Mod 15 (b1) Gas Turbines

JAR66 Module 15. JET ENGINES.

CONTENTS

15.0 INTRODUCTION

15.1 FUNDAMENTALS 3

15.2 ENGINE PERFORMANCE (B1) 15

15.3 AIR INTAKES 30

15.4 COMPRESSORS 44

15.5 COMBUSTION SECTION 58

15.6 TURBINE SECTION 65

15.7 EXHAUST 74

15.8 BEARINGS AND SEALS (B1) 83

15.9 LUBRICANTS AND FUELS 91

15.10 LUBRICATION SYSTEMS 95

15.11 FUEL SYSTEMS 106

15.12 AIR SYSTEMS 119

15.13 STARTING AND IGNITION SYSTEMS 128

15.14 ENGINE INDICATING SYSTEMS 138

15.15 POWER AUGMENTATION SYSTEMS (B1) 150

15.16 TURBOPROPELLER ENGINES 153

15.17 TURBOSHAFT ENGINES 163

15.18 AUXILIARY POWER UNITS 177

15.19 POWERPLANT INSTALLATION 188

15.20 FIRE PROTECTION SYSTEMS 206

15.21 ENGINE MONITORING & GROUND OPERATION 211

15.22 ENGINE STORAGE & PRESERVATION (B1) 226

Mod15 Gas Turbine Engines by COBC.

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INTENTIONALLY LEFT BLANK

Mod 15 Gas Turbine Engines by COBC 2

MODULE 15. GAS TURBINE ENGINE

15.1 FUNDAMENTALS

INTRODUCTION

The conquest of air by powered flight was ever the aim of man, and a great step forward was made by the Wright Brothers at Kittyhawk, America by their historic flight in 1903. Since this early date, aircraft have developed steadily and, in 1939, aircraft speeds of 464 mph were achieved by production aircraft. Aircraft could climb 56 000 feet and fly distances up to 7 000 miles non stop. At this time, international records in speed, altitude and endurance had all been set by Great Britain.

In attempts to improve aircraft performance, engines were increased in both size and power output, with various configurations being tried (e.g. various in-line and radial engines with from 7 to 36 cylinders per engine). Superchargers with coolers, water-methanol injection systems and many aids to performance were introduced. However, piston engines and propeller combinations suffered a loss in performance at high forward speeds and high altitudes; clearly a new type of aircraft propulsion unit was needed if aircraft performance was to advance even more; thus the jet engine (gas turbine) was born.

It is generally acknowledged that, in Great Britain, Sir Frank Whittle of the Royal Air Force designed and developed the first British gas turbine engine that was suitable for aircraft propulsion. Sir Frank was born in 1907 and he entered the Royal Air Force as an apprentice. As an apprentice he gained a cadetship to Cranwell College and, whilst there, he become interested in the prospect of jet propulsion for aircraft. He produced design drawings for a gas turbine engine and his first engine ran on static tests in 1937. In 1941 the Whittle gas turbine engine powered the Gloster E28/39 aircraft and many of the present-day Rolls-Royce Aero engines are developments of Sir Frank's design. Aero gas turbine engines have been the foundation, which has made modern high performance aircraft possible.

PRINCIPLES OF JET PROPULSION

SIR ISAAC NEWTON

Jet propulsion is a practical application of Sir Isaac Newton's third law of motion which states "For every force acting on a body, there is an equal and opposite reaction".

A fireman's hose is an example where reaction is felt. When a powerful jet of water is ejected from a hose, the hose tends to react and move away from the water jet and, so great is the reaction, that sometimes two men are needed to hold the hose and direct the water jet.


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HERO'S ENGINE (Fig 1)

The earliest known example of jet reaction occurred during the use of a toy called 'Hero's engine'. In 120 BC this toy showed how the momentum of steam issuing from a number of jet outlets could impart an opposite reaction to the jets themselves, and in doing so cause the engine to revolve. When this principle is applied to aircraft propulsion, the 'body' upon which the force acts is the atmosphere. Air is introduced into the intake duct of the gas turbine engine and then a force is applied to cause the air to accelerate within the engine.

The force which accelerates the air reacts in the opposite direction on the engine and moves the engine away from the accelerating column of air in the same manner as the fireman's hose moved away from the water jet.

Hero’s EngineFigure 1.1

JET REACTION

Jet reaction is an internal phenomenon and it is not, as sometimes assumed, the result of the jet efflux impinging upon the atmosphere. The jet engine is designed to accelerate a stream of air to an exceptionally high velocity and to obtain useful thrust from the reaction. There are many ways of increasing the velocity of the air but, in all cases, the resultant reaction is the propulsive thrust exerted on the engine. The thrust obtained is proportional to the mass of air passing through the engine and to the velocity increase of the mass of air flow, i.e. momentum = mass velocity. Thus, the same amount of propulsive thrust can be obtained by either:

Accelerating a large mass through a small increase in velocity.Or: Accelerating a small mass through a large increase in velocity


Jet ReactionFigure 1.2

FUNDAMENTALS

POTENTIAL ENERGY

Potential energy is the stored energy possessed by a system, because of the relative positions of the components of that system. If work done raises an object to a certain height, energy will be stored in that object in the form of the gravitational force. This energy, waiting to be released is called potential energy. The amount of this energy a system possesses is equal to the work done on the system previously.

Potential energy can be found in forms other than weights and height. Electrically charged components contain potential (electrical) energy because of their position within an electric field. An explosive substance has chemical potential energy that is released in the form of light, heat and kinetic energy, (see below), when detonated.

KINETIC ENERGY

Kinetic energy is the energy possessed by an object, resulting from the motion of that object. The magnitude of that energy depends on both the mass and speed of the object. This is demonstrated by the simple equation:

Energy = ½mv2

(Where ‘m’ is the mass of the object and ‘v’ is its speed in feet or metres per second)

All forms of energy convert into other forms by appropriate processes. In this process of transformation, either form of energy can be lost or gained but the total energy must remain the same.


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NEWTON’S LAWS OF MOTION

The full details of Newton’s Laws of Motion were covered in Module 2 (Physics), but as a reminder, they are listed below:

FIRST LAW

This law states that a body at rest tends to stay at rest and a body in motion tends to remain in uniform motion, (straight line), unless acted upon by some outside force.

SECOND LAW

This law states that the acceleration produced in a mass by the addition of a given force is directly proportional to the force and inversely proportional to the mass.

It can be demonstrated by the formula: FORCE = MASS X ACCELERATION

THIRD LAW

This law states that for every action there is an equal and opposite re-action.

BRAYTON CYCLE

A gas turbine engine is essentially a heat engine using a mass of air as a working fluid to provide thrust. To achieve this, the mass of air passing through the engine has to be accelerated, which means that the velocity, (or kinetic energy), of the air is increased. To obtain this increase, the pressure energy is first of all increased, followed by the addition of heat energy, before final conversion back to kinetic energy in the form of a high velocity jet efflux.

The working cycle of the gas turbine engine is similar to that of the four-stroke piston engine. There is induction, compression, ignition and exhaust in both cases, although the process is continuous in a gas turbine. Also, the combustion in a piston engine occurs at a constant volume, whilst in a gas turbine engine it occurs at a constant pressure.


The cycle, upon which the engine functions, in its simplest form, is the Brayton cycle, which is represented by the pressure/volume diagram, shown below.

The Brayton CycleFigure 1.3

The points A,B,C and D on the previous graph show the action of the pressure and volume of the charge during the cycle.

A B: Compression from atmospheric at ‘A’ to maximum at ‘B’.

B C: Combustion with heat being added.

C D: Expansion through the turbine and jet pipe.

During the C D part of the cycle, some of the energy in the expanding gasses is turned into mechanical power by the turbine; the remainder on its discharge to atmosphere, provides the propulsive force. Later in the notes, turbo-propeller engines will be covered which reverse the above statement. With these engines, the turbine(s) remove the majority of the power from the exhaust, to drive the propeller, leaving little residual thrust.


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PHYSICS TERMINOLOGY

The relationship between force, work, power, energy, velocity and acceleration were covered during Module 2 (Physics). The next section will be a revision of those physics notes.

FORCE

Force may be defined as a push or a pull upon an object. The units of force are the Newton, (Metric) and the Pound, (Imperial). The Newton is the force required to accelerate a mass of 1 kilogram (kg) at 1 metre per second, per second, (m/s2).

WORK

Work is the application of force to a body and the displacement of that body in the direction of the force. It can be demonstrated by the simple formula shown below:Work (W) = Force (F) x Distance (D)

The Metric system unit is the Joule, (One Joule being a force of 1N acting through 1m). One Joule therefore being 1Nm. (Newton-Metre), the Imperial measurement being the foot-pound (ft-lb.).

POWER

The rate of doing work. It is the work done per unit time and could be shown as the formula:

Power = Work doneTime to do the work

Work can be expressed in a number of units such as foot-pounds per second, Horsepower and Watts. The foot pounds per second are self explanatory, but theHorsepower is defined as 550 ft-lb. per second or 33,000ft-lb/min. The Watt is equal to 1/746 hp, therefore, 746Watts equals 1hp.

ENERGY

The term energy is defined as the capacity for doing work. As mentioned earlier, there are two forms of energy, potential and kinetic. Whilst there are many different types of energy, they are ALL either potential or kinetic.

VELOCITY

It is common to find people confusing the terms velocity and speed when describing how fast an object is moving. The difference is that speed is a scalar quantity, whilst the term velocity refers to both speed and direction of an object. The full definition of


velocity is that it is the rate at which its position changes, over time, and the direction of the change.

The simple diagram below shows how the aircraft, which flies the irregular path from ‘A’ to ‘B’ in an hour, (a speed of 350 mph), has an actual velocity of 200 mph in an East-Northeast direction.

Velocity/Speed DiagramFigure 1.4

ACCELERATION

This term describes the rate at which velocity changes. If an object increases in speed, it has positive acceleration; if it decreases in speed, it has negative acceleration. A reference to Newton’s Second law of Motion will explain the principles of acceleration. Acceleration can be in a straight line, which is referred to a linear acceleration and it can apply to rotating objects whose speed of rotation is increasing, (or decreasing), when it is called angular acceleration.


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CONSTRUCTIONAL ARRANGEMENTS

TURBOJET ENGINE

The simplest form of gas turbine engine is the turbojet engine, which has three major parts; the compressor, the combustion section and the turbine. A shaft connects the compressor and the turbine to form a single, rotating unit. These engines produce thrust in the manner described in the Brayton Cycle.

The simplest turbojet engine is the unit shown below with a single centrifugal(Double Entry)compressor and a single stage turbine. This type of engine can still be found in certain special installations but generally, they have been superseded by engines with axial compressors and multiple stage turbines. The advantages and disadvantages of the two types of compressor will be discussed in depth later in this module

Simple Centrifugal Gas Turbine EngineFigure 1.5


SINGLE-SPOOL AXIAL FLOW TURBO-JET

The illustration below shows the inner workings of a single-spool axial flow turbo-jet, the next engine in the development of gas turbines. It will be noticed that it has, in this example, 17 compressor stages and 3 turbine stages.

Single Spool Axial Flow TurbojetFigure 1.6

TWIN-SPOOL TURBO-JET

A low-pressure section consisting of a compressor and a turbine assembly and a high-pressure section, also consisting of a compressor and turbine. This type of construction allows the two sections to run at different and more efficient rotational speeds.

