that its base will be touching the

ground.

CAUTION: -WITH THIS POSITION OF THE TAIL PROP STRUT, THE HYDRAULIC

LIFTING JACKS MUST NOT BE MANIPULATED TO AVOID FUSELAGE

REAR PART DEFORMATION.

- IF THE AIRCRAFT RESTS ON THE LIFTING JACKS, THERE MUST BE NO


ASYMMETRIC LOAD ON ONE WING. THIS MIGHT CAUSE AIRCRAFT

TILTING OVER TO THE SIDE OF HAVIER LOAD. MAX. PERMISSIBLE

DIFFERENCE OF THE LOAD AT THE WING IS 50 KP.

FIG. 16-1 SCHEME OF LIFTING JACKS AND TAIL PROP STRUT LOCATION

(1) Left-hand hydraulic lifting jack HZ 4-3 or B 097 700 L, execution 3, (2) Right-

hand hydraulic lifting jack HZ 4-3 or B 097 700 P, execution 3, (3) Tail prop strut

L 410.9620, (4) Pin securing the prop in the fuselage, ( 5 ) Locking pin for the

prop legth adjustment, (6) Shortened tail prop strut B 596 340 N (not shown in

the fig. 16-2)

A - fulchm-in the landing gear nacelle


16.1.2. Aircraft lowering

CAUTION: BEFORE YOU START TO SINK THE HYDRAULIC LIFTING JACKS, MAKE SURE THAT THE TAIL PROP STRUT L 410.9620 HAS BEEN REMOVED. IF NOT, THE DEFORMATION OF FUSELAGE REAR PART MAY HAPPEN.

(a) Loosen the tightening nuts of the locking struts and loosen the locking struts of all 3 lifting hydraulic jacks.

(b) Loosen the overflow valve of all 3 hydraulic lifting jacks simultaneously to enable the plane to sin^ evenly.

16.1.3. Measures to be taken when lifting one wheel only

Position the lifting hydraulic jack under the jack point at the proximity either of the main or - - - - -

nose wheel. o n soft soil use a pad under thghydraulic lifting-jack base. - - - - - -

I When lifting the main landing gear

(a) Turn the nose wheel so that the lateral side of the nose wheel will remain turned towards the main wheel, which remains on the ground. Do not use chocks to secure the nose whetl. You will enable easier balancing of the aircraft position in lifting it in one point only

(b) Secure main wheel, which remains on the ground by means of chocks from both sides.

(c) When putting the lifting jack under the aircraft, position the jack such a way as to be slightly deviated ( l o approx) from the aircraft center line. The deviation will enable the lifting jack balancing when is being lifted.

When lifting the nose landing gear

Before lifting the aircraft front part it is necessary to install the shortened tail prop strut

B 596 340 N. Secure at least one main wheel by chocks from both sides. The chock in the

-front-must be approx. 5-cm from lhetyre, since main wheels willsh~ft slightlyfoyard - during -

the lifting.

16.1.4. Aircraft lifting by means of on - board hydraulic lifting jack

(a) Check on wheather the overflow valve is closed

(b) Open the vent valve

(c) Sh~i; on the control bar and start pumping. Use the whole pump stroke for pumping, if

possible.

NOTE: The jack is not equipped with a relief valve, therefore the jack maximum shift-out is

determined by dead stops. If you feel increased resistance during pumping, it

means the jack has reached the maximum of the hydraulic stroke. Stop pumping.

When - using - - the - lifting - - - - jack at temperatures, ranging from -45OC to -50°C, approx. - - - - - - - - - - - - - - - - -

60 strokes per minute must be performed in order guarantee the lifting jack full

efficiency and prevent its aeration.

(d) After lifting the aircraft at the necessary height, secure the lifting jack by security strut.

'NARNING' WITHOUT USING THE SECURITY STRUT NO PERSON IS ALLOWED TO

STAY UNDER THE AIRCRAFT


--

16.1.5. Aircraft lowering by means of on - board hydraulic jack

(a) Remove the security strut

(b) Slowly open the overflow valve to allow the aircraft to sink slowly.

16.1.6. Application of the rear shore

Tail prop strut XL 410.9620 and shoesned tail prop strut B 596 340 N serve to secure the

aircraft against tipping over on the fuselage rear part, when the aircraft rssts lifted on the lifting

jacks,when are carried on in the fuselage rear section (e.g. in the cargo compartment) and

when engines are dismantled in spite af the aircraft not being lifted on the liftiig jacks. It is - - - - - - - - -

- - - - - - - - - - - - - - - - - -

fixed i n a hole-ii the lower fin, near the 24th frame. The tail prop height must be adjusted as

follows:

-when the aircraft is on lifting jacks with the base touching the ground.

- when the aircraft is not on lifting jack - with the prop base elevated 10 cm from the ground, as

far as the terrain inclination permits for it (10 cm reserve allows for the possible

undercarriage dampers spring action).

Installation of on - board hydraulic jack HZ 4-3 in the aircraft

On - board hydraulic lifting jack and securing prop are placed between the 12th and 13th

frames on the passenger cabin right side under the seat no. 8. Hydraulic jack including the pad

for soft terain is fixed with two bolts provided with wing nuts into the thread holes in the - - - - - - - - - -

- - - - - - - - - - - - - passenger cabin floor, - - - - - - - - - - - - - - - - - -

NOTE: For fixing of on-board lifting, jack use always only bolts L 410.9515-01 with washers

8.2 CSN 021740.14. Security strut including the handle of lifting jack pump is stored in

artificial leather packing and is fixed with two straps to the passenger cabin floor.

16.2. LEVELLING The aircraft levelling must be done in following cases:

After disassembly and assembly of a wing, a fin rudder and stabilizer, in case the maximum

speed has been exceeded, or the operational load factor has been exceeded (acceleration),

after extraordinary hard landing or more serious aircraft damage, in case of flight performance

deterioration that has been caused by airframe deformation.

The positioning of the aircraft into horizontal position is done as follows: - - - - - -

- - - - -

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

- in transversal plane by means of levelling points no. 19 on the left and iight wing. Both levelling

points must be found in one plane (left and right point).

- in longitudinal plane by means of levelling points Nom. 1, 7.

The horizontal positioning of the alrplane itself is made by means of hydraulic lifting jacks in

accordance with item 16.1. A list of levelling in para 16.1. points and other dimensions for the

levelling are given in the levelling certificate, issued by the manufacturer for each airplane


separately. When checking on the levelling in operation, the previously measured levelling

values have to be relied upon; these are mentioned in the levelling certificate.

FIG. 16-2 HORIZONTAL POSITIONING OF THE AIRCRAFT


16.3. WEIGHING Aircraft weighing and counterbalancing is made by means of three balanceson which the

airplane is positioned by hydraulic lifting jack - see fig. 16-3.Balances inaccuracy for aircraft

weighing may be 0,2% max.

FIG. 16-3 AIRCRAFT POSITIONED ON BALANCES

(1) Front balance, (2) Left-hand balance, (3) Right-hand balance


Weiahina prcctdure

- Before starting the aircraft weighing, the aircraft must be in horizontal position according to

section 16.2.

- Find out whether the airplane outfit corresponds to the reguired configuration supposed to be

weighed.

- Record the values ascertained on the particular balances (Rl. R2)

- Put the ascenained values into the fo rm~ la and execute the calculation

The centerof gravityhom the referencepiane

Center of grawty in respect to the mean aerodynamic chord (MAC)

XL - A X -2,191 XT =- .I00 = L . lo0 =. . . . . . . SAT

b 1.91 8

center of gravity position from reference plane (levelling point no.2)

distance of the front prop axis from levelling point no. 2 (X1=0.865 m) - - - - - - - - - - - - - - - - - - - - - - - - - - - -

distance of the left and right-hand axis from levelling point no. 2 (X2=3,700 m)

reaction on front balance (after subtracting the hydraulic lifting jack weight)

reaction on rear balances (sum of reaction on the left and right-hand balance

after subtracting the hydraulic jack weight)

aircraft total weight

distance from levelling point no. 2 to the beginning of mean aerodynamic

chord (A=2.192 m)

length of mean aerodynamic chord (b=1.918 m)

- - - - - - - - - - - - - - - - - - - - - - - - - - - -


FIG. 16-4 SCHEME FOR COUNTERBALANCING CALCULATION

(1) - Center of gravity, (2) levelling point no. 2.

For other symbols see above text.

MINIMUM TURNING RADIUS WHEN AIRCRAFT TAXIING ON THE GROUND Fig.16-5 shows the necessary space shown by circles for the turning of the airplane.The

smaller circle shows the circle written by the outside wheel of main landing gear, the bigger

circle shows the circle, written by the tip of the wing. The values are given by the maximum

deviation of the nose wheel, i.e. 50° - 5O Ithe marginal tolerance of -So. 1.2 4 5 O is considered

FIG. MINIMUM TURNING RADIUS FOR AIRCRAFT TAXIING

The values are given in metres and rounded-off to tenths

OF THE GROUND

of metre.

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E. E9. E2O pppp-p--ppp--- ~

16.5. TOWING

16.5.1. TRACTOR TOWING Toswing h e aircraft by a tractor is carned out by means of towing equipment L 410.9525 or

B 097 581 N (hard terrain) or by XL410.9521 or B 096 150 N (soft terrain).

(a) Towin eaui~ment L 410.9525 or 6 337 581 N

Towing equipment L 410.9525 or 8 097 581 N is secured against shocks by a spring

damper; located in the front part of the towing shaft-bar. The shaft bar itself (tube) of the

towing equipment is connected with the suspension device by a connecting element with

2 security pins. The material of security pins is selected such that the front pin will be

damaged in case of excess load in the straight direction and the rear pin will be damaged

if excess force is developed in the lateral direction.

A label is attached on the bar referring to the pins material to be unconditionally used if

the pins must be replaced:

I SECURITY PINS

MATERIAL: STEEL

TENSILE STRENGHT SpT=392 MPa (40 kp/mm2)

Attachement of towing equipment is carried out as follows:

Position the shaft-bar of the towing equipment against the hole in the swinging arm of the

nose landing gear and insert the pin (2) - see fig.16-6 into hole. Secure the pin (2) by a

security plunger (3).

FIG. 16-6 TOWING EQUIPMENT L 410.9525 or B 097 581 N

(1) Towinq eaui~ment. (2) Pin. (3) Security plunger. (4) Front security pin.

(5) Rear security pin. (6) Label, (7) Nose landing gear leg


(b) Towina eauioment XL 410.9521 or 6 096 150 N (see fig. 16-7) (if installed)

The towing equipment is equipped with a security pin (6) - se fig. 16-7 against landing

gear damage in case the maximum deviation angle of the nose wheel is exceeded.

CAUTION: IF THIS PIN HAS BEEN DAMAGED. THE PIN MAY BE REFLACED ONLY BY

THE PIN DRW. NO. L 410.9522-09 OR BY A PIN ,MADE OF MATERIAL

WITH TENSILE STRENGHT MAX. 392 MPa (40bplmm2).

Attach towing equipment as follows:Position the shaft-bar of the towing equipment (2) against

the hole in the swinging arm of the nose landing gear, insert the pin (4) and secure it by

putting down the latch.Position the yokes (7) which are at the end of the towing cable on the

main landing gear by means of pins (8) inserted into holes in the hollow pins fastening the

landing gear damper and secure them with a wedge (9).

FIG. 16-7 TOWING EQUIPMENT 8 096 150 N

(1) Towing equipment, (2) Shaft-bar, (3) Cable, (4) Pin for attachement to the

nose landing gear, (5) Nose wheel, (6) Security pin, (7) Yoke, (8) Pin for the

attachement to the main wheel, (9) Wedge, (10) Main wheel, (1 1) Coupling for

the left-hand side, (12) Locking L 410.9524-20 Ifor the right-hand sidel,

(13) Cable, (14) Security pin L 410.9522-10

(P) - Detail of the attachement to main wheel

(R) - Detail of the attachement to the nose wheel


(c) Instructions for towina with a tractor

When towing the aircraft by a tractcr, the following regulations must be observed:

- ln the cockpit cabin a pilot or mechanician must be seated. familiar with the attendance

of the airplane brakes.

- The blocking strut of the steering L 410.921 0 must be removed.

- The servo of the nose wheel as well as the nose wheel from the pedals must be shut-off,

i.e. FEDAL NOSE WHEEL STEZING on the control console must be in neutral

position.

- The brakes accumulator must have a pressure of at least 9.80 MPa (100 kp/cm2).

NOTE: The pressure gauge of brakes accumulator indicates only with the 36V convertor

on. After checking on the pressure, switch-off the convertor immediately. If the

pressure in the brakes accumulator is not sufficient, the aircraft may be towed

exceptionally, but the pilot (mechanicial) must bear in mind that in emergency

case when the situation will require braking the aircraft he will apply the parking

brake.

- The maximum turning angle of the aircraft nose wheel as the tractor starts its way with

the aircraft attached must not excced 25O.

NOTE: During manual towing it is allowed maximum turning angle of the nose wheel of

30° when starting pull.

- Before you instruct the tractor driver to start towing check on whether the parking brake

is off and the aircraft door is closed.

CAUTION: THE AIRCRAFT MUST NOT BE MOVED OR TOWED \NITH ENTRANCE

DOOR OPEN.

- During the towing, the nose wheel may be swivelled to the maximum of 300.

- When towing the aircraft, the tractor driver must avoid rash movements with the steering

wheel and sudden braking. It is recommended not to exceed the speed of

10-1 5 km/hour.

- When towing the aircraft at limited visibility, switch on the aircraft position lights.

- Rearward towing (i.e. tail ahead) of the aircraft by a tractor is not permitted.

16.5.2. Aircraft handing by means of manual tow bar

To manipulate the aircraft as it is towed out of the hangar, etc., a manual tow bar

L 410.9521 is used, which is substantially lighter than the towing equipment, used for towing

the aircraft by a tractor. The bar is attached by means of a pin in the swmging arm of the nose

landing gear leg - see fig. 16-8.


FIG. 16-8 MANUAL BAR L 410.9521

(1) Manual bar, (2) Pin, (3) Nose landing gear leg

16.6. PARKING AND MOORING 16.6.1. Airplane parking

To secure the airplane for parking at an apron the following equipment is fixed- a parking

equipment for parking during a flight day

- a parking equipment

- all coating wrappings

- earthing equipment

16.6.2. Parking equipment for parking during a flight day

To secure the airplane for parking during a flight day, a blocking strut B 596 476 N is used.

16.6.3. Parking equipment \

To secure the airplane parking on the apron. the following parking equipment is used:

I No. Name i Type (DW~.NO.) Qty I Propeller blocking B 596 276 N

or B 596 695 N

2 Wheel chock L 410.9250 6


I No. Name Type (Dwg.No.) Qty I Aileron security clamp

Elevator blocking

Rudder blocking

Cover of engine air inlet

Exhaust cover (for winter season)

Exhaust coating .. (for summer season)

Pitot nozzle cover

Blind flanges for static pressure pickups

Coating ior windshield flight compartment

Toilete ventilation cover

Cower of starting generator air inlet

L 410.9220

L 410.9220

B 596 790 N

B 596 452 N

B 096 127 PIL

B 096 008 N

B 096 360 N

L 410M.9537

L 410.9516

B 596 420 N

B 596 580 N

FIG. 16-9 PARKING EQUIPMENT FOR PARKING ON THE APRON

Legend to fig.: see table in para 16.6.3.

Detail 6 - the B 596 695 N (1) propeller blocking is supplied with the aircrafts

from the 25th series.

AIRCRAFT TRAINING MANUALL 410UVP-E, E9, E20 - - .

16.6.4. Coating assembly (if installed)

Coating assemblies serve for airplane protection on the apron. The coatings are marked with

serial numbers 1 - 24 and are inserted into coating wrappings L 410.9320-19, marked with

serial numbers I - VI.

The airplane coatings assemblies include:

I 1 L. H. wing coating B 596 397 L W. D. 1

2 1 L.H. fuel tank coating B 596 338 L W. D. 1

Coating wrapping

23 Coating of wing tip B 596 451 L W. D. 1

Name Coating number

2 L.H. wing-fuselage coating B 096 076 L W. D. 1

3 Engine coating B 096 077 L W. D. 1

I I 4 R.H. wing coating B 596 397 P W. D. 1

Drawing number

22 R.H. fuel tank coating B 596 338 P W. D. 1

Pcs

24 Coating of wing tip B 596 451 P W. D. 1

5 R.H. wing-fuselage coating B 096 076 P W. D. 1

6 Eng~ne coating B 096 077 P W. D. 1

Ill 7 Propeller blade B 596 398 N W. D. coating w L,

8 Propeller cap coating B 596 400 N b.v 2

9 Propeller blade coating with strap B 596 399 N W. D. 2

10 Flight compartment coating B 596 405 N W. D. I

IV 11 Fuselage rear part coating B 096 078 N W. D. 1

12 Fuselage central part coating B 096 075 N W. D. 1

13 Antenna coating L 410M.9322-37W. D. 2

V 14 L.H. stabilizer coating B 196 080 L W. D. I

15 R.H. stabilizer coating B 096 080 P W. D. 1

16 Fin coating B 096 081 N W. D. 1

V1 17 L.H. landing gear nacelle coating B 596 396 L W. 0. 1

AIRCWFT TRAINING MAFllJdL L d'lO UVP-E, E9, E20

I Coating i Coating wrapping number

I Name I Drawing number I Pcs

VI 18 R.H. landing gear nacelle mating B 596 396 P W. 0. 1

19 Nose landing gear coating B 096 07; N W. D. 1

20 Wheel cca:ing 8096085LlPW.C 1+1

16.6.5. Earthing equipment

Earthing equipment B 096 432 N is designet for earthing the airplane.

FIG. 16-10 COVERS FOR AIRPLANECOVERING

Legend to fig. see table in item 16.6.4.


2 Mooring device for the wing B 596 278 N or B 596 670 N 2

3 .Additional rear mooring device B 596 356 N or B 596 673 N 2

4 Additional front mooring device B 596 358 N or B 596 672 N 1

5 Mooring device of tail unit B 596 372 N or B 596 674 N I

6 Mooring peg LDN 691 1 12

16.6.6. Mooring

The airplane mooring is to be carried out when the wind with gusts exceeding 20 mls (72 kmlh) is

assumed. For airplane mooring the followins equipment is used:

NOTE The ground equipment for the aircraft starting from the 20th series includes

new mooring equipment. The differences in design of the mooring devices

delivered with the aircraft up to the 19th series and those starting from the

20th series have no effect on their function, and therefore both kinds of the

No.

mooring devices can be used on any aircraft regard less of its Serial Number.

FIG. 16-1 1 AIRPLANE MOORING

Details of the mooring device (1) through (5) are drawn in the fig. 16-12 and

16-1 3.

1 Moor~ng device of nose wheei B 596 277 N or B 596 671 N 1

QtY Name Type (Dwg. No)


FIG. 16-12 AIRPLANE MOORING - - - - - - - - - - - - - - - - - - - - - - -

- - - -

(made till the 19th series)

(1) Mooring device of nose wheel B 596 277 N, (2) Mooring device for the wing

B 596 278 N, (3) Additional rear mooring device B 596 356 N, (4) Additional

front mooring device B 596 358 N, (5) Mooring device of tail unit B 596 372 N,

(6) LDN 691 1 mooring peg. (7) Turnbuckle. (8) Eye. (9) Pin. ( lo) Pin, (11) Nut,

(12) Side plate, (1 3) Fork, (14) Fork, (1 5) Screw. (16) Securing pin, (1 7) Pin,

(18) Nut, (19) Eye, (20) Hook.