BY-PASS ENGINE

Some engines of this twin spool construction can also be found with a by-pass duct, which passes some of the air from the rear of the first compressor around the combustion and turbine sections of the engine. The by-passed air joins the exhaust from the turbine section, which helps to improve the propulsive efficiency of the engine, to produce better specific fuel consumption and to make it a little quieter.

A large number of this type of engine are still in service, although high by-pass engines which are covered later, have generally superseded them. An engine of the twin-spool, low by-pass type is shown below which, with re-heat (afterburner) can be found installed on Concorde.


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Twin Spool By-Pass TurbojetFigure 1.7

TURBO-FAN ENGINE

A turbo-fan engine is an advanced development of both the conventional propeller and the by-pass principle previously mentioned. The turbofan accelerates a smaller mass of air than a propeller but it does accelerate a much larger mass of air than earlier pure turbojet or low by-pass engine designs. The turbo-fan engine, illustrated overleaf, has a large fan in front of the intermediate compressor, and is in effect, a low-pressure compressor. Basically, the engine is of a triple-spool construction, which means that there are three compressors and three turbines, the Low pressure (LP), Intermediate pressure (IP) and High pressure (HP) units, mounted on their respective spools.

As with the twin-spool set-up, the three main rotating parts of this engine all rotate at their optimum speeds, giving an even greater thrust and efficiency than the two spool designs. The high mass of air that passes only through the fan gives the engine a very high by-pass ratio.

Triple Spool Turbo-FanFigure 1.8

Turbo-fan engines in a range of sizes, are installed in aircraft varying from small business-jets to the largest transport aircraft including the Boeing 777 and the Airbus A340.


TURBO-SHAFT ENGINE

Turbo-shaft engines are different from the previous types of engine mentioned. These all produce power through a driveshaft, which is used to drive gearboxes when they are installed on helicopters. (Turbo-propeller engines will be covered later). Many engines of this type are modifications of turbo-propeller and turbo-fan engine designs.

The primary purpose of this design of engine is to produce shaft horsepower. This is often obtained by the addition of an extra, ‘power’ turbine, or free power turbine, which extracts power from the exhaust gasses and, is connected via a reduction gearbox, to an output shaft.

On some designs of turbo-shaft engines,(Direct Coupled) the output shaft does not have a separate power turbine, but is simply connected to the engine’s own turbine(s). This means that the output shaft will be rotating whenever the engine is rotating.

Whilst both type of turbo-shaft engines have been used most successfully to power helicopters, the ‘free-turbine’ design is the more widely used. This is due to the fact that the engine can be started whist the rotors are held stationary by a rotor brake mechanism. This allows ground runs and other work to be carried out with the engines running, but with the rotors, which can be dangerous to personnel on the ground, stationary.

The illustration below shows a twin-spool turbo-shaft engine with a two-stage free ‘power’ turbine, which, in this design, drives a shaft through the engine, leaving the drive connection on the ‘front’ of the engine.

Twin-Spool Turbo-Shaft (With Free-Power Turbine)Figure 1.9


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The final design that will be considered is the turbo-propeller engine, often called the turbo-prop. This design is very similar to the turbo-shaft engine, except the output shaft is usually driven through a reduction gearbox and connected to a propeller, as shown below.

Twin-Spool Axial Flow Turbo-PropellerFigure 1.10


15.2 ENGINE PERFORMANCE (level B1 only)

METHOD OF CALCULATING THE THRUST FORCES

The thrust forces or gas loads can be calculated for the engine, or for any flow section of the engine, provided that the areas, pressures, velocities and mass flow are known for both the inlet and outlet of the particular flow section.The distribution of thrust forces shown in the figure can be calculated by considering each component in turn and applying some simple calculations. The thrust produced by the engine is mainly the product of the mass of air passing through the engine and the velocity increase imparted to it (i.e. Newton’s Second Law of Motion), however the pressure difference between the inlet to and the outlet from the particular flow section will have an effect on the overall thrust of the engine and must be included in the calculation.To calculate the resultant thrust for a particular flow section it is necessary to calculate the total thrust at both inlet and outlet, the resultant thrust being the difference between the two values obtained. Calculation of the thrust is achieved using the following formula:

Thrust =

Where A = Area of flow section in sq. in.P = Pressure in lb. per sq. in.W = Mass flow in lb. per sec.VJ = Velocity of flow in feet per sec.

g = Gravitational constant 32.2 ft. per sec. per sec.


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CALCULATING THE THRUST OF THE ENGINE

When applying the above method to calculate the individual thrust loads on the various components it is assumed that the engine is static. The effect of aircraft forward speed on the engine thrust will be dealt with later. In the following calculations ‘g’ is taken to be 32 for convenience.

Compressor casing

To obtain the thrust on the compressor casing, it is necessary to calculate the conditions at the inlet to the compressor and the conditions at the outlet from the compressor. Since the pressure and the velocity at the inlet to the compressor are zero, it is only necessary to consider the force at the outlet from the compressor. Therefore, given that the compressor –OUTLET Area (A) = 182 sq. in.

Pressure (P) = 94 lb. per sq. in. (gauge)Velocity (vj) = 406 ft. per sec.Mass flow (W) = 153 lb. per sec.

The thrust

=

=

= 19,049lb. of thrust in a forward direction.

COMPRESSOR


International Standard AtmosphereMod15 Gas Turbine Engines by COBC.

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ENGINE THRUST ON THE TEST BENCH

The thrust of the turbo-jet engine on the test bench differs somewhat from that during flight. Modern test facilities are available to simulate atmospheric conditions at high altitudes thus providing a means of assessing some of the performance capability of a turbo-jet engine in flight without the engine ever leaving the ground. This is important as the changes in ambient temperature and pressure encountered at high altitudes considerably influence the thrust of the engine.

Considering the formula for thrust under “choked” nozzle conditions:

Thrust = )A +

It can be seen that the thrust can be further affected by a change in the mass flow rate of air through the engine and by a change in jet velocity. An increase in mass airflow may be obtained by using water injection and increases in jet velocity by using after-burning.As previously mentioned, changes in ambient pressure and temperature considerably influence the thrust of the engine. This is because of the way they affect the air density and hence the mass of air entering the engine for a given engine rotational speed. To enable the performance of similar engines to be compared when operating under different climatic conditions, or at different altitudes, correction factors must be applied to the calculations to return the observed values to those, which would be found under I.S.A. conditions. For example, the thrust correction for a turbo-jet engine is:

Thrust (lb.) (corrected) = thrust (lb.) (observed) x

Where P0 = atmospheric pressure in inches of mercury (in Hg) (observed)30 = I.S.A. standard sea level pressure (in Hg)

The observed performance of the turbo-propeller engine is also corrected to I.S.A. conditions, but due to the rating being in s.h.p. and not in pounds of thrust the factors are different. For example, the correction for s.h.p. is:

S.h.p. (corrected) = s.h.p. (observed)

Where P0 = atmospheric pressure (in Hg) (observed)T0 = atmospheric temperature in deg. C (observed)30 = I.S.A. standard sea level pressure (in Hg)

273 + 15 = I.S.A. standard sea level temperature in deg. K273 + T0 = Atmospheric temperature in deg. K

In practice there is always a certain amount of jet thrust in the total output of the turbo-propeller engine and this must be added to the s.h.p. The correction for jet thrust is the same as that specified earlier.

To distinguish between these two aspects of the power output, it is usual to refer to them as s.h.p. and thrust horsepower (t.h.p.). The total equivalent horsepower is Mod 15 Gas Turbine Engines by COBC 18

denoted by t.e.h.p. (sometimes e.h.p.) and is the s.h.p. plus the s.h.p. equivalent to the net jet thrust. For estimation purposes it is taken that, under sea level static conditions, one s.h.p. is equivalent to approximately 2.6 lb. of jet thrust. Therefore:

t.e.h.p. = s.h.p.

The ratio of jet thrust to shaft power is influenced by many factors. For instance, the higher the aircraft operating speed the larger may be the required proportion of total output in the form of jet thrust. Alternatively, an extra turbine stage may be required if more than a certain proportion of the total power is to be provided at the shaft. In general, turbo-propeller aircraft provide one pound of thrust for every 3.5 h.p. to 5 h.p.

Comparison between thrust and horse-powerBecause the turbo-jet engine is rated in thrust and the turbo-propeller engine in s.h.p., no direct comparison between the two can be made without a power conversion factor. However, since the turbo-propeller engine receives its thrust mainly from the propeller, a comparison can be made by converting the horse-power developed by the engine to thrust or the thrust developed by the turbo-jet engine to t.h.p.; that is, by converting work to force or force to work. For this purpose, it is necessary to take into account the speed of the aircraft.

The t.h.p. is expressed as

Where F = lb. of thrustV = aircraft speed (ft. per sec)

Since one horsepower is equal to 550ft.lb. per sec. and 550 ft. per sec. is equivalent to 375 miles per hour, it can be seen from the above formula that one lb. of thrust equals one t.h.p. at 375 m.p.h. It is also common to quote the speed in knots (nautical miles per hour); one knot is equal to 1.1515 m.p.h. or one pound of thrust is equal to one t.h.p. at 325 knots.

Thus if a turbo-jet engine produces 5,000 lb. of net thrust at an aircraft speed of 600

m.p.h. the t.h.p. would be

However, if the same thrust was being produced by a turbo-propeller engine with a propeller efficiency of 55 percent at the same flight speed of 600 m.p.h., then the t.h.p. would be:

Thus at 600 m.p.h. one lb. of thrust is the equivalent of about 3 t.h.p.

ENGINE THRUST IN FLIGHTSince reference will be made to gross thrust, momentum drag and net thrust, it will be helpful to define these terms:

Gross or total thrust is the product of the mass of air passing through the engine and the jet velocity at the propelling nozzle, expressed as:


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)A +

The momentum drag is the drag due to the momentum of the air passing into the

engine relative to the aircraft velocity, expressed as where:

W = Mass flow in lb. per sec.V = Velocity of aircraft in feet per sec.G = Gravitational constant 32.2 ft. per sec. per sec.

The net thrust or resultant force acting on the aircraft in flight is the difference between the gross thrust and the momentum drag.From the definitions and formulae stated earlier under flight conditions, the net thrust of the engine, simplifying, can be expressed as:

All pressures are total pressures except P, which is static pressure at the propelling nozzle

W = Mass of air passing through engine (lb. Per sec.)

VJ = Jet velocity at propelling nozzle (ft. per sec)

P = Static pressure across propelling nozzle (lb. Per sq. in)

PO = Atmospheric pressure (lb. Per sq. in)

A = Propelling nozzle area (sq. in)

V = Aircraft speed (ft. per sec.)

G = Gravitational constant 32.2

The balance of forces and expression for thrust and momentum drag

Effect of forward speedSince reference will be made to ‘ram ratio’ and Mach number, these terms are defined as follows:Ram ratio is the ratio of the total air pressure at the engine compressor entry to the static air pressure at the air intake entry.


Mach number is an additional means of measuring speed and is defined as the ratio of the speed of a body to the local speed of sound. Mach 1.0 therefore represents a speed equal to the local speed of sound.From the thrust equation, it is apparent that if the jet velocity remains constant, independent of aircraft speed, then as the aircraft speed increases the thrust would decrease in direct proportion. However, due to the ‘ram ratio’ effect from the aircraft forward speed, extra air is taken into the engine so that the mass airflow and also the jet velocity increase with aircraft speed. The effect of this tends to offset the extra intake momentum drag due to the forward speed so that the resultant loss of net thrust is partially recovered as the aircraft speed increases. A typical curve illustrating this point is shown in the figure.