FIG. 16-1 3 MOORING DEVICE

(made from the 20th series)

(1) Mooring device d nose wheel B 596 671 N. (2) Mooring device for the wing

B 596 670 N, (3) Additional rear mooring device B 596 673 N. (4) Additional

front mooring device B 596 672 N, (5) Mooring device of tail unit B 596 674 N,

(6) Cable, (7) Turnbuckle with chain, (8) Pin, (9) Pin, (10) Bracket, (1 1) Fork.

(12) Screw, (13) Fork. (14) Side plate. (15) Nut. (16) Securing pin, (17) Pin.

(18) Nut. (19) LDN 691 1 mooring peg.


16.7. EXTERIOR MARKINGS

TURN TWICE THROUGH 360 DEG TO RELEASE LOCKS

All airplane exterior markings are showed bellow in the chart.

OIL 11 I CHECK OIL FILLER CAP CLOSED

POS. ( F ig No. / Inscription reading

WATER INJECTION COMPRESSOR WASH

Note

GROUND HYDR. SOURCE 14.7 MPa COMPRESSED AIR GROUND SOURCE 0.54 MPa HYDRAULIC TANK GAUGE PRESSURE

Red or greencross

CUT HERE TO BREAK IN

VHF AERIAL

SUPPORT HERE

1. PUSH

2. TURN

CLOSED

OPEN

NITROGEN 1.47 MPa

DO NOT STEP HERE

NITROGEN 4.9 MPa

INTERCOM

Concerns locks of covers for access to board batteries and el-radio equipment. Locks are releasable with a screwdriver with rounded edge

Ground electric installation socket

Access to engine oil filler

Access to three-way cock for ,engine water injection(( or ,compressor wash((.

Access to ground hydraulic source connection and ground pressure air connection for hydraulic tank pressurizing

First-aid box location

Area destined for forced airplane entry (after emergency landing to break in)

Marking of VHF antenna

Supporting jack point for the rear strut

Main door opening and shutting

Brakes battery filling

Area not allowed to be stepped on by feet

Hydraulic battery filling of main conduit

Main wheel tyre pressure

Nose wheel lyre pressure

Acces for connection of ground personnel headphones


FUEL DRAIN Access to fuel tanks drain valves

PRESS FIRE EXTINGUISHING BY Access for extinguishing with GROUND MEANS ground extinguishers

OPEN Emergency exit door opening and shutting

TURN

CLOSED

L

Red circle Reference levelling point no.2 (for calculation of the centre of gravity position - see section 8-20-00)

DESTILLED WATER MAX. 10 1 Access to tank water filler for engine water injection

.;'/IATION KEROSENE Fuel for engines

Pos.

FUEL 200 1 MAX. PRESS. 0.39 MPa Fuel for engines

FUEL 628 1 MAX. PRESS. 0.39 MPa Fuel for engines

HYDRAULIC FLUID Access to hydraulic fluid level checking neck

AIR 0.49 MPa Access to the connection for checking of leading edges pneumatic deicmg device

OIL DRAIN Access to engine oil drain valve

DO NOT PRESS HERE Covering of marked-off area is not allowed to be pressed in order not to cause its deformation

HYDRAULIC FLUID DRAINAGE AND Access to hydraulic tank drain FILLING valve

Airplane product.ion number (if installed)

1. PUSH AND HOLD (if installed) Emergency exits opening

Note Fig. No.

2. TURN (if installed) Emergency exits opening

Inscription reading

3. PULL (if installed) Emergency exits opening

Identification label (if installed) Identification label denotes imatriculation sign. type and production No. of the aircraft

ELT (if installed)

HYDRAULIC TANK GAUGE PRESSURE Location of hydraulic gauge


42 16-1 5 DANGER-PROPELLER (if used) Front, emergency

Pos.

43 16-14 Red edging (if used) 16-15

marking zone anti-ice beet

I Fig. No.

Fig. 16-14 MARKINGS ON THE AIRPLANE-LEFT-HAND SIDE

A - transition skin between the wing and fuselage

Inscription reading Note


FIG. 16-1 5 MARKINGS ON THE AIRPLANE-RIGHT-HAND SIDE

A - right-hand engine nacelle-view from the fuselage side

I3 - trans~tion skin between the wing and the fuselage


FIG. 16-16 MARKINGS ON THE AIRPLANE-TOP VIEW


FIG. 16-17 MARKINGS ON THE AIRPLANE-BOTTOM VIE'N


16.8. INTERIOR PLACARDS AND MARKINGS All intenor placards and markings are showed bellow in the chart.

Between 6 and 7 Serigraphy on upholstery panel

frames-at the door

handle of the 1, OPEN emergency exit CLOSED

Location

Between 7 and 8 frames-atthe door - - - -

handle of the '+

Reading

emergency exit OPEN

Note

All emergency exits REMOVE

handle cover

Between 7 and 8 EXlT

frames near the erne-2ency exit dccr

Serigraphy on upholstery pwe! - - - - - -

Red colour on ABC white

material

Organic glass printins

(serigraphy) lighting inscription

Between 7 and 8 DANGER - PROPELLER Made of red coulor on light

frames above yellow ground emergency exit door

7 th frame on the FASTEN SAFETY BELTS Organic glass printing - - - - - - -

c o n tr6lscover - - - - - - - (engravedblighting inscription

7th ':ame on the NO SMOKING controls cover

Organic glass printing

(serigraphy)

Between 10. and 11. PUSH TO CALL CREW Organic glass printing

and 13. and. 14. (serigraphy) red letters on

frames-upper right white material

16th frame-the mam FIRE Red colour (serigraphy) on

door EXTINGUI- the panel

SHER

Above the main door EXlT

- - - - - - - - - - - - - -

Between 13 and 14 EXIT (if installed)

Organic glass printing

(serigraphy) lifting inscription - - - - - - -

Lighted transparency

frames on LH and

RH side above

emergency exit door


Main door

I

Cover on the main door

Location

Main door framing flang side

First-aid box ccver

I Reading

Edge shelf of

baggage compartment

Side of baggage comparment

Note

Between 19th and 20th frames - on the ceiling in front of WC

21 th frame WC door

Passengers seat rear side

7 th frame on the controls cover and on LH. RH side before

the first seat row and on the front side of

rest of each passanger seat

OPEN ;

PRESS

CLOSED Organic glass printing (serigraphy), coated with luminous material

Organic glass printing (serigraphy), coated with luminous material

Identification label denotes Metal printing (serigraphy) imatriculation sing, type and production No. of the airplane (if installed)

t @ t Serigraphy on the box cover

UPPER SECTION LIMIT LOAD Duralumin sheet printing 60 kg LOWER SECTION LIMIT (serigraphy) LOAD 90 kg

WC OCCUPIED or LAVATORY Organic glass printing OCUPIED (if installed) (serigraphy) lighting

inscription

RETURN TO YOUR SEAT Acrylon printing (serigraphy) (if installed) lighting inscription

White organic glass printing

WC -0- with black colour (serigraphy)

1 2 3 (front seats) 4 5 6 7 8 9 10 Siphoflexe printing

11 1213141516 1718 19(rear (engraved) seats)

r , Label (serigraphy) I LIFE -VEST UNDER YOUR I SEAT

(if installed)

AIRCRAFT TRAINING MANUAL L 410 UVP-E t9, E20

On the ceiling

opposite the rear

baggage compartment

19th frame

I

- - - -

o n jthframe on controls cover

Location Reading

On the emergency exits

Note

On the entry door and all emergency exits

I

At-emergency exits- -

between frames No. 13 and 14 on both sides of fuselage

LOCATE p, OXYGEN 8 n

SEKERA El Label (serigraphy)

Organic glass printing (serigraphy)

Organic glass-printing (serigraphy)

EMERGENCY EXIT

LOCKED UNLOCKED

- - CLOSE- OPEN

1. REMOVE 2. PUSH AND HOLD

3. TURN 4. PUSH

(if installed)

Orga'nic glass printing (ser~graphy)

Organic glass printing (serigraphy), coated with luminous material

Organic glass printing - - - - - - - - - -

(serigraphy), coated with luminsus material


ldentification of ~ i ~ i n c l and hoses of individual assemblies

Piping system 1 Piping system

Hydraulic . light blue chrome light blue

system flow yellow rings

direction

I Identification with self-sticking tape (normal

temperature range)

yellow rings M I

Identification with paint (Raised temperatures) -

I ldentification with self-sticking tape (normal

Fuel system

I temperature range)

Oil system I - I

Fire extinguishing svstem

irsfi . light blue chrome light blue

I

I ldentification with paint

Total and static pressure system:

I 1 (Raised temperatures)

Pressure probe piping R

1 I

Static pressure piping

I

Total pressure

piping

tm2d Air conditioning (heating)

flow direction(b1ack -El medium chrome arrow in white field) yellow

cherrish redlrays E l cherrish red

four stars flow direction flow direction

medium chrome yellow medium chrome

yellow alternating squares

coffee brown coffee brown

rhombuses

light

red light grey

zig-zag line

light red light grey zig-zag line

light

red light grey

zig-zag line

coffee brown light grey W


16.9. AIRFIELD SERVICING 16.9.2. Review of materials for operation

Fuel T-1 (ST SEL 5024-85, or GOST 10227-86)

TS-1 (ST SEV 5024-85, or GOST 10227-86 or CSN 656520)

RT (ST SEV 5024-85, or GOST 10227-86 or 653520)

PL-6 (PND 25 005-76)

PL-7 (PND 25005-76)

JET A (ASTbID 1655-89) - - - - - - -

- - - - - - - JETA-l (ASTMD1655-89prDW22494) - - - - - -

PSM 2 (PN-WC-96026)

The mlxlng and filling of the fuelsis permitted.

€3-3V (TU 38-101295-72) or TU 38-101295-85). Aeroshell

Turbine Oil 590, 555, 560 MIL-L236399C.

The mutualmixing of the oils is permited according to

MI L-L-23699C specification.