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Obviously, the ‘ram ratio’ effect, or the return obtained in terms of pressure rise at entry to the compressor in exchange for the unavoidable intake drag, is of considerable importance to the turbo-jet engine, especially at high speeds. Above speeds of Mach 1.0, as a result of the formation of shock waves at the air intake, this rate of pressure rise will rapidly decrease unless a suitably designed air intake is provided; an efficient air intake is necessary to obtain maximum benefit from the ram ratio effect.


As aircraft speeds increase into the supersonic region, the ram air temperature rises rapidly consistent with the basic gas laws. This temperature rise affects the compressor delivery air temperature proportionally and, in consequence, to maintain the required thrust, the engine must be subjected to higher turbine entry temperatures. Since the maximum permissible turbine entry temperature is determined by the temperature limitations of the turbine assembly, the choice of turbine materials and the design of blades and stators to permit cooling are very important.

With an increase in forward speed, the increased mass airflow due to the ‘ram ratio’ effect must be matched by the fuel flow and the result is an increase in fuel consumption. Because the net thrust tends to decrease with forward speed, the end result is an increase in specific fuel consumption (s.f.c.), as shown by the curves for a typical turbo-jet engine in the figure.At high forward speeds at low altitudes, the ‘ram ratio’ effect causes very high stresses on the engine and, to prevent over-stressing, the fuel flow is automatically reduced to limit the engine speed and airflow.

Effect of afterburning on engine thrust

At take-off conditions, the momentum drag of the airflow through the engine is negligible, so that the gross thrust can be considered to be equal to the net thrust. If after-burning is selected, an increase in take-off thrust in the order of 30 percent is possible with the pure jet engine and considerably more with the by-pass engine. This augmentation of basic thrust is of greater advantage for certain specific operating requirements. Under flight conditions, however, this advantage is even greater, since the momentum drag is the same with or without after-burning and, due to the ram effect, better utilisation is made of every pound of air flowing through the engine.

Effect of altitude

With increasing altitude the ambient air pressure and temperature are reduced. This affects the engine in two inter-related ways:-The fall of pressure reduces the air density and hence the mass airflow into the engine for a given engine speed. This causes the thrust or s.h.p. to fall. The fuel control system adjusts the fuel pump output to match the reduced mass airflow, so maintaining a constant engine speed.The fall in air temperature increases the density of the air, so that the mass of air entering the compressor for a given engine speed is greater. This causes the mass airflow to reduce at a lower rate and so compensates to some extent for the loss of thrust due to the fall in atmospheric pressure. At altitudes above 36,089 feet and up to 65,617 feet, however, the temperature remains constant, and the thrust or s.h.p. is affected by pressure only.Graphs showing the typical effect of altitude on thrust, s.h.p. and fuel consumption are illustrated.


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Effect of temperature

On a cold day the density of the air increases so that the mass of air entering the compressor for a given engine speed is greater, hence the thrust or s.h.p. is higher. The denser air does, however, increase the power required to drive the compressor or compressors; thus the engine will require more fuel to maintain the same engine speed or will run at a reduced engine speed if no increase in fuel is available. On a hot day the density of the air decreases, thus reducing the mass of air entering the compressor and, consequently, the thrust of the engine for a given r.p.m. Because less power will be required to drive the compressor, the fuel control system reduces the fuel flow to maintain a constant engine rotational speed or turbine entry temperature, as appropriate; however, because of the decrease in air density, the thrust will be lower. At a temperature of 45C, depending on the type of engine, a thrust loss of up to 20 percent may be experienced. This means that some sort of thrust augmentation, such as water injection, may be required.The fuel control system, controls the fuel flow so that the maximum fuel supply is held practically constant at low air temperature conditions, whereupon the engine speed falls but, because of the increased mass airflow as a result of the increase in air density, the thrust remains the same. For example, the combined acceleration and speed control fuel system schedules fuel flow to maintain a constant engine r.p.m., hence thrust increases as air temperature decreases until, at a predetermined compressor delivery pressure, the fuel flow is automatically controlled to maintain a constant compressor delivery pressure and, therefore, thrust, illustrates this for a twin-spool engine where the controlled engine r.p.m. is high-pressure compressor speed and the compressor delivery pressure is expressed as P3. It will also be apparent from this graph that the low pressure compressor speed is always less than its limiting maximum and that the difference in the two speeds is reduced by a decrease in ambient air temperature. To prevent the LP compressor overspeeding, fuel flow is also controlled by an LP governor which, in this case, takes a passive role.



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PROPULSIVE EFFICIENCY

Performance of the jet engine is not only concerned with the thrust produced, but also with the efficient conversion of the heat energy of the fuel into kinetic energy, as represented by the jet velocity, and the best use of this velocity to propel the aircraft forward, i.e. the efficiency of the propulsive system.The efficiency of conversion of fuel energy to kinetic energy is termed thermal or internal efficiency and, like all heat engines, is controlled by the cycle pressure ratio and combustion temperature. Unfortunately this temperature is limited by the thermal and mechanical stresses that can be tolerated by the turbine. The development of new materials and techniques to minimise these limitations is continually being pursued. The efficiency of conversion of kinetic energy to propulsive work is termed the propulsive or external efficiency and this is affected by the amount of kinetic energy wasted by the propelling mechanism. Waste energy dissipated in the jet wake, which represents a loss, can be expressed as

W(vJ - V) 2 where (vJ - V)

2g

is the waste velocity. It is therefore apparent that at the aircraft lower speed range the pure jet stream wastes considerably more energy than a propeller system and consequently is less efficient over this range. However, this factor changes as aircraft speed increases, because although the jet stream continues to issue at a high velocity from the engine, its velocity relative to the surrounding atmosphere is reduced and, in consequence, the waste energy loss is reduced.


FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP

Primary engine design considerations, particularly for commercial transport duty, are those of low specific fuel consumption and weight. Considerable improvement has been achieved by use of the by-pass principle and by advanced mechanical and aerodynamic features and the use of improved materials. With the trend towards higher by-pass ratios, in the range of 15:1, the triple-spool and contra-rotating rear fan engines allow the pressure and by-pass ratios to be achieved with short rotors, using fewer compressor stages, resulting in a lighter and more compact engine.S.f.c. is directly related to the thermal and propulsive efficiencies; that is, the overall efficiency of the engine. Theoretically, high thermal efficiency requires high pressures which in practice also means high turbine entry temperatures. In a pure turbo-jet engine this high temperature would result in a high jet velocity and consequently lower the propulsive efficiency. However, by using the by-pass principle, high thermal and propulsive efficiencies can be effectively combined by by-passing a proportion of the LP compressor or fan delivery air to lower the mean jet temperature and velocity. With advanced technology engines of high by-pass and overall pressure ratios, a further pronounced improvement in s.f.c. is obtained. The turbines of pure jet engines are heavy because they deal with the total airflow, whereas the turbines of by-pass engines deal only with part of the flow; thus the H.P. compressor, combustion chambers and turbines, can be scaled down. The increased power per lb. of air at the turbines, to take advantage of their full capacity, is obtained by the increase in pressure ratio and turbine entry temperature. It is clear that the by-pass engine is lighter, because not only has the diameter of the high pressure rotating assemblies been reduced, but also the engine is shorter for a given power output. With a low by-pass ratio engine, the weight reduction compared with a pure jet engine is in the order of 20 per cent for the same air mass flow.With a high by-pass ratio engine of the triple-spool configuration, a further significant improvement in specific weight is obtained. This is derived mainly from advanced mechanical and aerodynamic design, which in addition to permitting a significant reduction in the total number of parts, enables rotating assemblies to be more effectively matched and to work closer to optimum conditions, thus minimising the number of compressor and turbine stages for a given duty. The use of higher strength lightweight materials is also a contributory factor.For a given mass flow, less thrust is produced by the by-pass engine due to the lower exit velocity. Thus, to obtain the same thrust, the by-pass engine must be scaled to pass a larger total mass airflow than the pure turbo-jet engine. The weight of the engine, however, is still less because of the reduced size of the H.P. section of the engine. Therefore, in addition to the reduced specific fuel consumption, an improvement in the power-to-weight ratio is obtained.

SPECIFIC FUEL CONSUMPTION

When comparing engine performance, one of the most important considerations is how efficiently the power is produced. The amount of fuel consumed to produce a given horsepower lbs. thrust is known as “specific fuel consumption” or SFC. A typical aircraft fuel system measures the volume of fuel consumed. This is displayed in pounds per hour or PPH. To calculate fuel flow, specific fuel consumption found on the customer data sheet, is multiplied by the horsepower lbs. thrust produced.


27

SPECIFIC FUEL CONSUMPTION – DEFINITION

SFC = SPECIFIC FUEL CONSUMPTION defined as LBS (fuel) per HP/lbs. thrust per hour

FLAT RATING

“Flat rating” is used by aircraft manufacturers when they select an engine that has a capability greater than the requirements of the aircraft. They then limit the power output of the engine. There are three distinct benefits derived from flat rating. One is the engine will have the ability to make take-off power at lower turbine temperatures over a wide range of outside air temperatures and pressure altitudes. Performance at altitude will be greatly enhanced. These two benefits result in the third benefit, longer engine life.

PERFORMANCE RATINGS

In the chart, performance ratings are compared on –1 through –12 engines. Notice the modifiers on the –1, -5, -6, -8 and –10 engines. These temperatures represent the effects of flat rating engines. Each engine will make take-off power below their turbine temperature limits to the ambient temperatures indicated. Engines that are not flat rated, such as the –3 or –11, would be unable to make take-off power below their turbine temperature limits when operating in conditions above 59F outside air temperatures.

PERFORMANCE RATINGS (example from a turbo-prop).


INTENTIONALLY LAFT BLANK


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15.3 AIR INTAKES

COMPRESSOR AIR INLET DUCTS

The main requirement of an intake is that, under all operating conditions, delivery of the air to the engines is achieved with the minimum loss of energy occurring through the inlet duct. To enable the compressor to operate satisfactorily, the air must reach the compressor at a uniform speed/pressure(maximum 0.5 mach) distributed evenly across the whole inlet area.

Although there are no exact figures, it is normally taken that the speed range between Mach 0.8 and Mach 1.2 is termed the transonic range whilst that between Mach 1.2 and Mach 2.5 is the supersonic range

The ideal air intake for a turbo-jet engine fitted to an aircraft flying at sub-sonic or low supersonic speeds, is a short, pitot-type circular intake. This type of intake makes the fullest use of ‘ram-effect’ on the air due to the forward speed, and suffers the minimum loss of ram pressure with changes of aircraft altitude. It will deliver air to the compressor inlet face at approximately Mach 0.5.

The pitot-type intake can be used for engines that are mounted in pods either attached to wing pylons or fuselage stub wings. Some installations require that the intakes deviate from the ideal circular form to meet structural and aerodynamic requirements.

At sonic speeds, the efficiency of this type of intake begins to fall, because of the formation of a shock wave at the intake lip. This is when the characteristics of the airflow changes and the formation and control of shock waves has to be considered. Supersonic intake design will be covered later in this chapter.

Each inlet configuration has been designed to achieve the same result, the delivery of the maximum mass of air to the front compressor face. This is sometimes known as ‘ram recovery’ or ‘total pressure recovery’.


PITOT TYPE INTAKES

Intakes in their simplest form, the circular pitot intake, allow the air to enter the engine in a straight line and sub-sonic. It can be seen in the illustration overleaf that the intake duct is a diverging duct, and hence, it both slows the airflow into the engine, (allowing flight to high sub-sonic speeds without choking the engine), and forces the slowing air to increase in pressure, again increasing the overall efficiency.

Pitot Type IntakeFigure 3.1


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DIVIDED INTAKE DUCTS

Some single engine aircraft have a pitot type of intake, but due to the long intake duct and the loss of space in the front of the aircraft, many more have a divided type of intake on each side of the fuselage. The disadvantage of this type of intake is that during any yawing manoeuvre, a loss of ram pressure occurs on one side of the intake, causing an uneven distribution of airflow into the compressor. The Vantage business jet shown below has divided intakes, located at high mid-fuselage, which could suffer from this effect.

Divided Air IntakeFigure 3.2


CURVED INTAKE DUCTS

There are a few aircraft, notably three-engine types, which have the centre (number 2) engine, fitted with a complex, “S shaped” curved intake duct. This duct has quite high losses compared with a simple pitot type of intake. The designer accepts these losses against the advantage of having the engine within the fuselage shape with little extra drag. In one notable case, the McDonnell Douglas DC-10 / MD-11 series, the complete engine assembly is located about a quarter of the way up the vertical stabiliser

The three-engine layout mentioned is also used in a selection of smaller business-jet aircraft, as well as the wide-body aircraft. The intake duct for the centre engine of this type of aircraft is in a form of a large ‘S’ shape and this has two penalties. Firstly the air has inertia, meaning it has to be forced to follow the curves of the duct, resulting in losses not suffered by the number 1 and 3 engines. Secondly, the air cannot be made to interface accurately with the front face of the compressor, resulting in uneven airflow. The illustration of a Dassault Falcon business jet below shows the complex, centre (No.2) engine intake system, compared with the Nos. 1 & 3 engines.

Falcon Jet Centre Intake LayoutFigure 3.3


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SUPERSONIC PITOT TYPE INTAKE

At higher supersonic speeds, the pitot type of air intake is unsuitable due to the severity of the shockwave that forms and progressively reduces the intake efficiency as speed increases. A more suitable intake for these higher speeds is known as the external/internal compression intake. This type of intake produces a series of mild shock waves without excessively reducing the intake efficiency.

Supersonic Pitot Type Air IntakeFigure 3.4


SUPERSONIC AIR INTAKE WITH RAMPS

As aircraft speeds increase still further the intake compression ratio also increases. At high Mach numbers, say over M=1.5, it is necessary to have an air intake that has a variable throat area and spill valves fitted to accommodate and control the changing mass of air.

The airflow velocities encountered in the higher speed range of the aircraft are much higher than the engine can efficiently use. The air velocity must therefore, be decreased between the intake and the engine air inlet. The angle of a variable throat area intake automatically varies with aircraft speed. It positions the shock wave to decrease the air velocity at the engine inlet and maintains maximum pressure recovery within the inlet duct.

The basic principle of these intakes is that the first part of the intake, being a converging duct, slows the supersonic air to Mach 1. Beyond this point the duct diverges, slowing the sub-sonic air to about Mach 0.5 by the time it reaches the face of the compressor. The cross section of a supersonic intake shows the layout, especially the moveable ramps.

Supersonic Intake with RampsFigure 3.5

Continued development enables the same effect to be achieved by careful design of the intake and ducting. This, coupled with auxiliary air doors to permit extra air to be taken in under certain engine operating conditions, allows the airflow to be controlled without the use of variable geometry intakes.


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TURBO-PROPELLER AIR INTAKES

The intake ducts used on turbo-propeller engine installations can have several different configurations. The main reason for this is that some turbo-propeller designs have their output shafts running through the centre of the intake. Others, however, have a reduction gearbox installed on the front of the engine, which leaves the output shaft offset to one side and space for a clear, pitot type of intake. Examples of both of these types are illustrated below.

Turbo-Propeller Air IntakesFigure 3.6


BELLMOUTH AIR INTAKE

Finally, one form of intake that may be seen on slow moving aircraft, especially helicopters, is the bellmouth type of intake. These intakes are converging in shape and are fitted on to aircraft that fly below ram- recovery speed. This type of inlet produces a large amount of drag, but this disadvantage is overcome by their high degree of aerodynamic efficiency. The illustration below shows a typical bellmouth intake fitted to a Rolls Royce Gnome 1200.

Bellmouth Air IntakeFigure 3.7

This type of intake is also fitted to the front of gas turbine engines under test. Because this type of intake has very little duct losses, the performance figures taken under test usually assume zero duct losses.


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AIR INTAKE ICE PROTECTION

Icing of the engine and the leading edge of the intake duct can occur during flight through super-cooled clouds or during ground operations in freezing fog. Protection against ice formation may be required since icing of these regions can considerably restrict the airflow through the engine, causing a loss in performance and possible malfunction of the engine itself. Additionally, damage may be caused by ice breaking away and being ingested into the engine or striking the acoustic material that lines the intake duct.

To prevent ice forming upon the vital parts of the engine and its nacelle, a system must be developed that will apply heat to the intake lip and essential parts of the compressor front face. The system must be reliable, easy to maintain, present no excessive weight penalty and cause no serious loss in engine performance when in operation. The parts of the engine installation that may require protection are:

Intake Lip Centre nose cone Sensor probes Guide vanes Intake struts

Not all of the previously mentioned parts of all installations will require specific protection. For example, the nose cone of the ALF 502 engine has hot oil circulating inside it, whereas the nose cone of the Rolls-Royce Tay has a special coating that resists the build-up of ice and has no heating.

As a general rule, turbo-jet engines, which have a ready source of hot, bleed air, use the hot air systems, whist the turbo-propeller engines normally use electrical power, (although bleed air may be used for some specific tasks).

Although there are exceptions, in general:

The hot air systems are generally used to prevent ice forming and are known as anti-icing systems.

The electrical power systems are used to break up ice that has formed on the surfaces and are known as de-icing systems.


THE HOT AIR SYSTEM

The hot air systems provide surface heating of the engine and powerplant, (and the airframe), where ice is likely to form. Rotor blades rarely require protection due to the high centrifugal forces present.

The hot air for the anti-icing system is usually taken from the high-pressure compressor stages. It is ducted through pressure regulator and shut-off valvesto the parts of the engine requiring protection. Once used, the spent air can be ducted into the intake duct or, on some designs, overboard. The illustration shows a typical hot-air anti-icing system protecting the intake lip, nose cone and intake guide vanes.

Hot Air System Figure 3.8


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ELECTRICAL SYSTEM

The electrical system of ice protection is generally used on turbo-propeller engine installations. One reason for this system being used is the additional protection required on the propellers. As a general statement, protection on turbo-propeller installations is applied to the intake cowling, the propeller blades and spinner. It may also be required to protect the intake of the oil cooler.

Electrical heating pads are bonded to the outer skin of the cowlings. They consist of strip conductors sandwiched between either layers of neoprene or glass cloth impregnated with epoxy resin. Due to the eroding effects of rain and hail, protection is required on the leading edges and consists of special polyurethane-based paint.

When in operation, some parts of the pads are heated continuously and some are heated cyclically. The continuously heated parts keep the actual leading edge free of ice whilst the cyclic heating breaks the formed ice off, using the assistance of the airflow. The illustration below shows where the elements might be located on an intake.

Electrical Heating ElementsFigure 3.9


CYCLIC TIMING

Whilst the electrical loads are provided by the aircraft electrical generating systems, to prevent excessive loads being placed upon the generators, (or the generators needing to be too large and heavy), the loading is cycled between the engine intake and the propeller blades & spinner. A typical cycle example is shown below.

ELECTRICAL HEATING CYCLEFigure 3.10

The cyclic timing of the intermittently heated elements is arranged, firstly, to ensure that the engine can accept the amount of ice that collects during the ‘heat-off’ period. Secondly, it ensures that the ‘heat-on’ period is long enough to give adequate ice shedding. This ensures a clean, ice free, surface, without “run back” forming behind the protected surfaces.

Some systems cater for the fact that ice build-up can be at different rates depending on circumstances. To cater for this, there are sometimes ‘fast’ and ‘slow’ cycles, selectable from the cockpit, and used depending on the outside air temperature (OAT) and any precipitation in the air.


41

MIXED SYSTEMS

On some aircraft, because of limitations at the design stage, the engine/powerplant assembly is either anti-iced or de-iced by a mixture of hot air, oil and electrical services. The example below is shown as using all three systems, although this design introduces complexity that could cause difficulty in service over time. Some engines may be found using any of the above three systems, to prevent ice accretion.

Mixed Ice Removal SystemFigure 3.11




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15.4 COMPRESSORS

The compressor section of the turbo-jet engine has many functions. Its primary purpose is to supply air in sufficient quantity to satisfy the requirements of the combustion process. Specifically to fulfil its purpose, the compressor must increase the pressure of the mass of air received from the air inlet duct and then to discharge it to the burners in the combustion chambers, in the mass and pressures required.

Compressors may be identified by the direction of the airflow through them. The two basic types have a centrifugal flow or an axial flow. Some engines may use both types on one compressor assembly.

CENTRIFUGAL COMPRESSORS

These compressors receive the air at their centre and accelerate it outwards by centrifugal force. The air is then expelled into a divergent duct, called a diffuser, where velocity is exchanged for energy.

A complete centrifugal compressor assembly consists of an impeller rotor, a diffuser and a manifold. The impeller can be single or double sided and can be installed in either one or two-stage assemblies. Whilst this type of compressor can generate a high mass-flow from a small diameter engine, it cannot take advantage of ram effect due to the tortuous route that the airflow has to follow through the compressor.

Whilst compression ratios in the vicinity of 5:1 were the norm on earlier designs, it is now possible to produce centrifugal compressors with compression ratios of 15:1, which are quite competitive with axial compressors. It is very rare to find more than two stages of compression, due to the huge losses caused by the continued re-direction of the airflow through the stages, the added weight of the impellers and the power required from the turbine to drive the compression stages.


The illustrations below show a single stage, dual-sided compressor and a two-stage, single sided centrifugal compressor.

Single Stage Dual Sided Compressor Two Single Sided Compressors

Figure 4.1


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AXIAL FLOW COMPRESSORS

The axial flow compressor has two main components, a rotor and stator. The rotor has blades attached to a spindle or drum, which impels the air rearwards in the same manner as a propeller. The stator blades act as diffusers at each stage, partially converting high velocity to pressure.

A set of rotor and stator blades constitutes a pressure stage, each stage being capableof producing a pressure rise of about 1.25:1. The number of these stages is dictated by the amount of air and the pressure rise that is required. A normal maximum number of stages to be found is between 16 and 18.From the front to the rear of an axial compressor, the space between the rotor shaft and the stator casing becomes smaller. This is necessary to maintain a near constant axial velocity of the air as the density increases with compression. This is shown in the illustration below.

Axial compressorFigure 4.2


ADVANTAGES AND DISADVANTAGES

ADVANTAGES

CENTRIFUGAL COMPRESSOR AXIAL COMPRESSORHigh pressure rise per stage High ram effect efficiencyGood efficiency over wide speed range High peak pressuresSimplicity of manufacture, low cost Small frontal areaLow weightLow power for startingDamage tolerant

DISADVANTAGES

CENTRIFUGAL COMPRESSOR AXIAL COMPRESSORLarge frontal area Complex manufactureLimited to two stages Relative high weight

High starting powerLow pressure rise/stage


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COMBINED COMPRESSORS

There are a number of gas turbines that use both types of compressor by having an axial compressor, followed immediately by a centrifugal compressor. The aerodynamic advantages of this arrangement are too complex to discuss at this stage but, suffice to say that this layout can generally be found on turbo-shaft engines, which power helicopters. The example, illustrated below, is of the Lycoming T-55 engine that powers the Chinook helicopter. It can be seen that it has seven axial stages followed by a single centrifugal stage.