Oil

Hydraulic liquid and the liquid AMG-10 (GOST 6794)

for the shock absorber of

main nose landing gear

Liquid for engine injection Deionizated water (PND 31-1 151 -65)Destilled water - - - - - - - - - - - GOST 6709-72)- - - - - - - - - - - - -

Deodorant liquid for the toillete

Greases:

- bearings of landing gear NH2 (PND 25-024-69) or.

wheels NK-50 (up to -1 5OC)

mixture 75% NK-50 and 25%

ClATlM 201 (-1 5OC and less)

- all lubricated joints ClATlM 201 (GOST 6267-59)

of the airplane ClATlM 221

MOLYKA (PND 33-053-62) or - - - - - - - - - - -

- VNICNP242- - - - - - -

OKB 122-16

Oils for lubrication of

airframe parts CKB 132-08 (GCST 1875-73)


16.9.3. Review of holes for servicing and checking

The following review comprises the holes, enabling the access to the maintenance and

checking points and for dismantling of certair: airplane parts. The numbers, intended for

carving out the routine maintenance have been squared in the review. For example: @ Individual hole covers are divided according to the way of opening as follows:

I - A cover with a compression closure (can be opened without using any tool).

ll - A tilting cover for opening of which a screwdriver is needed

111 - Screw fixed covers

IV - A cover fixed with the help of rotary closing locks, releasing of this locks is carried out by

:urning of the lock trunnion with the help sf a screwdriver with the rounded edge by 1.5 - 2

revolutions to the left. When assembled. tightening is made as that on the normal screws.

'.CITE: In case that durinr; :he tightening the trunnion falls out after releasing the

pressure of the screwdriver. it is necessary to turn the trunion backward to the

left till the noticable stop and then carry out new tightening by turning the

trunnion to the right.

Access to the fuelmeter and assembly of the fuel tank


Access to the engine control. control of the wing flaps and ailerons

Access to the mud collector for interconnection of fuel tanks and to the electric installation

Assembly of the fuel tank and air bleeding outlet

Access to the control of the ,,ving flaps, to the fuel system and electric installation


Assembly of the fuel tank

Access to the engine control. the control and the electric installation

Access to the neck for checking the level of the hydraulic liquid (small lid)

Assembly of hydraulic liquid tank (outer lid) Ill.

Access to the filling neck of the fuel tank (small lid)

Access to the fuelmeter (larger lid)

Access to the fuel tank filling neck (small lid)

Assembly of the fuel tank (larger lid)

w

111.

Ill.

Ill.

Ill.

111.

111.

Ill.

Ill.

111.

II.

II.

Ill.

II.

Ill.

7

'Over 1 Fig.No. VPe i

Ssr. No Intended for


I

Ser. i No. j

Intended for

Access to the position light Ill.

Checking and assembly of the wing flaps and ailerons ccntrol. electric installation, deicing of leading edges and the engine control cables I.

Access to the oil filling neck in the engine I.

The assembly of electric installation and fuel system behind the wing rear spar-(pons are in the wing shroud) - - I l l ;

Access to the electro-mechanical strut of the aileron trim tab. II.

Access to the searchlights

Access to the front baggage compartment

Access to the front baggage compartment

Access to the wing connection with the fuselage to the work cylinder of the wing flaps

Access to the wing connection with the fuselage - - - - - - - - - - - -

Access to the hydraulic system, aileron control, socket connection of the electric installation and to the connection for connecting the air pressure to the system of the pneumatic deicing

Access to the suspensions of the elevator

Socket of the outer source for the electric installation

Access to the cockpit battery

Access to the fuselage rear section and the wing tip position lamp

Inspection of the left engine, fuel cleaner

- - - - - - - - - - -

Pressure checking in the fire extingusher of the left-hand engine

Access to the instruments behind the front fire wall, discharge of the hydraulic liquid, to the connection of the external source of the hydraulic system, hydraulic cleaner, battery filling, replenishing of hydraulic liquid

Ill.

I.

I.

Ill.

Ill.

Ill.

Ill.

I.

IV.

Ill.

II.

I.

IV.


Access to the propeller controller of the left hand engine

Ser. No.

Access to the engine, to the suspensions of the left-hand engine bed

Access to the engine control, starter-generator of the left- hand engine

Access to the unions for blowing of engine fire extinguishers

Intended for

Access to the radio equ~pment and condensate sump of the p~tot-system. The cover is iwided into two halfs. For routme maintenance only the upper half is to be removed.

Access to the radio-electro equ~pment and the condensate sump of the pito-system

Access to the propeller controller of the right-hand engine

'Over I Fig.No. type I

Access to the engine and the suspension of the engine mount of the right-hand engine

Access to the engine control, starter-generator of the right-hand engine

Access to the instruments behind the front fire wall

Access to the unions for blowing of engine fire extingushers

Inspection of the right-hand engine, fuel cleaner

Pressure checking in the fire extinguisher of the right - hand engine

Access to the VHF antenna

Access to the tad plane suspensions, the fin, electric installation

Access to the elevator control

Access to the control details of the rudder trim tab

Access to the control of the elevator trim tab

,Access to the drums of the rope control of the elevator trlm tab

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, €20

60. Access to the landing gear hydraclic system

61. Access to the landing gear hydraulic system

Cover I Fig.No. tY Pe

Ser. No.

62. Access to the space of the landing gear left-hand nacelle 111. 16-19,

16-20, 16-22

Intended for

63. Door of the left-hand landing gear controlled with the landing gear I. 16-1 9,

16-20, 16-22

65. Access to the space of the landing gear left-hand nacelle 111. 16-19

66. Access to the left-hand supporting dish for airplane lifting 111. 16-19

67. Access to the space of the landing gear right-hand nacelle 111. 16-19, 16-21, 16-22

68. Door of the right-hand landing gear controlled with the landing gear 16-19,

16-21, 16-22

170 Access to the water injection in the engine system (tank. pump, etc.) 111. 16-19,

16-21, 16-22

71. Access to the landing gear hydraulic system 111. 16-19

72. Access to the space of the landing gear right-hand nacelle 111. 16-19

73. Access to the right-hand supporting dish for airplane lifting 111. 16-19

176 Access to the mud removing and fuel discharging valve I. 16-20

r/7 Access to the mud removing and fuel discharging valve I. 1 6-2 1

1-/8 Access to the fuel pumps 111. 16-20

179 Access to the fuel pumps 111. 16-21

80. Access to the devices behind the fire wall

81. Access to the devices behind the fire wall

AIRCRAFT TRAINING MANUAL L 411 0 UVP-E, EQ. E2O

Access to the mixing chamber, damper of the heating system I

Ser. No.

Access to the piping for heating and venting 111. 16-22

Access to the three-way cock for the compressor washing and the water injection in the engine I. 16-20,

16-2 1

i

Intended for

Access to the adjusting elements of the control pull rods and ropes 111. 16-22

Access to the connections of the hydraulic system 111. 16-22

'Over

type

Access to the filling neck of the water injection tank in the engine Ill. 16-22

Fig. No.

Access to the rope stretching of the trim tabs 111. 16-22

Assembly of hand operated hydraulic pump

Access to the adjusting elements of the control pull rods and ropes

Access to the control levers

Rope stretching of the trim tabs

- - - - - - - ~ ~ ~

Access to the antenna of the marker receiver 111. 16-22

Access to the connection of heating piping 111. 16-22

Access to the connection of heating piping 111. , 16-22

Access to the loop frame of the radio-compass 111. 16-22



Access to the fuselage rear section IV. 16-22

CAUTION: WHEN WORKING IN THE REAR SECTION OF THE FUSELAGE (BEHIND THE 21TH BULKHEAD), USE THE ASSEMBLY BOARD B 596 331 N TO COVER THE SPACE ABOVE THE ANTENNA OF THE RADIO ALTIMETER (BETWEEN THE 21TH AND 22ND BULKHEAD) AND COVER THE CONTROL ROPES OF THE ELEVATOR TRlM TAB NOT TO DAMAGE THE CABLE TO THE ANTENNA OF THE

- - - - RADIO-ALTIMETER AND ROPES OF ELEVATOR. TRIM TAB CONTROL. AFTER FINISHING WORK REMOVE THE ASSEMBLY BOARD.

Access to the seats suspensions Ill 16-22

Covers of landing gears, controlled by landing gear 16-19

Access to the lock of the nose landing gear 111. 16-19


Access to the landing gear hydraulic system 111. 16-22

Ser . No.

Access to the control pull rods and ropes for controlling the engines and for controlling the elevatcr trim tab. 111. 16-22

Access to the control locking with the help of the blocking strut I. 16-22

Intended for

Inspection of the wing bottom skin ;ii. 16-22

16-19 Aceess toiheoper&ioncyhnder of the ABC tab - I. - - - -

Cover

Access for assembly of the socket of the outer source and the static pressure sensor I! I 16-20

Fig.No.

Assembly of fire extinguishing system, operation cylinder and the control of the interceptor and hydraulic system 111. 16-18

1.Y Pe -

Access to the adjustment of the pump stage for water injection in the engine

Access to the discharge valve of water injection system in the engine

Access to the mud removal of piping of the stall speed probe I. 16-22

Access to the mud removing and discharge valve I. 16-19

Inspection-of the floor and €uselage skin

Access to the lock on the pull rod for closlng the front covers of the nose landing gear I. 16-1 9

Air pressure check in the hydraulic tank pressurizi~c; system I. 16-20

Access to the door guide Ill. 16-22

Access to the filling neck of wing tip fuel tank II. 16-18

.Access to the fuelmeter of wing tip fuel tank 111. 16-18

Fuel piping assembly 111. 16-18

Assembly of banking plane 111. 16-18

Assembly of control cylinder for banking plane and ~nstallatlun o f electro~magnetic - valves - on fuel

- - - - - - - -

piping to terminal fuel tank In. 16z18

NOTE: Beside the above covers, there are removable upholstered panels, in the passenger compartment:whch are fixed by the means of screws.

Acces to discharging washing the toilet (if installed) II. 16-21

Access to neck of single-point fuelling (if installed) 111. 16-21


FIG. 16-18 REVIEW OF HOLES FOR SERVICING (TOP VIEW)


FIG. 16-19 REVIEW OF HOLES FOR SERVICING (BOTTCM VE\N)

I - Right-hand landing gear nacelle (from the bottom)


FIG. 16-20 REVIEW OF HOLES FOR SERVICING (LEFT-HAND SIDE)

I - Left side of left-hand engine nacelle

ll - Right side of left-hand engine nacelle


FIG. 16-21 REVIEW OF HOLES FOR SERVICING (RIGHT-HAND SIDE)

I - Left side of left-hand engine nacelle

ll - Right side of left-hand engine nacelle


FIG. 16-22 REVIEW OF HOLES FOR SERVICING (PASSENGER COMPARTMENT

FLOOR AND LANDING GEAR NACELLES)


16.9.4. Scheme of areas, prohibited to be stepped on. Scheme of places which cannot be pressed

FIG. 16-23 AREA PROHIBITED TO BE STEPPED ON

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, €9, E20

00 NOT PRESS HERE

FIG. 16-24 PLACES WHICH CANNOT BE PRESSED (BOTTOM VIEW)


16.9.5. Scheme of the access to the tanks

Diagram of tanks .ocation w~th inflammable matters is shown on fig. 16-29.

16.9.6. Dangerous area around the airplane, when the engines are running.

FIG. 16-25 SCHEME OF DANGEROUS AREA AROUND THE AIRPLANE. \NHEN

ENGINES ARE RUNNING (IDLING RUN)


FIG. 16-26 SCHEME OF DANGEROUS AREA AROUND THE AIRPLANE. WHEN

ENGINES ARE RUNNING ON MAX. TAKE-OF REGIME


16.9.7. Review of airfield servicing means of general purpose application.

! Characteristics connection dimensions of Specification 1 nozzles and reducers 1

1. Road tanker

--

Fuelling gun, ND 38 mm

2. Ground source of power supply Characteristics as per GOST 19705-74, outlet voltage 28/29 V, permissible load 700 A, ShRAP 500

3. The car for spraying the airplane with antifreezing liquid Not specified

4. General purpose heating car, engine driver OD of heating hose outlet hole is 200 mrn

5. Mobile sxygen filling station Connecting union thread: 21.8 x 1/14 pipe thread

6. Aggrecate VZA for filling the system of distilled water injection Not specified

16.9.8. Scheme of the location of the airfield servicing means

FIG. 16-27 SCHEME OF THE LOCATION OF THE AIRFIELD SERVICING MEANS

(1) Road tanker, (2) Ground source of power supply. (3) Car for oil filling,

(4) Hydraulic truck

AIRCRAFT TWINING MANUAL L 410 UVP-E, E9, E20

16.9.9. Instruction for refueling

(a) The operating persons must be properly trained about the work safety and fuel handling,

must not wear the dress made of artificial fibres (silon), must wear rubber gloves

(b) No open flame is allowed near the airplane. In case the airplane is parking on concrete

apron, then the airplane must be earthed.

(c) When using the ladder for the access to the wing upper side, special care must be paid to

prevent the fall down. Rubbersoled shoes must be used when moving round the wing.

(d) When handling the filling hoses care must be paid to avoid any damage on the rubber

deicers of the wing leading edges

16.10. SERVICING IN EMERGENCY SITUATIONS

16.10.1. General

Thls sixtion comprises the descr~ption of service at smergency situar~ons after an emergency

landing, when the crew and passengers must be rescued by means of emergency means and

the airplane must be removed from the runway.The following technological procedures are to

be taken up:

- penetration into the airplane and passengers rescue on the board

- extinguishing the on-fire airplane on the ground

- a~rpiane removal after landing with the landing gear

- airplane lifting with one-side retracted main landing gear

16.10.2. Pene~ration mto the airplane and passenger rescue on the board

(a) In case of door blocking (i.e. nor the main entrance door neither the emergency ex~t - see

fig. 16-28 can be opened - pos. 1 and 2), penetrate into the airplane from outside by

cdtting a square opening in the fuselage rear section from the !eft-hand side at the place

whlch is marked by yellow marks (3).

YE-E: Emergency square opening can be made by means of manual operated power saw

There are no installations in the indicated area, damage of which may endanger the health

s:::,?r of p?rsocs the airplar,e or of ;escuir;g ~erscnne! !t -us: bx taken intc cmsideraticn.

thcugh, that on the inside wall the passangers wear may be hung.

(b, h e to the fact that the opening for the forced penetratlon Into the a~rplane IS close to the

,:round. when the a~rplane is In normal poslt~on, there are no ladders nor the steps or

sllpways requ~red for rescurlng the passengers

(c) 'Nith regard to the width of the emergency exit (970 mm) only two workers of the rescuing

squad may work together with axes. \Nhen using powered circular saw ~t is advantageous

,i the rescue worker works alone. The safe access way to the place for the forced

penerat~on into the airplane is from the left rear side.


FIG. 16-28 PWCE FOR FORCED PENETiiATlON INTO THE AIRPLANE

(1) Main entrance door, (2) Emergency exit. (3) Place for forced penetration into

the airplane, (4) Emergency exits (if installed)

16.10.3. Extinguishing the on-fire aircraft on the ground

(a) For extinguisning the on-fire airplane the iocation of ihe inflammable mater~als containing

tanks ~nside :he airplane is decisive. This 'ccation is acparent frcn the fig. 16-29. The

decisive amcct of inflammables is represented by the fuel (aviation petrol), stored in the

wings. It is aiso very important for fire extinguishmg, that the fuei jwmg up fuei tanks

excluded) is stored in bag tanks, this being also inflammable. Hence the main stream of

fire ext~ngu~sk~ng agent must be directed :owards the wmg and cue the location of forced

penetration into the airplane it must be first of all directed to the left-hand wing.

FIG. 16-29 LOCATION OF INFLAMMABLE MATERIAL TANKS

(1) Fuel tanks in left-hand wing, (2) Fuel tanks in r~gnt-hand wing, (3) Terminal

fuel tank in left-hand wing, (4) Wing tip fuel tank in nght-hand wing, (5) Main

hydraulic tank. (6) Hydraulic accumulators of the mam network, (7) Brakes

hydraulic accumulator. (8) Emergency hydraulic tank. (9) Oil tank of the

r:cnt-hand engine. (10) Oil tank of the left-hand engine (part of the engine.

pos.no. 9 and 10).


(b) In case of fire in engine nacelle, the fire is to be extinguished with ground extinguishing

equipment sum a way, that the hose of the ground extinguishing equipment will be driven-

in into special lid on the engine nacelle, wh~ch will fall out by hose strike.

16.10.4. Airplane removal after landing with the landing gear up

(a) In the retraced position (see fig. 16-30) the landing gear wheels are not completely

retracted in the landing gear nacelle. On the extending section of the landing gear wheels

there are still more brake rollers, these being in case of landing with retracted landing

gear the source of increased friction and will be together with the adjacent wheel structure

destroyed or heavily damaged during the landing.

FIG. 16-30 RETRACTED POSITION OF THE MAIN LANDING GEAR

(1) Tyre, (2) Brake rollers

i5) Remove :he airplane with retracted landing gear as follows

- make sure that the propellers do not rotate

- place the ropes accordmg to fig. 16-31 around both landing gear nacelles - min.

lenght sf 20 rn. The contact oomts of ropes w~th front edge of the landing gear

nacelles nave to be protec!ed w~th rubber belt of dimens~on 0.5 x 1 m and the

th~ckness 5 - 8 mm. or other suitable mater~al. When using the steel ropes, add stdl

more wooden bars of dimensions 25 x 50 mm lenght 0.5 m


- connect the ropes behind the fuselage and attache them to the towing vehicle

- to the airplane rearward (i.e. tail ahead)

NOTE: - There is no diffrence between the ways of airplane towing from the grass or

concrete surface

- due to relatively low weight of the airplane and thas small specific pressure,

the airplane can be towed for a shorter distance without loading it on a

special transport bogie

- even when following all precautions concerning the protection of the landing

gear nacelles, fracture of their skin on front edges can occur during the

towing.

FIG. 16-31 TOWING THE AIRPLANE WITH RETRACTED LANDING GEAR BY THE

TOWING VEHICLE


16.10.5. Airplane lifting with-one side gear retracted

(a) Lift the airplane with one-side main leading Sear retracted as mentioned below:

- put the chocks under the retracted wheel o i the main landing gear from both sides

- adjust the propeller manualy in order not to obstruct the movement in the vicinity of engine

nacelle

- drive-in with high lift truck of capacity min. 1500 kg and the lift min. 2600 mm on the outer

side of engine nacelle (in the distance at least 0.5 m from the engine nacelle) to the area

in front of wing leading edgeon the-side of relracted landing gear-see fig. 16-32. - - - - -

- extend the high lift truck fork to the elevaticn of about 0.4 - 0.5 m under the wing lower

edge

- position the wooden prisms of size 10x1 Ox: 30 mm such a way that after their pressing to

the wing they lay in the centerline of front and rear beam (centerline to be recognised

according to rivets row - the forder one approx. 0.5 m from the leading edge. the hinter

one approx. 1.3 m from the leading edge). In case the truck fork is not long enough, the

support in place of front beam would be sufficient, but care must be paid so that the fork

free ends will not demage the skin

- lift the fork of the truck such a way that the prism will be slightly forced-to the wing lower

side. Check the prlsm position and repair possible defects - - - ~ ~ ~

- check the reliability of blocking of retracted landing gear wheel against movement

- lift the airplane into position sufficient for landing gear extension in small steps and

continuously check the proper wlng position on the high lift truck

- by the means of assembly lever remove the landing gear from the landing gear nacelle

and take it out manually to the position when the operating cylinder lock is Iccked. Work

has to be performed by skilled aircraft maintenance engineer

- lower the truck for to allow the airplane seat completely on the landing gear (in case , when

the aircraft landing gear cannot be lowered or it is heavily damaged and unmovable, then

place the special bogie, used for crashed airplane transportation under the landing gear

nacelle).Take away the wooden prisms and drive the truck away

,Remove the wooden chock from-under-the- landing gearwheel

NOTE: - Such prepared airplane is ready for towing in accordance with item 16.10.4

- In case that the hlgh lift truck is not available, carry out the lifting with suitable jack,

supported in the same points. Mobile crane may be used, too. In such case the fork of

high lift truck has to be replaced by wooden beam placed under the prism and sufficiently

overhanging the leading and trading edge (for the safe attachment into the suspension

rope).