Combined Axial/Centrifugal CompressorsFigure 4.3


OPERATING PRINCIPLES

CENTRIFUGAL

The impeller is rotated at high speed by the turbine and the air is continuously induced into the centre of the impeller. Centrifugal action causes it to flow outwards along the vanes to the tip, which causes it to accelerate and the pressure to rise.

Once it leaves the impeller it passes through the diffuser section, which is divergent, causing the pressure to rise again. This demonstrates how this arrangement has half the compression occurring in the impeller and half in the diffuser.

This type of compressor works best at high rotational speeds. It is normal for a centrifugal compressor to have impeller tip speeds of around 1,600 ft per second, (well over Mach1). This is one of the reasons why centrifugal compressors generate a high level of noise when operating.

AXIAL

The rotor is rotated at high speed by the turbine, continuously drawing air into the front of the compressor. After each rotor stage, which has caused the pressure to rise, the air passes through a stator stage, which diffuses, (decelerates), the air and causes the pressure to rise yet again. This process continues throughout the number of stages of the compressor, each stage comprising a rotor and a stator, each stage achieving a compression ratio of approximately 1.25:1

The stators have a second duty, which is to straighten out the ‘swirl’ which is the result of axial compression. As the air leaves each rotor stage with increased velocity, it also has a rotary motion that, if not corrected, will reduce the efficiency of each progressive stage. The stator turns the air in the reverse direction, resulting in the airflow flowing axially through each stage.

When axial compressors are required to produce a high level of compression, it becomes very difficult to control the air throughout all of the stages. This is due to the variables that any aircraft can meet. These include the speed of the compressor, due to throttle demand from the flight deck; the speed of the aircraft, especially in the climb or descent and the density of the air or altitude at which the aircraft is operating.


49

CONSTRUCTIONAL FEATURES

CENTRIFUGAL

The centrifugal compressors are usually mounted on ball or roller bearings and are driven by the turbine stage(s). The connecting shaft may be manufactured in two parts, to allow engine disassembly, whilst having a self-aligning coupling to join the parts together.The discs are forged with the vanes straight for ease of manufacture. Normally a separate set of rotating guide vanes, which cannot be easily forged, are attached to the front of the impeller. These draw the air into the impeller unit.

Diffuser assemblies are often part of the compressor case, with integrally cast vanes to act as both diverging ducts and to direct the airflow into the elbows and the combustion chambers.

AXIAL

The Axial compressor consists of firstly, the rotating rotor, made up from the main shaft supported by ball and roller bearings and either separate discs or a drum assembly, to which are affixed the blades of differing sizes. Secondly, the casing assembly, in a number of pieces (to allow splitting, for access to the rotors), contains all of the stator vanes attached to the inside face of the case. The case also provides part of the strength of the complete engine and, on some designs, has attachments or mounting points built into the case design.

The vanes are affixed to the rotor discs and stator case(s) by a variety of methods, all giving positive retention against centrifugal force, (rotors) and rotation, (stators).The rotor blades are of aerofoil section and are twisted, much the same as a propeller, to give an even thrust along their length. This is shown by the different stagger angles between the root and tip of the blades. The roots of the blades are formed into a shape that matches the recesses in the rotor disc and they only have to be retained on the disc by plates that restrict fore and aft movement. This can be seen in the illustration overleaf.


Blade Details Figure 4.4

1. STATOR VANES

Stator vanes are also of aerofoil section and are located in slots around the compressor casing. There is no chance for the blades to move fore and aft due to the retention of the grooves, but there is a tendency for the blades to slide radially around the grooves. This tendency is caused by the air loads, generated by the blades straightening the airflow after each rotor stage.

This movement is prevented by retaining set screws, which hold a number of blades in place, preventing any movement by the others. This is shown in Figure 4.5, where a retaining ring, held by the screw, holds the blades in place.

Stator Blade RetentionFigure 4.5


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FAN BALANCING

The fan consists of the single front stage of the compressor. Normally, it is the low- pressure (LP) compressor and is part of a twin or triple-spool engine. It will usually consist of a small number of blades that can be removed, often individually, if they become damaged in service. The engine shown below, a Rolls Royce Tay, has a wide chord fan which can be both repaired, (by blade replacement) and balanced in situ.

Fan blades may be manufactured from Titanium, sometimes as a skin with a honeycomb core, although some have been manufactured from composite materials. Titanium is used normally because of the bird strike requirements that dictate very strong blades on the first stage of the engine.

Rolls Royce Tay FanFigure 4.6


The need for the engine to be precisely balanced because of its high rotational speed, means that the replacement of individual blades must be undertaken with care. In most cases, the blades will have been pre-weighed by the manufacturer and the value engraved upon the blade. The blades will be divided into weight “groups” so that, providing the replacement blade is of the same “group” as the one removed, there should be no need to balance the assembly.

In some cases, due to the engine having built-in vibration sensors, it will be possible to carry out balancing ground-runs. This will allow the engineer, following the maintenance manual, to add or remove small balance weights, at specific points around the fan assembly, until the assembly is in perfect balance.

Shown below is the fan assembly of the Tay engine, showing the use of weights to give a balanced assembly.

Tay Fan AssemblyFigure 4.7


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STALL AND SURGE

Each stage of a multi-stage compressor possesses certain airflow characteristics that are dissimilar from those of its neighbour thus; to design a workable and efficient compressor, each stage must be matched to the next stage. This matching is fairly simple when the engine is running on a test bed, it is much more difficult when speed, altitude, temperature, etc. are included, such as when the aircraft is operating normally.

In extreme conditions, the airflow through the compressor can become disturbed and vibration can be set-up. This stalling of the blades can either be positive or negative, depending whether the fault is at the intake, (front), or at the high compression, (rear), end of the engine.

If the engine demands a pressure rise from the compressor greater than the blades can sustain, surge will occur. This is an instantaneous breakdown of flow through the engine and high-pressure air in the combustion system is expelled forwards through the compressor,T.G.T. would rise and may be accompanied with a loud bang, resulting in a loss of thrust. To overcome this problem, engines have a declared ‘safety margin’ to ensure the area of instability is avoided. This is shown graphically below.

Surge Margin DiagramFigure 4.8


To control these disturbances, which occur most often on single shaft engines with high compression ratios, a variety of methods are used on different engines. This control can take the form of variable inlet guide vanes for the first stage and variable stator vanes for other stages. As the compressor slows from its optimum, the blades change their angle of attack to vary the airflow on to the rotor blades, so that they do not stall and remain at their optimum angle of attack.

BLEED BANDS/VALVES

In addition, an interstage bleed may be fitted to the compressor casing, usually located at the higher compression stages, permitting excessive pressure to be bled overboard. This avoids the choking which may occur during rapid acceleration. Due to the loss of performance during normal operations, bleed valves will usually only be opened during starting and acceleration. The operation of these air bleed systems can either be actuated by hydraulic, pneumatic or electronic methods

VARIABLE INLET GUIDE VANES (VIGVS)

The number of stages that have variable incidence vanes depends on the design of the engine. Some may only have the first stage inlet guide vanes moveable, whilst others can have four or more stages that are variable. The illustration below shows an engine with variable inlet guide vanes and three variable stator vanes.

Guide Vanes and Stator Vanes Figure 4.9


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OVERFUELLING SURGE

All engines have to be over fuelled by a small margin to cause them to accelerate. If the over fuelling is above the correct figure, due perhaps to a badly adjusted fuel control unit, then the inertia of the rotating parts of the engine will resist acceleration.

The excessive fuel will cause choking at the turbine; this will cause a slowing of the compressor air velocity, resulting in a progressive stall through the engine from the front. The resulting reversal of the airflow is a surge.

TWIN SPOOL AXIAL FLOW COMPRESSORS

Relief from surging troubles can be obtained from the devices described earlier. A better solution is the twin-spool axial flow compressor, part of the twin spool engine type described earlier. The compressor has two sections, each section is completely independent from the other and driven by its own turbine assembly, each mounted on its own co-axial shaft. The LP compressor is driven by the aft, LP turbine and the HP compressor is driven by the forward, HP turbine. Each shaft assembly will be rotating at its optimum speed.

Whether the engine is at high or low altitude or whether it is moving through the air at high or low speeds, the two spools will be matched to the external atmosphere parameters and aircraft performance.

At idle, for example, the HP system is doing most of the work whilst the LP spool runs slower, this makes its angle of attack of the airflow on to the first stage much better and, due to the faster moving HP spool, there is less chance of ‘choking’. Equally, at higher altitudes, when the LP spool rotates faster, due to the reduced air density, the greater mass airflow to the HP section restores some of the losses that a single spool engine would suffer at this altitude.


INTENTIALLY LEFT BLANK


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15.5 COMBUSTION SECTION

The combustion section houses the process by which the energy contained within the air/fuel mixture is released. The primary function of the combustion section is to burn the air/fuel mixture adding heat energy to the air. To carry this out efficiently it must:

Provide the means for proper mixing of the air and fuel to assure good combustion.Burn this mixture efficiently.Cool the hot combustion products to a temperature that the turbine blades can withstand under operating conditions.Deliver the hot gases to the turbine section.Combustion sections are located between the compressor-diffuser and the turbine section. They are usually located co-axially with the compressor and the turbine.All combustion chambers contain the following elements:

An outer casing A perforated inner liner A fuel injection system (This topic will be covered later) Means of ignition (This topic will also be covered later) A fuel drainage system


COMBUSTION CHAMBER AIRFLOW

To ensure that the components of the combustion chamber and the turbine assembly are not overheated, the airpaths are divided into primary and secondary. Around 20% of the total air is fed to the fuel nozzles for combustion. The remaining, secondary airflow forms a cooling air blanket around the liner, centres the flame and, finally, mixes with the primary airflow, so that the total airflow to the turbine is at an acceptable temperature. The illustration below shows how the total compressor output is divided into sections for different purposes.

Apportioning the AirflowFigure 5.1

An additional purpose of the combustion system, which has become more and more important recently, is total and complete combustion. The very low air pollution levels required before certification can be granted, will not allow the smoke trails that were typical of older generation aircraft.

There are currently three basic types of combustion chambers, with variations within these types being in detail only:

The Multiple Combustion Chamber The Tubo-Annular Combustion Chamber The Annular Combustion Chamber


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MULTIPLE COMBUSTION CHAMBERS

The multiple chamber layout is found on centrifugal compressor engines, such as the Roll-Royce Dart turbo-propeller engine and also early axial compressor designs. The chambers are disposed around the engine and the compressor delivery air is directed, by ducts, into each individual chamber. Whilst the airflow enters the chamber at high velocity, swirl vanes and baffles slow the combustion air to a speed at which the flame can safely exist.

If any problem exists with a single chamber, such as overheating and bulging of the outer case, it can be changed easily with the engine remaining in situ.

Each chamber has an inner flame tube surrounded by an air casing. All of the flame tubes are interconnected to both equalise the air pressure throughout the combustion system and to allow the combustion to propagate around all of the flame tubes during engine starting. Fuel manifolds carry fuel to the burners and a complex drain system ensures that all the chambers are emptied to a collector tank, after shut-down or a “wet start”. All these features can be seen in the illustrations below.

Combustion chamberFigure 5.2


Multiple Combustion ChambersFigure 5.3


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TUBO - ANNULAR/CAN ANNULAR

The Tubo-annular (Can-Annular) combustion chamber is a mixture between the multiple combustion chamber and the annular combustion chamber layout. A number of flame tubes are installed inside a common air casing. The airflow through this chamber is similar to the multiple chamber system described earlier.

Pratt & Whitney and Rolls Royce produced many designs based on this principle. This arrangement combines the ease of overhaul and testing of the multiple system with the compactness of the annular system. As with the previous designs, there are normally two igniters, in opposing chambers to initiate ignition at engine start. A cutaway illustration of a typical Tubo-annular combustion is shown below.

Tubo – Annular/Can AnnularFigure 5.4


ANNULAR

Finally there is the Annular Combustion Chamber which consists of a flame tube (commonly referred to as a combustion liner) circular in construction which is contained within an inner and outer casing. This design is in common use today in both small and large engines. They are the most efficient design, both from a standpoint of thermal efficiency and weight.

They are also shorter than comparable engines. For the same power output, the length of an annular chamber is 75% that of a Tubo-annular system of the same diameter.

The smaller amount of surface area requires less cooling air and is also the most efficient use of space. However, this type of combustor must be removed as a single unit for repair or replacement, requiring a complete separation of the engine at major flanges. (This is a much more complex operation than that required for the single chamber design).

Another advantage of this type of design is the elimination of combustion propagation problems from chamber to chamber.

This type of design can be found in two totally different forms. They can be found with a straight-through flow or with a reverse flow.

STRAIGHT-THROUGH FLOW

The straight-through flow annular combustor takes in air at the front and discharges it at the rear. The annular combustor consists of an outer housing with a perforated inner liner, sometimes called a ‘basket’. Both of these parts encircle the engine.

Multiple fuel burners project into the basket, together with Ignitor(s). Both the primary and secondary airflow’s behave the same way as they do in the other combustor designs. Because of their efficient fuel burning, annular combustion chamber engines are one of the most efficient designs in the world, which also produce the cleanest, least polluting exhaust possible.

REVERSE FLOW

The reverse flow combustor serves the same function as the through flow unit, but it differs by the air, from the compressor, flowing around the chamber and entering from the rear. This results in the combustion gasses flowing in the opposite direction to the normal flow through the engine.

Note: Sir Frank Whittle used this form of combustion chamber on his earliest jet engine designs.


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The reverse flow form of design also allows the turbine wheels to be located inside the combustor, allowing for a shorter overall length, a lighter engine and pre-heating of the compressor discharge air. These advantages offset some of the losses resulting from the reversal of the airflow. Illustrated below are a through flow combustor and a reverse flow combustor.

Straight Through Flow and Reverse Flow CombustorsFigure 5.5


15.6 TURBINE SECTION

The turbine section has the task of providing the power to drive the compressor, the accessories and, in the case of turbo-propeller/Tubo-shaft engines, providing shaft power to the propeller or rotor. It does this by extracting energy from the hot gases released from the combustion system and expanding them to a lower pressure and temperature.

Very high stresses are involved in this process and, for efficient operation; the turbine blade tips may have a velocity of over 1,500 feet per second. The continuous flow of gas to which the turbine is subjected may have an entry temperature of between 8500C and 1,7000C and may reach a flow velocity of more than 2,500 feet per second in parts of the turbine.

To provide the driving torque, the turbine may consist of several stages, each employing one row of stationary nozzle guide vanes and one row of rotating blades. The number of stages depends upon the relationship between the power required from the gas flow, the rotational speed at which it must be produced and the diameter of the turbine permitted.

The number of main shafts, and therefore the number of turbine stages, depends upon the type of engine. High compression ratio engines will often be twin-spool engines with low-pressure and high-pressure sections, containing their own compressor and turbine stages. On high by-pass ratio engines it is normal for there to be three shafts with again, their own compressor and turbine assemblies.

There are numerous design compromises in the turbine section of an engine. These will involve such variables as mean blade speed, centrifugal stresses, blade thickness/strength and others. The design turbine inlet temperature will dictate its thermal efficiency, the higher the temperature, the more efficient it is. By-pass engines will have a better propulsive efficiency and, thus can have a smaller turbine for a given thrust.


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6.1 TURBINE BLADES

Turbine blades are aerofoil shaped components designed to extract the maximum amount of energy from the flow of hot gasses. The blades may be either forged or cast depending on the alloy from which they are manufactured. Early blades were made from steel forgings whilst one of the current material in use is cast nickel based alloy. The blades today are usually precision-cast and finish-ground to the precise shape.

A modern development is the manufacture of non-metallic blades using a reinforced ceramic material. The ability of these materials to operate satisfactorily at higher temperatures allows the engine designer to plan for higher turbine inlet temperatures and hence greater thrust.

Ceramic coatings applied to nozzle guide vanes also allow them to tolerate higher temperatures, whilst the development of metals which have solidified either with the crystal structure aligned with the main stresses, directional solidification, or as a single crystal, all help to resist ‘creep’. This will be covered later.

Blades are classified as impulse, reaction or a combination impulse-reaction type.

6.1.1 IMPULSE

In the impulse type, the total pressure drop across each stage occurs in the nozzle guide vanes. Because of their convergent shape, they will increase the velocity of the exhaust gasses whilst decreasing their pressure. The gas is then directed on to the blades, which experience an impulse force caused by the impact of the gas on the blades.

6.1.2 IMPULSE/REACTION

Normally, turbine engines do not use pure impulse or pure reaction type blades, but incorporate a design using an impulse-reaction combination. With this combination blade the workload can be evenly distributed along the length of the blade. Also, the axial velocity and the pressure drop across the blade, from root to tip are also considered uniform. A typical impulse-reaction turbine blade is illustrated on the below, and it can be seen how the root end is of impulse design and the outer section is of reaction design.

6.1.3 REACTION

Reaction turbines produce their turning force by an aerodynamic action. The turbine nozzle guide vanes are shaped in such a way that they only aim the gas in the correct direction, not increase its velocity. The gases pass between the blades of the turbine, which do form a converging passage, this increases the velocity of the gases. As the gasses flow over the aerofoil shaped blades. A force reaction in the direction of the plane of rotation causes the turbine to spin.


Turbine Blade SectionsFigure 6.1


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BLADE ATTACHMENTS

The turbine blades are fitted into the turbine wheel (disc) with a form of fastening that allows them to be loose when the engine is cold but to be firmly attached when at their operating temperature. The most common form of attachment is the fir tree root.

These blades are retained within their housings by a variety of methods, the most common being either, peening, welding, lock tabs or riveting.

The turbine blades may be either open or shrouded at their outer ends and either or both types of blade may be used in a single engine. Normally, open-ended blades are used on the high-speed wheels, whilst the shrouded blades are found in wheels having slower rotational speeds.

Shrouded blades form a band around the perimeter of the wheel, which helps to reduce blade vibration. The extra weight of the shrouded tip is offset by the blades being both thinner and more efficient. The illustrations below shows a series of shrouded blades, which are attached to the disc using fir tree roots, (left) and a series of open tip blades, also using fir tree roots.

Shrouded and Open Tip BladesFigure 6.2


6.2 GUIDE VANES

The stator element of the turbine is known by a variety of different names, such as turbine nozzle vanes, nozzle diaphragm, turbine guide vanes and, probably the most common name, nozzle guide vanes. They are located directly aft of the combustion chamber(s) and immediately forward of the first stage turbine wheel. If the engine has more than one turbine stage, there will be a set of stationary nozzle guide vanes ahead of each stage.These nozzles have two functions:

Firstly, after the combustion chamber has introduced the heat energy into the mass airflow and delivered it evenly to the turbine nozzles, it becomes the job of the nozzles to prepare the mass airflow for driving the turbine rotor.

The vanes of the nozzles are set at such an angle that they form a number of small nozzles, discharging the gas at extremely high speed. The nozzles convert a varying

portion of the heat and pressure energy into velocity energy. This energy is converted into mechanical energy through the rotor blades.

Secondly, the nozzles’ purpose is to deflect the gasses to a specific angle in the direction of turbine wheel rotation to ensure that the gasses strike the turbine blades at the optimum angle.

Construction usually consists of an inner shroud and an outer shroud between which are attached the guide vanes. The number and size of the vanes vary with different engine designs, as does the method of attachment and mounting. They all have to make allowance for the expansion that takes place when the engine is operating, which can take the form of loose fitting vanes or expansion slots in the continuous shroud.

The illustration (overleaf) shows two methods of attaching the vanes to the shrouds, loose fitting, (top), and welded, (lower). The loose vanes will become tight within their shrouds when they get to their operating temperature. The welded installations will probably have the inner and/or the outer shroud ring cut into segments, allowing expansion as it heats up in use.


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Nozzle Vane AssembliesFigure 6.3

STRESS AND CREEP

The blades, when at their normal operating temperature, will be glowing red-hot and carrying large centrifugal forces due to their high rotational speeds. A small turbine blade weighing 60gm may exert a load of over 2000kg at maximum engine speed. It must also withstand the high bending loads applied by the gasses to produce the many thousands of turbine (shaft) horsepower necessary to drive the compressor, the accessories and, in some cases, the propeller/rotor assembly also.

The blades are also subject to fatigue, both mechanical and thermal, plus corrosion and erosion. Apart from being manufactured from quite exotic materials, to be proof against the above demands, the blades have to be made by forming and machining using current manufacturing methods.

Following from the above, it can be seen that for a particular blade material and an acceptable safe life there must be an associated maximum permissible turbine entry (inlet) temperature. This maximum T.I.T. also limits the amount of power that the engine can produce. Metallurgists are constantly searching for better materials as well as better blade cooling, to raise the engine T.I.T. and, hence the power.


Over a period of operational time, the turbine blades slowly grow in length. This phenomenon is known as “Creep” and there is a finite, useful life limit before failure of the blade occurs.

The creep, which occurs throughout the life of the blades, can be divided into three phases:The initial, fast, primary creep occurs when the blade is first in service.

Over a much longer period of time, the secondary creep occurs, although

the amount of this creep will be less, overall, than the primary creep.

Finally, towards the end of the blade’s life, the tertiary creep shows an accelerating increase of extension over time, which finishes up at the point of fracture.

Blades in service are not permitted to reach the tertiary creep zone. Most blades will be given a finite life, occurring towards the end of the secondary creep zone. It will however, be seen that overspeeding, frequent temperature and RPM changes, careless handling of the engine caused by mishandling of the engine controls, can quickly erode the safety margins between the retirement life and the point of fracture.

The illustration below left, shows how the blade creeps during service. The chart is NOT drawn to scale, it simply represents the phases the blade goes through. The second chart, on the right, shows how the introduction of better materials has resulted in turbine blades with much better creep characteristics and hence much longer finite lives.

Creep CharacteristicsFigure 6.4


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To reduce the heating effect on the turbine blades, and hence the creep that results, many blades have some form of cooling applied to them. This allows the blades to operate at temperatures above the critical temperatures for the metal alloy used in the construction of the blade.

BLADE COOLING

There are a number of methods used to cool the turbine blades, some of the most popular are listed below:

Internal airflow cooling – Air flows through the hollow blades and vanes exhausting into the gas flow.

Surface film cooling – Air flows from small exit ports in the leading and/or trailing edges of the blades or vanes to form a heat barrier on the surfaces.