AIRCRAFT TRAINING MANUAL L 41 0 UVP-El E9. E2O

FIG. !8-32 .AIF2L.ANE LIFTING 9" ME.4NS OF HIGF LIFT TRL'CK

(1) Wooden prism, (2) Chocks under the wheel. (A) The wing detail at the high

lift truck fork point (cross section)


16.1 1. GROUND EQUIPMENT AND TOOLS Review of tools

Rewiew of tools specified in the table has been compiled as follows:

A - tools for the airplane technician (engineer)

B - tools for the aierame

C - tools for the fuel system

D - tools for the control system (Not delivered in 1 : 1 set)

E - tools for the landing gear and hydraulic system

F -tools for heating, a~rcondit~oning and deicing system (Not delivered In 1:1 set)

G - tools for electric equipment

H -tools for radio equipment

I - tools for instrumentation

NOTE. - Tools for the engine and propeller is not incorporated in the tool review. Review of

th~s tool is specified in manufacturer s documentat~on of this products.

- For cargo and medevac airplanes additional tools are supplied as follows:

spanner No. 8 - B 097 602 N, spanner No. 9 - B 097 603 N. spanner No. 10 - B 097 604 N, spanner No. 12 - B 097 605 N, spanner No. 13 - B 097 606 N.

Tool set

1 .Double-ended spanner 5 . 5 ~ 7 CSN 23061 1.6

6x9 CSN 23061 1.6

8x10 CSN 23061 1.6

11x12 CSN 23061 1.6

13x1 7 CSN 23061 1.6

14x1 7 CSN 23061 1.6

19x22 CSN 23061 1.6

19x24 CSN 23061 1.6

24x27 CSN 23061 1.6

24x30 CSN 23061 1.5

5 . 5 ~ 7 TONA 611

6x9 TONA B1 1

8x1 0 TONA 61 1

11x12 TONA611

1.3~17 TONA 61 1

t4xl.5 TONA.611

6 CSN 230626.6

8 GSN 23U626.6

9 CSN 230625.6

9 CSN 230626.6

17 CSN 2306266

22 CSN 230626:6

27 CSN 230626.6

30 CSN 230626.6

32 CSN 238626.6

36 CSN 230626 6

41 CSN 230625.6

50 CSN 230625.7

10 CSN 230626.6

3. Spanner for screws

fixing the seat B 096 039 N

4. Spanner of fuel

cleaner hose B 096 404 N


5. Spanner of lock safety

pin of landing gear

6. Main wheel disc spanner

7. Spanner for airfilter

8. Spanner

9. Special spanner

10. Control column spanner

1 1. Open spanner

12. Spanner

13.Spanner

14. Eye spanner

15. Double-ended barrel wrench

16. Box wrench

B 096 322 N

L 410.9559

6 097 730 N

434-601 -Dl

B 096 018 N

B 096 106 N

B 096 121 N

B 096 111 N

B 096 320 N

B 096 321 N

6x9 CSN 23063T.7

14x17 CSN 230637.7

8x10 CSN 230653.T

11x12 CSN 230653.7

14x17 CSN 230653.7

9 CSN 230651.7

10 CSN 230651.7

12 CSN 230651 7

14 CSN 230651.7

17 CSN 230651.7

19 CSN 230651.7

27 CSN 230651.7

41 CSN 230651 .T

L 410.9143-02

L 410.9143-04

L 410.9143-05

L 410.9143-07

L 410.9144-09


17. Socket wrench

18. Open spanner

19. Wrench

20. Screwdriver

21 Screwdriver

22. Modified screwdrwer

23. Automatic screwdriver

24. Gross cut screwdrlver

B 096 341 N

5 396 120 N

6 0 9 6 113 N

7 GSN 230650 7

8 CSN 230650 7

10 GSN 230650 T

L 410 9143-03

B 097 408 N

8 a97 409 N

No 697 3 5x80

No 697 3 5x100

No 697 4 5x1 20

No 697 6x1 20

'vo 658 3x162

No 698 12x160

No 710 2 3x105

No 710 3 5x105

No 71045x105

L 410 9141-03

L 410 914145

L 410 9141-06

L 410 9146-04

B 096 480 N

Z 37 9110-10

870011, 2 3

No 716 slze 1

No 716 s~ze 2

No 716 slze 3

No 716 slze 4


25. Combination pliers

26. Pliers with long flat jaws

27. Wire cutting pliers

28. Cutting wippers

29. Pipe tongs

30. Rudder assembly pliers

31. Pliers for Seger rings

32. Pipe tongs

23. Side xrt : ; lg wippers

34. Pliers

35. Flat scraper

36. Packing of lever press hose

37. Hammer

38, Centre punch

39. Dural mandrel

40. Handle

dl . Assembly mandrel for elevator

-1:. Metal sheet shears

43. Portable lamp

44 Fixture for 'iyre d~srnantling

180lrubber CSN 230382

CSN 230363.41

Type 3211160

No. 3013

No. 357

L 410.91 10-02

No. 2327

200 CSN 23041 3.2

CSN 230327.1

155 CSN 230343.2

20 CSN 229435

6 096 540 N

300 CSN 2301 10

B 096 481 N

Z 37.91 10-01

Z 37.91 10-02

Z 31.91 10-03

4 GSN 230659.7

5 CSN 230659 7

5 CSN 230659 7

8 CSN 230659.7

10 CSN 230659.7

12 CSN 230659.7

16 CSN 230659.7

B 096 378 N

L 410.9141-02

250 GSN 2261 16.1

L 410.9618

K 20-7100


45. Pressure gauge

46. Small measuring instrument

47. Brush

48. Brush for collector cleaning

49. Soldering iron

50. File

51. Needle file

52. Magn~fytng glass

53. Lever press

54. Lubricating hose

55. Dial calliper

56. Flat chisel

57. Dynamometer

58. Extension piece

59. Head 12.10.9,8.7

60. Extension piece

61. Thickness gauge

62. Screwdriver

B 596 336 N

PU 110

8/50 ON 233710

12/50 ON 233710

20150 ON 233710

No. 82411

L 410.9126-03

150 ON 2291 10.3

150 ON 221960.3

120 ON 229180.3

No. 6

500 CSN 231462'

500 CSN 231492.

150 GSN 251238

150 CSN 232820.1

L 410.9712

B 097 514 N

B 097 515 N

B 097 516 N

B 097 521 N

B 097 522 N

B 097 523 N

B 097 524 N

B 097 525 N

B 097 51 1 N

0.05 - 1.00 x 100 CSN 251670

B 596 634 N


Review ground equipment 1:l

:.Blocking S i i ~ t

2. Elevator and ailerons blocking

3. Rudder blocking

4. Propeller Mocking

5. Wheel chock - - - - - -

6. Mooring device of the nose wheel

7 . Moooring device for the wing

8. Cock extension

9. Steps

10. Exhaust cover

: 1. Airplane canvas complete

12. Extensicr; 'or cabin heating - - - - - - - - - - - - -

13. Map for certificates

14. Assembiy platform

15. Map for cockpit documentation

16. Hand operated towing device

17. Towing device to a tractor

18. Hydraulic liquid discharge hose

19. Oil discharge hose - - - - - - - - - - - -

20. Oil filling funnel

21 .Hose

22. Vat

23. Fuel discharge hose

B 096 476 N

B 922 150 N (L 41 0.9220)

B 596 790 N or B 096 107 N

B 596 276 N or B 596 695 N

B 925 028 N (3 pcs) or B 596 895 N (3 pcs) (XL 41 0.9250)

B 596 177 N (till the 19th series) or B 596 671 N (from the 20 th ser~es)

B 596 278 N (till the 19th series) or B 596 670 N (from 20th ser~es)

B 928 265 N (XL 41 0.9280)

B 596 281 N

B 596 008 N

B 596 279 N or B 596 687 N

I3 935 760 N (L 410.9351)

B 004 035 N (L 410.9380 W D )

B 596 331 N

B 004 036 N (L 41 0.9390 \N D )

B 952 053 N (XL 410.9521)

8 952 200 N (L 410.9525)

B 954 089 N (XL 410.9543)

B 954 095 N (L 41 OM 9544)

B 096 242 N

B 964 201 N (L 4 10 9643)

B 961 805 N L 410.9617

B096 116 N


24. Rear support strut B 962 104 N (XL 410.9620)

25. Distilled water funnel

26. Control lever blocking

27. Rear support strut

28. Convas on the cockpit windshield

29. Toilet venting cover

30. Blind flange of static

31. Exhaust cover

32. Vessel for oil

33. Vessel for kerosene

34. Breaking piece

35. Anchoring pin

36. Lever press

LDN 691 1

9527 3405 500 CSN 231462

37. Lubrication hose 9527 3417 1 000 CSN 231492.3

33. Cockpit hydraulic jack 9527 3996 (Hz 4 - 3 execution 2)

39. Material for cockpit jack fixing:

- screw

- packing B 591 453 N (L 410.9515-02)

8397 1181 (18 x 250 x 5 ONL 3401.1)

- belt with claps

3185 0418 (M4 x 18 ONL 3147)

- screw

- washer 3570 1008 (8.2 CSN 021740.14)

- washer 3555 1004 (4.3 CSN 021 702.14)