Combination convection and surface cooling.It should be pointed out that the term “cooling air” does NOT infer that the air is cold or even cool. Air used to cool turbine components, which can be operating at temperatures of around 10000C, can be cooled by air, tapped from the higher compression stages at around 3000C+.

The two illustrations below show, on the left, an example of internal blade cooling, and on the right, an example of internal and surface blade cooling.


Turbine Blade Cooling Figure 6.5


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15.7 EXHAUST

The exhaust section of a turbo-jet engine is made up of several components, each of which has its own function. These components also have one common purpose; they must direct the flow of gasses rearwards in such a manner as to prevent turbulence and at the same time, impart a high final, (exit) velocity to the gases on turbo-jets, less so on turbo-propeller engines.

The exhaust section is located directly behind the turbine section, and ends when the gases are ejected at the rear. The individual parts of the exhaust include the exhaust cone, the tailpipe and the exhaust/jet nozzle.

EXHAUST CONE ASSEMBLY

The exhaust cone assembly consists of an outer shell or duct, an inner cone and a number of radial hollow struts or fins. The outer shell/duct is manufactured from heat resistant steel and attaches to the turbine case flange.

The duct is slightly divergent, due to the inner cone profile, even if the outer duct appears to be convergent. This slows the gas flow, slightly decreasing the velocity and increasing the pressure.

The radial struts serve two purposes, firstly to support the inner cone and, secondly, to straighten out the airflow which leaves the turbine with some ‘swirl’. The illustration below shows the main components of an exhaust cone assembly.

Exhaust Cone AssemblyFigure 7.1


7.1 TAILPIPE

The tailpipe is used, where necessary, to pipe the exhaust gases out of the airframe. Its use imposes a penalty on the efficiency of the engine in the form of heat and duct (friction) losses. These losses result in a measurable loss of final thrust.

In some designs a tailpipe is not required. For example, when the engine is installed in nacelles or pods, a short tailpipe is all that is required.

7.2 JET NOZZLE

The exhaust or jet nozzle imparts to the exhaust gases the all-important final boost in velocity. The jet nozzle, like the tailpipe, (when fitted), is not part of the basic powerplant but part of the airframe.

There are two types of jet nozzle design, the convergent design, for sub-sonic gas velocities and the convergent-divergent design for supersonic gas velocities. The jet nozzle openings may be either fixed or variable area, the fixed being the simpler of the two designs.

Convergent

The convergent nozzle accelerates the airflow, reaching Mach 1 and becoming “choked” at about the exit of the nozzle. When the gas exits the choked nozzle, it spreads out and accelerates. Attempting to accelerate the air any faster than Mach 1 would be uneconomic and reduce the engine life, due to the higher temperature. For faster speeds, a convergent-divergent nozzle is required.

Convergent-Divergent

This design is used mainly on supersonic aircraft, although it may be found on engines with high-pressure ratios. This type of nozzle is used to recover some of the otherwise wasted energy, by generating a further increase in gas velocity and, hence, thrust.

This type of nozzle uses the same principle as supersonic intakes. If the sub-sonic airflow is accelerated to reach Mach 1 at the narrowest point of the convergence, the air will then accelerate further, as the duct diverges, to reach a high supersonic airflow at the exit.


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The two types of nozzle are illustrated below, with the convergent at left and the convergent-divergent on the right. A third form, a variable area nozzle will be covered in the section on afterburning.

Jet NozzlesFigure 7.2

By-Pass Exhaust System

The modern fan or by-pass engine has two gas streams venting to the atmosphere, the high temperature gases being discharged by the turbine and the low temperature gases discharged from the fan section. These gasses may be exhausted separately or together.

In a low by-pass engine, the flows of cool air and hot air are combined in a mixer unit that ensures the mixing of the two streams prior to exiting the engine. This mixing also helps to reduce the exhaust noise.

An example of a low by- pass exhaust is illustrated overleaf.


High by-pass engines usually exhaust the two streams separately. The hot and cold nozzles are co-axial. A common nozzle may be used to partially mix the hot and cold gases prior to their ejection.

High By-Pass ExhaustFigure 7.4

Engine Noise Reduction

Noise is measured in effective perceived noise decibels, (EPNdB), which takes into account the pitch as well as the sound pressure, (decibels) and also makes allowance for the duration of an aircraft flyover. The noise produced by different types of engines is quite marked, as is the effect of the installation of noise suppressors. This can be seen on the chart below. ENGINE NOISE SUPRESSION

Engine Noise ReductionFigure 7.5


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The most significant sources of noise from an engine are the fan/compressor, the turbine and the exhaust. The noise exhaust is the effect that can be reduced by the largest amount. It is mostly generated by the shearing action between the jet exhaust and the outside air, although the eddies in the air can also cause high frequency noise for small eddies and low frequency for large ones.

A reduction in noise can be achieved if either the mixing rate of the two airflow’s can be accelerated, or the exhaust velocity, relative to the air can be reduced.

Methods of Noise Suppression

Noise suppression of internal sources is looked at in two ways, engine design and the use of acoustic panels to absorb noise. As both of these are beyond the requirements of the syllabus, they will not be covered. It is however, important to be aware of the fact that the honeycomb noise absorbing panels are fragile and must be treated carefully. (Illustration overleaf).

The exhaust, being a large generator of noise, can be made quieter by mixing of the high speed and the low speed air over a shorter distance. This is achieved by increasing the contact area of the atmosphere with the exhaust stream by using a propelling nozzle, which incorporates a corrugated or lobe type noise suppressor. A corrugated type of noise suppressor, as fitted to the Rolls-Royce Tay, is illustrated below.

The exhaust illustrated shows the corrugations which mix the two streams, hot and cold, by encouraging the hot stream to expand outwards whilst the cold, (by-pass), air is drawn inwards. This mixing improves efficiency and reduces the noise emitted by the engine.

Exhaust SilencerFigure 7.6


As mentioned earlier, the honeycomb acoustic panels are fairly fragile and, as can be seen from the illustration of a typical installation below, they line most of the by-pass ducts and care must be taken, during inspections, that they are checked for Foreign Object Damage(F.O.D.) or other causes.

Acoustic PanelsFigure 7.7

Thrust Reversal

Modern aircraft brakes are very efficient, especially carbon units but, on wet, icy or snow covered runways this efficiency may be severely reduced by the loss of adhesion between the tyres and the runway surface. To ensure continued operation during inclement weather, an alternative to friction brakes had to be found.

A simple and effective way to achieve this was to reverse the direction of the exhaust gas stream, thus using engine thrust as a decelerating force. An additional bonus to this system was that it could be used at all times and can, therefore be used to shorten landing runs, even when the runways are dry.

On rare occasions, it has been authorised to use thrust reversers in flight but, normally, they are only active when the aircraft’s landing gear is on the ground and the weight of the aircraft is on it, or the aircraft is ‘Weight On Wheels’(W.O.W.)


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An example of the effectiveness of reversers is illustrated below where the landing run can be reduced by more than 20%.

Landing Run DistancesFigure 7.8

The methods of reversing the flow vary with each engine but basically if the engine is a turbo-jet, the hot stream is reversed by either clamshell or bucket doors. On a high by-pass engine, because the by-pass, (cold), thrust is a large part of the total, only the cold stream is reversed, normally by blocker doors. This method is sometimes called the translating cowl system.

On propeller powered aircraft, the reverse thrust action is obtained by changing the pitch of the propeller blades, usually by hydro-mechanical means. This system moves the blades to a negative angle, but only after touchdown and pilot selection, producing a flow of air forwards, decelerating the aircraft.

The exhaust flow of jet engines can be directed forwards, (approximately 450) but, due to the risk of foreign object damage(F.O.D.), many engines are restricted in reverse, to about 70 or 80% of maximum power. This results in the fact that only a percentage of the forward engine thrust is available for reversing action.

The illustrations overleaf show three common methods of thrust reverser operation. All three methods are stowed with little drag but some weight penalty when not selected and actuated. The clamshell system, a hot stream method, reverses the airflow ahead of the exhaust nozzle. This method can be seen on Concorde.

The second method, also a hot stream system, has the buckets at the very rear of the exhaust/jet pipe. It also deflects the stream forwards and can be seen on many early Boeing small jets such as the 727 and 737 models.


Finally, the translating cowl/cold stream system uses a set of ‘blocker doors to direct the fan, (cold), airflow through a series of cascade vanes, often uncovered by a sliding cowling which smoothes the airflow over the vanes during flight.

Thrust Reverser TypesFigure 7.9


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INTENTIALLY LEFT BLANK

Mod 15 GAS TURBINES by COBC 82

15.8 BEARINGS AND SEALS (Level B1 only)

BEARINGS

Introduction

A bearing is any surface that supports or is supported by another surface. Bearings are designed to produce a minimum of friction and a maximum of wear resistance.Bearings must reduce the friction of moving parts and also take thrust loads or a combination of thrust and radial loads. Those which are designed primarily for thrust loads are called thrust bearings. The two different types of bearings used on gas turbines are ball and roller.

Ball Bearings

A ball bearing consists of an inner race, an outer race and one or more sets of balls; and bearings which are designed for dismantling, a ball retainer or cage. The purpose of the retainer or cage is to prevent the balls touching one another. Ball bearings are used for radial and thrust loads; a ball bearing specially designed for thrust loads would have very deep grooves in the races.

Roller Bearings

The bearings are manufactured in various shapes and sizes and can be adapted to both radial and thrust loads.The bearing race is a guide or channel along which the rollers travel; the roller is situated between an inner and outer race, both of which are made of case hardened steel. When the roller is tapered, it rolls on a cone shaped race inside an outer race.Straight roller bearings are used only for radial loads and taper roller bearings will support both radial and thrust loads. Roller bearings will withstand greater radial loads than ball bearings because of greater contact area.

STRAIGHT ROLLER BEARING


`Types of Bearings

Bearings are broadly classified by the type of rolling element used in their construction. Ball bearings employ steel balls which rotate in grooved raceways, whilst roller bearings utilise cylindrical, tapered or spherical rollers, running in suitably shaped raceways. Both types of bearings are designed for operation under continuous rotary or oscillatory conditions, but, whilst ball bearings and tapered roller bearings accept both radial and axial loads, other types of roller bearings accept mainly radial loads.


Air Sealing

LP, HP and IP air are all used to prevent the hot exhaust gases flowing inward between the stages of the turbine, by means of a labyrinth seal. It is also used to seal engine bearing housings to prevent oil leaks (sump areas).


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The figure represents a typical sump area, although in this case only one bearing is shown. Sump areas contain as many as five bearing assemblies.

TYPICAL SUMP AREAAs the air pressure is greater than the oil pressure, we can see from the figure that cooling air is directed into the air cavity of the sump. Oil is also directed into the oil cavity via the oil jet. For as long as the air pressure is greater than the oil pressure, the oil will be retained within the oil cavity. The oil seal is so designed to reduce the amount of air escaping across it. However, a certain amount of air will get into the oil cavity, which is then vented overboard via the oil cavity air vent orifice.


Air to Air or Air to Oil Sealing

The air at these different pressures must be prevented from mixing. This is achieved by inserting differential pressure seals at appropriate places within the engine.These seals are multi-groove types, but commonly known as labyrinth seals.

Labyrinth Seals

Labyrinth seals are constructed of metal non-rotating lands, which are secured to various parts of the engine case and a series of cylindrical rotating knife-edge steps that mate with the lands. With this type of seal, there are no contacting parts. A precise clearance is designed into the seals to control the pressure, as the compressor air passes over the cascade of knife-edges, the pressure is reduced.