B 097 365 N JO. Steps

41. Pitot tube cover

42. Packing of lever press hose

33. Airplane earthing equipment

44. Air inlet cover


45. Ladder for fuel filling B 097 343 N

46. Steps B 097 300 N

47. Antenna of th II. band 9523 6605 (Block 12-1 2)

9801 4024 (No. 26)

9801 4025 (No. 28)

50. Connection of fuel mud removal B 096 591 N (for main fuel tank) - - - - - -

- - - -

51. Hose for hydraulic pumps venting B 097 485 N

52. Vessel for venting of hydraulic Pumps B 097 480 N

53. Hose for fuel draw-off B 097 441 N

54. Equipment for hydraulic system pressurizing

55. Nose cover strut

56. Calibrated vessel for distilled water

57. Front curtains L.H:, R.H.

58. Side curtain L.H., R.H.

59.-Additionahear mooring device

60. Additional front mooring device

61. Mooring device of tail unit

62. Winter cover of the oil cooler, front part

63. Winter cover of the oil cooler, rear pan

64. Oil cooler insulation

B 097 630 N (till the 17th series) or B 596 546 N (from the 18th series)

B 097 500 N

B 092 041 PIL

B 092 042 PIL

B 596 67x N

B 596 672 N

B 596 674 N

B 596 570 N

65. Air inlet cover of the starter B 596 580 N generator

66. Reduction piece B 596 525 N

67 Interconnecting piping B 596 568 PIL

AIRCRAFT TRAINING MANUAL L 4'10 UVP-E, E9, E20

Ground equipment-special set (if installed)

1. Propeller lifting device B 536 430 N

2. Stand for the propeller XL 41 0.9660

3. Engine suspension ropes L 410 M.9671

4. Truck for the engine L 410 M.9681

5. Fixture for the rudder angle B 096 022 N

measurement

6. Plate for putting aside of B 096 023 N

6 $96 023 N

7. Angle meter for the measurement B 596 301 N

of ailerons and elevator deviation

8. Hydraulic truck 6999 AF

9. Filter washing device L 410.9539

10. Filter washing device L 410.9540

1 1. Testing fixture B 596 455 N

12. Connection for fuel mud removal B 596 560 N

13. Hydraulic jack (hangar type) B 097 700 P.H.1L.H.

14 Assembly carpet L 410.9340

15. Device for checking and filling XL 41 0.9551

landing, gear dampers

16. Dynamometer L 410.9712

17 Dynamometer for clearance L 410.9624

checking of landing gear legs

18. Comtrol column locking L 410.9650

19 Truck for hydraulic liquid B 596 200 N

repien~shmg

20. Pip~ng blind flanges L 410.9626 W. D

3 Fixture for wiper assenbly B 096 024 N

22. Serv~ce truck L 410.9682

23. Carpet on the wing L 41 0.936219361


24. Window protective cover L410.9623

25. Checking fixture B 097 322 N

26. Test lamp B 096 221 N

27. Rubber wheel L 410.9631

28. Knurled wheel L 410.9632

29. Canvas for engine preheating B 596 614 N

30. Assembly bogie L 410.961 5

31. Device for washing the L 410.9591

compressor

32. Towing dev~ce B 097 582 N

33. Blocking fixture B 096 267 N

34. Ring remover of main wheel B 097 586 N

bearing

35. Ring remover of nose wheel B 097 592 N

bearing

36. Lock wasner pin remover of the B 097 327 N

main landing gear

37. Wheel remover of the nose wheel I3 097 550 N

38. Wheel remover of the main B 097 560 N

landing gear

39. Static pressure connector B 096 670 N

40. Connector B 096 672 N

41. Stall speed hole plug B 097 490 N


16.1 2. AIRPLANE MAITENANCE The airplane maintenance 'ncludes:

- routine maintenance

- periodic maintenance

- inspection of airplane

- seasonal maintenance

- unscheduled inspection

- maintenance during storage

(a) Routine maintenance (for the aircraft with overhaul)

Maintenance A

- follows immediately after each landing

- durmg the training flights this maintenance is done when refuelling only

Maintenance 8 is done-

- beiore each fhghi (after each landmgj if a h~gher type of mamtenance is not required

- beiore flight after the periodic maintenance was done

- d u r q tra~nmg (tipe ratlng) flights this check is done when refuelling only

Maintenance V is done:

- once a day, after finishing the flying operations, preferably at the base airfield or at the

arfield of the last landing of the day, if higher type of maintenance is not required

- during preparation for fl~ght, as a supplementary work to the Maintenance E if the pause

between flights is from 1 to 15 days, if the aircraft has not been prepared for storage.

Maintenance G is done:

- at the base airfield once every 7 days + - 1 day when the aircraft is in regular service

(i.e. one flight a day at ieast), or after every 50 + - 10 landings, provided the hours flown do

not make a higher type of maintenance mendatory. This period may be prolonged by the

number of non-flymg days, to a maximum of 10 consecutive days

- when putting the aircraft to flight status after storage

- after the test-flight follcw~ng an englne change

Mamtenance D is done.

- immediately before each flight, irrespective of the type of mamtenance already done


Maintenance E is done:

- immediately before the first flight, irrespective of the type of maintenance done.

Maintenance Z is done:

- when airplane is handed over to the flight technology base

- for parking, if the period of waiting before the next flight exceeds two hours

- for maintenance

- for towing the airplan-e to another parking aFea

(b) Routine maintenance

The routine maintenarct consists of:

Routine maintenance A

- done before every flight unless stated in the remark with the given operation othervise

Routine maintenance 8

- done after every landing and in case of handing over the aircraft for parking

R~utine maintenance S

- done once a day after concluding the flight day or before the first flight after an interval

between flights from 1 to 15 days.-This maintenance has validity-24-+ $2 hours. - -

- - - -

The routine maintenance type A and B is allowed to Se done by a pilot on condition that the

was trained for this work and approved by the manufacturer or else organization havmg the

certification for L 410 UVP-E, E9, E20 aircraft type checks.

In zase any defect c c x s during the routine maintenance A and 9 which is done by the pilot,

then these defects shall be repared only by a trained engineer.

(c) Periodic maintenance [for the aircraft with overhaul)

Periodic maintenance IS done according to the number of hours flown, and according to the

calendar tlme, in dependence on specific conditions of the exploitation of arcraft at the user

organizations.

Periodic maintenance consists of maintenance work of the basic type, done every - - -

300 5 30 fltght hours. 300 +S@landings, 6 zQ.6 calendar mothsof-ailframe operation tme

ard of the supplementary work's time. This work results from the number of hours flown.

number of landings and calendar time related to every 600. 900. 1200 etc. flight hours

(landings), to I? , 18. 24 ?tc. months. The whole cycle of period~c maintenance covers the

per~od of arcraft operarlon between two general overhauls. Domg the mamtenance work

according to the number of landings andlor calendary period does not rel~eve of doing

mamtenance work according to the number of hours flown.

--

(d) Periodic maintenance

Tile periodic ma~ntenance consis:s ;i:

- line check 1 - is done periodically after every 10 + 1 days.

- check 2 - is done after every 300 + 30 flight hours.

- check 3 - is done after every 1 200 2 30 flight hours.

- check 4 - is done after every 2 400 + 30 flight hours.

Remark: - With some systems it is necessary to carry out the check also depending on the

operation time (months) or on the number of landings. The time periodicity, or the

number of landings is given in the column Note of the specified operation. The

time works and works according to the number of landings may be done with the

next lower periodical check.

(e) Unscheduled maintenance

Unscheduled maintenance is done after flight in a region of storm activity, after a lighting

strike, after landing at a weight higher than the maximum landing weight and after a hard

landing.

In other exxtraordinary cases IS the amount and contents of work established accordins :o

the decision of a com~sicn

(f) Mamtenance durmg storage

After putting :he aircraft cut of operatlon for more than 15 days ~t IS necessary to execute

works connected w~th the preparat~on for the storage durmg storing and when puttmg the

arcraft mto operation

The %tent of ,vcrk conr5c:ed wth '-e preparat~on for the storage deoevh on the plarred

storage time

There are the followmg types of storage

- for a t ~ m e of 15 to 30 days

- for a durat~on up to 3 monthss

- for a duration more than 3 months (I e 6 months, 9 months, 12 months)

The arcraft must be stored provlded wth all blmds and covers

Sass2 3n results 3 the 1nscec:lon it is -ecessary to carty 3ut the rerars requ~red to ensLrl 'he

adequate level of 'light safety and coni,nued arworthmess of the arcraft

The insoection c i :he arcraft IS carried sut accordmg to an approved moec t~cn program ~n the

arcrar: manufacturer s facilvty or ln 3r wthor~yed servlce facllity

The lnspectlon dependmg on the operatron tlme may be carr~ed out w~thm 5 15 days of the

specifled tlme hmit Inspecttons dependmg on the number of landmgs withm + 50 landmgs of the

specifled limit if the ~nspection interval for a spec~f~c check IS d~fferent from that glven for checks


1, 2, 3, and 4, a multiple of the inspection period is shown at the particular check, which is

defined in a footnore at the xt tom of the corresponding ;age (e. g. 2+ .... means that this is t3

be done at every other check 2).

However, the accomolisher of maintenance depending on the number of landings or the

calendar does not relieve the operator from the duty of accomplishing the work specified affer a

certain number of flyght hours.

Lubrication is to be done during maintenance checks of every type, in accordance with the

lubrication chart (lubrication dependingon calendar period, lubrication at an interval other than

300 hours).

Time limit for each periodic maintenance check is to be always counted

- starting from the entry of a new aircraft into service

- starting from the last carried - out mayor inspection of the aircraft


CHAPTER 17

AIRCRAFT CHARACTERISTICS

17 1 Basic Performance

17.2. Flight Technical Performance

17 3 Operation Performance




17.1 BASIC PERFORMANCE Aircraft Easic Technical data

Type designation

Design

Category of Utility

Number of seats:

Crew

passengers

Power plants

Turboprop engine, two-shaft ccaxial with free turbine.

reverse air flow, gearbox, type INALTER M 601 E

Number of engines

All metal propeller, five blade w~th speed regulator

adjustable feather and reverse ;Itch, type V 510

Number of propeller

Fuel system

Main system - number of wing fuel cells

Auxiliary - number of wing tip tanks

Total fuel volume in the wing tanks

Total fuel volume In the wingtip tanks

Maximum value of nonconsumable fuel

- in wing tanks

- in wing tip tanks

L 410 UVP-E, L 410 UVP-E9,

L 410 UVP-E2O

ground airplane

Public traffic service,

operation in approved aircraft

modification (versions)

2

19 (when the aircraft is

operated in various versicns

the number of passengersis

{educed according to a

specific aircraft configuration)

8 (4in each wing half)

2 (1 at the wing tip)

1000 kg

313,8 kg

Fuel pre-heater


Oil system

Oil cooler

Thermostatic valve

Hydraulic system

Nominal pressure

Volume of main hydraulic tank

- - -

Volume oferriergenq tank- -

Air working pressure in the hydraulic tank

Air pressure from engine compressor to overpressure

hydraulic tank

Pressure in the parking brake circiut

Operational pressure In brakers

Landing gear

a )Ma in lanbnggear - - - -

Tyres

Brakes

b)Nose landing gear

- - - - - -

Shimmy damper

Tyre

14.7 MPa (150 kplcm2)

10 1 - - - - - - -

3.2-1

0.1 +0.12/ -0.03 MPa

(1 +1.2/ -0.3 kp/cm2)

max 0 . S MPa

(max 5.5 kp/cm2)

4.9 -0.2 MPa (50 -2 kplcm2)

0 - 4.4 70.3 MPa (0 - 45 +3 kp/cm2)

- - - -

Left and right legwith external oil-

pneumatic twochamber shock

absorber with floating piston

12.5-10 model 4 electricaly

conduc:~ve TUBELESS operational

pressure: 420 +30 kPa

(4.3 +0.3 kp/cm2)

Disc hydraulically operated with

position for parking braking

- leg with oil - pneumatic shock

absorber- : - - - - -

- shock absorber is a part of

servosteering system

9.00 - 6 model 4 TUBELESS

electrically conductiveoperationaI

pressure: 420 -30 kPa

(4,3 -0.3 kp/cm2)


Airframe deicing system

Engine water injection system

Tank capacity

Quantity of injected water (delivered by pump)

Primary electrical system

a) Primary regulated DC power supply

Starter - generator

Nominal voltage value

b) Primary three phase alternating current supply

Nominal voltage

Nominal frequency

c) Secondary one phase alternating current suoply

Nominal voltage

Nominal frequency

d) Secondary three phase alternating current supply

Nominal voltage

Nominal frequeficy

Basic dimensions

Wing span

Overall lenght

Overall height

Distance between engine and airframe axis (spanwise)

Wheel track (operational position)

Wheel base (relieved positton)

Nose wheel turning

- by hand lever

pneumatic

3.3 1 first stage

6.6 1 second stage

10 1 third stage

19 479 mm

(19 980 mm with wing tip tanks?

14 424 mm

5 829 mm

2 408 +/-I0 mm

3 650 mm

3 666 mm

50° - 5O (to both sides)

4O30' +I-1°30' (to both sides) - by pedals


Wing area

Aspect ratio

Pilot cabin volume

Cockpit entrance door dimensions (left and right side of

the security partition)

Emergency exit dimension

Passenger cabin volume

Passenger cabin lenght

Passenger cabin width

- maximum

- minimum

Passerger cabin heicht

Entrance door dimension

Cargo aoor dimension

Isle widht (at heigh of 400 mm above floor)

Front baggage compartment volume

Front baggage compartment lenght

Front baggage compartment width

- maximum

- minimum

Front baggage compartment height

- maximum

- minimum

Aft baggage compacment vclume

Aft baggage cornpacment width

Aft baggage compacment herght

Aft baggage compartment depth

- maximum

- mmimum


Note: The values for aft baggage compartment apply to the baggage compartment located at the

rear of the passenger compartment, forward of frame No. 21, on the RH side

Weights

Basic weight (weight of equipment aircraft without cren) 4 000 +I-20 kg

of standart equipped aircraft (without fuel in the fuel

tanks, without wing tip tanks, without water to be injected

into the engines, without lifting jack and without

B 596 281 N entrance steps)

Maximum take-off weight 6 400 kg (L 41 OUVP-E, E9, E20)

Maximum landing weight 6 200 kg (L 41 OUVP-E, E9, E20)

Maximum weight without fuel 5 870 kg (L 410 UVP-E, E9)

5 900 kg (L 410 UVP-E20)

Centre of gravity

Operational C. G. range for L 41 0 UVP-E aircraft is 17 to 28%.

Operational C. G. range for L 410 UVP-E9, E20 aircraft is 17 to30'/0

Note: The max. take-off we~ght for the L 410 UVP-E9, E20 aircraft can be increased to 6,600 kg.

17.2. FLIGHT TECHNICAL PERFORMANCE

Strength Conditions

Speed designation

vmax. F

Name of speed

I Vrnax.MO I Max~mum operating speed

Operational load factor

-1.24 1

i

Max. permissible cruise speed w~th wing flaps extended to 18', 42'

+ 3.1

maximum positive

+ 2

-1.24 i

maximum negative


Limitation of operating speed

VmaX,rnax 1 Design limit speed (calculated) 1 400 only for L 41 OUVP-E9 I I

Speed designation Name of speed

v m a X , ~ 0

Speed in kmlh IAS

I Vmax,FE 18

1 v m a X . ~ o I Maximum permissible landing gear operating 1 250

Maximum operating speed

Vmax.FE 42

/ V m a x , s p 0 ~ ~ I Maximum permissible spoiler extension speed 1 190

350 (L 41 0 UVP-E) 335 (L 410 UVP-E9, E20)

Max. permissible cruise speed with wing flaps extended to 18'

* At speed v=205 krnl!, !AS is extension blocket

250

Max. permissible cruise speed with wing flaps extended to 42'

Flight performance

220*

Flight performance are val~d under foilowmg conditions.

- international standart atmosphere

- maximum take-off we~ght 6 400 kg

- customer bleed is off

Both engines are operation

(1 j Real take-cff fur; 2: jpeec! >IF? = 150 kmlh IAS (161 kmlh EAS) *450 m

Note: Maximum take-off engine rating, propeller speed: 2 080 RPM, take-off configurationmg

flaps 18O, aerodrome altitude 0 m ISA, no wind, concrete runway steep: 0.

(2) Real landing run after landing speed of 140 kmlh IAS when using wheel brakes and reverse

thrust of both engmes *240 m

* Required runway length 800 m (1 00 m clearway)

Maximum landing weight is 6 200 kg, both,engines ~dle, landing configuration -wig flaps

extended to 42O, spo~lers extended before touch-down (as per Flight Manual), airport alt~tude:

0 m ISA, no wmd. concrete runway steep 0, approach speed vapr= 155 kmlh IAS

(1 6 1 kmlh EAS)


(3) Climb rate

Maximum continuous engine rating

Propeller speed: 1 900 RPM, cruise configuration (wing flaps and landing gear retracted)

I I

Flight altitude I Climb time from 400 m curcuid in m I

i minutes Torque %

(4)Stall speed

Engines idle. straight flight

Wing flaps position 1 Stall speed IAS kmlh

I 00 cruise 133 +I-5

(5) Range

Stall speed EAS kmlh

155

I 42' landing 102 +I-5

Range value is valid under following conditions:

J 121

- 8 marn fuel tanks in the wing

- max. take-off weight 8 400 <g

- ISA atmospheric conditions. no wind

- two engines operative

- levelled flight block altitude 4 200 m

- cruise speed 365 kmlh IAS

- climb rate at climbing from 400 m to 4 200 m: 3 mls

- rate at descent from 4 200 rn to 400 m: 3 mls

- propeller speed 1 900 RPM

- descent at engine ratrng: cruse speed 310 kmih, propeller speed nprop = 1 700 RFJl

- range tolerance: 3%

- crurse configuration (wing flaps retracted)

AIRCRAFT TRAINING MANUAL L 41 0 UVP-E, E9, E2O - --

(a) Range with maximum payload 1 615 kg and 625 kg of fuel (without wing tip tanks)

- with fuel reserve for 30 minutes at speed 287 kmlh IAS

at altitude of 2 100 m 546 km

- with fuel reserve for 45 minutes at speed 287 kmlh IAS

at altitude of 2 100 m

(b) Maximum range without wing tip tanks and

payload of 1 330 ks and 91 0 kg z i fuel

- with fuel reserve for 30 minutes at speed of 287 kmlh IAS

at altitude 2 100 m

(c) Maximum range w~rh wing tip t a r x w~th

payload of 885 kg and 1 300 kg of fuel

- with fuei reserve for 30 mimutes at speed 287 kmlh IAS

at altitude 2 100 m

Flight with one engine in operative

(1) Climb speed

Intermediate Continquency Power Setting

Propeller speed: 2 080 RPM

Left engine stopped - feathered propeller, flight speed

v = 200 kmlh IAS (210 kmlh EAS), cruise configuration

(wig flaps and landing gead gear retracted)

17.3. OPERATION PERFORMANCE

Operation performancs has been implemented into the design of the airplane in particular:

F!ight altitude m

~ Q O

1 000

2 000

3 000

- by reliao~lity of ~ n s t r ~ r e n t s and agregates

- standby agregates a m systems

Climb lime from 400 m c~rcuid in I minutes I Torque Oh

- sufficient techncal endurance of mdividual elements of airframe and systems (related to the

airplane as a whoie)

1 .o 6.0

19.0

40.3

< l o o i I

< 100

< 100 I

< 100 1


- low service and maitenance manhours thanks to easy accessible instrument and aggregates,

their interchangeability and contrcllability.

Operation technical performance:

- Aircraft Service life 20 000 flight hours

20 000 landings



Documents

Aircraft Training Manuel LET 410 UVP-E