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The labyrinth seal may be used in conjunction with an abradable coating on the stationary member as shown in the figure, or with a honeycomb shroud as shown in the figure.


Carbon Seals

Another method of air sealing is achieved by using a carbon seal arrangement. They are used on the rotating assembly of a gas turbine and protection of engine drive components on an accessory gearbox.

Carbon seals are manufactured of a mixture of carbon and graphite powder, bonded together with a viscous substance, such as coal tar. The carbon seal is fixed and held against the rotating seal by springs. Both the rotating seal and the carbon seals are machine ground and precision lapped to a micro finish.

Spring Ring Seal

This type of seal would normally be used around a main bearing assembly within the engine. It may be used in conjunction with a labyrinth or screw back type of seal. The location of a spring ring seal is shown in the figure.

Construction and Operation

This type of seal is similar to a large stepped piston ring; it is located on a rotating shaft. When the shaft is stationary, the seal clamps tightly to the shaft. As the shaft rotates, the spring ring can expand slightly, under centrifugal force, when it then forms an effective seal with the adjacent stationary housing.


`Hydraulic Seal

This type of seal may also be found protecting the bearings on the main rotating assembly of an engine. It is fitted between the rotating shafts on a twin or triple spool engine. A hydraulic seal would be used in conjunction with another type of seal, as shown in the figure.


The seal consists of a circular baffle ring mounted on a rotating shaft; the rim of this ring sits in the centre of a circular depression in an outer rotating shaft. Oil from the bearing will fill this depression and be held there by centrifugal force. This oil reservoir will form a liquid seal with the rim of the rotating baffle ring. Any tendency for the oil to leak across this seal will be counteracted by air leakage across a back-up seal.

Screw Back Seal

This type of seal will be found close to a bearing. It is mounted on a rotating shaft and will be backed up by one of the types of seal previously described.The figure shows the location of a screw back seal.

Screw Back Seal Location


This seal consists of a raised screw thread on a rotating shaft; the thread form is facing towards the bearing. Any oil from the bearing, which gets onto this threaded section, isthen ‘screwed back’ to the bearing, thus preventing oil loss. Air pressure on the outer end of this seal will also be screwed towards the bearing and helps to prevent a loss of oil.


15.9 LUBRICANTS AND FUELS

WARNING

Safety Precautions

Handling of synthetic lubricants requires precautions not needed for conventional lubricating oils. Synthetic lubricants have a high solvent characteristic, which causes them to penetrate and dissolve paints, enamels and other materials, including crepe soles on footwear.

Additionally, if synthetic oils touch or remain on the skin, physical injury can result. It is essential, then that these lubricants are kept away from the skin, either by the use of ‘barrier creams’ or by wearing protective garments and safety glasses. Any part of the skin that is affected by a spillage of synthetic oil, should be treated in accordance with local instructions and the relevant COSHH leaflet. (Containment of Substances Harmful to Health)

When handling fuels and oils, all normal precautions regarding flammable substances that were covered earlier, in Module 7, should be followed. This should include correct storage, carriage and dispensing, together with Earthing and cleanliness at all times.

PROPERTIES AND SPECIFICATIONS

Kerosene Fuels

Aviation turbine fuels are used for powering turbojet, turboprop and turboshaft engines. There are two main types of turbine fuels in use, JET A and JET A-1, which are kerosene types and JET B which is a blend of gasoline and kerosene fractions. In the United Kingdom these fuels are sometimes referred to as AVTUR, (JET A & JET A-1); and AVTAG, (JET B). JET B is also known as JP4 in the United States Military.

Finally, there is a high flash point, low freezing point fuel known as JP 5 or AVCAT. Its use is limited, due to its high flash point making it less liable to ignition in an accident, but can be used on aircraft carriers and other aircraft carrying sea craft.


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ALL fuels should meet the requirements stated below:

Ease of flow under all operating conditions (Sometimes called ‘pumpability’). Quick starting of the engine under all conditions. Have a high calorific (energy) value. Must be non-corrosive. Combustion by-products should not be harmful to internal engine parts. Minimal chance of fire. Provide lubrication of the moving parts of the fuel system (pumps etc.).

Pumpability

There are several factors affecting ‘pumpability’; fuel viscosity, solids (wax and gum) and ice particles (due to water in the fuel). The lowest temperature at which the fuel can be pumped is known as the “pour point”.

Volatility and Starting

Quick starting of an engine depends on fast ignition and the quality of the fuel. It must remain volatile, (easily evaporated), at starting temperatures so that the fuel spray from the burners will readily ignite.

Combustion

Once ignited the fuel must burn completely, giving both the highest energy value and also producing benign combustion products (Carbon).

Calorific Value

This is the amount of heat released during combustion. Turbine fuels have a slightly lower calorific value, per unit weight, than piston engine fuels, (Petrol/Gasoline). However, as turbine fuel is heavier with a higher specific gravity, it releases more heat per unit volume. Fuel with a high calorific value is most suitable for aviation turbine use.

Non-Corrosive

To reduce corrosion within the fuel system, the fuel must be a good lubricant. As basic kerosene has little lubricity, additives are used to improve its lubrication properties.

Fire Hazards

Fire is always a risk during handling, often due to spillage, the presence of electrical sparks, contact with hot engine parts, etc. Gasoline ignites more readily then kerosene, as it has a lower flash point.


OILS

Lubricating oils, like other petroleum products, are a mixture of various hydro-carbons. The viscosity basically depends upon the process used to refine the oil and blending can further control it.

Viscosity and Viscosity Index are the factors that decide the lubricant for a particular purpose. The desired viscosity of oil for a particular engine is decided by the designer considering many factors, including:

Bearing loads and clearances Sliding speeds Oil pump capacity Operating temperatures Engine RPM

The specification of an oil indicates the properties it possesses, thus ensuring that it is able to safely protect, cool, lubricate, etc. the moving parts of the engine. It is essential then, that the correct oil is always used to top-up or re-fill an engine and oils are NEVER mixed together.

For the lubrication of the main shafts of jet engines running in contact bearings, low viscosity oil is required. Early engine designs operated on straight mineral oils, but these were unsatisfactory when low temperature starting was involved, either on the ground or when re-lighting in the air. To overcome this problem, and that of the high contact pressures at the bearings, low viscosity, extreme pressure, (EP), oils were developed.

When mineral oils reached the end of their useful life, they were replaced with synthetic oils. These were initially developed from esters of sebacic acid and, as the basic oils were unsuitable for carrying the bearing loads in the engines, other, more complex esters were added to assist the load carrying and to raise the viscosity. This ‘first generation’ synthetic oil became AeroShell 750, (service designation OX-38), and was produced by most petroleum companies under different names.

What is known as ‘second generation’ oils came about because of the by-pass and turbo-fan engines, which, due to the insulating effect of the by-pass air, caused the oil temperatures to rise. This meant that a new oil, which could operate at higher temperatures and resist oxidation had to be developed. These are known as Type 2 lubricants. Their specification includes anti-oxidants, load carrying additives, corrosion inhibitors, metal deactivators and foam inhibitors. A typical example might be AeroShell 500.

There are ‘third generation’ oils developed for use during supersonic operation, both with the military and in the Olympus engines on Concorde. These have resistance to very high oil temperatures, in the range of 2600 to 3150C, and, whilst still being developed from ester based oils, they have poorer properties both at low temperatures and for lubricity. An example might be AeroShell 555.


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In summary, oil has a range of characteristics for different applications, it can be derived from three main sources and can contain a wide range of additives to change its properties.

Characteristics:

Ability to wet the surface to be lubricated Viscosity that maintains film lubrication High viscosity index Low rate of evaporation at high temperature Prevents formation of gum and sludge Must remain stable in use

Sources:

Mineral Synthetics Vegetable (rarely used on aircraft)

Additives

Extreme pressure Anti-corrosion Viscosity improvers Pour point depressants Anti-foaming Anti-oxidants

Fuel Additives

Additives to the basic fuel specification are pre-mixed by the supplier of the fuel. All the engineer can do is to be aware that certain fuels, with specific additives, are only to be used where specified by the engine manufacturer.

For example, anti-icing agents are added to fuels to limit the freezing of entrained water, without recourse to fuel heating, at low temperatures. Also, the addition of anti-microbiological agents to the fuel helps to kill the microbes, fungi and bacteria, which form slime or, occasionally, a matted waste in the fuel tanks and pipework.

Occasionally, an additive has not been added during refinement, meaning the engineer has to add the applicable agent, in the correct quantity, during refuelling. A popular brand of a combined anti-icing and anti-microbiological mixture is called PRIST. It is designed to be added during servicing. However, the engineer must determine the type and amount, after consultation with the maintenance manual, the operator’s manual or the Type Certificate Data Sheet.

Alternatively, many gas turbine engine manufacturers approve an anti-biological compound called “Biobor”, as an additive to the aircraft fuel supply.


15.10 LUBRICATION SYSTEMS

The gas turbine engine lubrication systems are required to provide lubrication and cooling for all gears, bearings and splines. They must also be capable of collecting foreign matter, which, if left in a bearing housing or gearbox, can cause rapid failure. In addition, the oil must protect the lubricated components that are manufactured from non-corrosion resistant materials.

With turbo-propeller engines, there are somewhat different requirements to any other types of gas turbine. This is due to the additional lubrication of the heavily loaded propeller reduction gears and the need for a high-pressure oil supply to operate the propeller pitch control mechanism.

Most gas turbines use a self-contained recirculatory lubrication system, in which the oil is distributed around the engine and returned to the oil tank by pumps. There are a few engines that use a system known as the total loss or expendable system in which the oil is dumped overboard after the engine has been lubricated.

TYPES OF SYSTEMS

There are two basic types of recirculatory system. They are known as the pressure relief valve system and the full flow system. The major difference being in the control of the oil flow to the bearings.

Because, in both designs, the oil temperature and oil pressure are critical to the safe running of the engine, provisions are made to display both parameters in the cockpit.

PRESSURE RELIEF VALVE SYSTEM

In this system, the oil flow to the bearings is controlled simply by limiting the pressure in the feed line to a given design value. This is achieved by the use of a spring-loaded valve, which allows the oil to be directly returned from the pressure pump outlet to either the oil tank or the pressure pump inlet, when the design pressure is exceeded.

A limitation of this system is that when the engine is at idle, the oil is being pumped around the engine due to the valve being off its seat. Once the engine speed increases however, the bearing chamber air pressure increases, slowing down the flow of oil. To overcome this problem on some engines, the rising pressure is fed to the back of the oil pressure relief valve, which effectively increases the oil pressure in the feed line.


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The basic components that could make up this type of engine lubrication systemwould be:

The pressure pump, which draws oil from the tank, through a strainer, to the pressure filter.

The pressure relief valve, which maintains a constant delivery pressure. A second pressure relief valve, is sometimes fitted, set well above system pressure.

This opens if the system becomes blocked. The filter by-pass valve, which opens if the filter becomes blocked. The scavenge pumps which return the oil to the tank via the oil cooler.

Illustrated below is a typical pressure relief valve type oil system, installed in a turbo-propeller engine.

Typical Pressure Relief Valve SystemFigure 10.1


FULL FLOW OIL SYSTEM

Although the pressure relief valve system operates satisfactorily for engines that have a low bearing chamber pressure that does not unduly increase with engine speed. It becomes an undesirable system for engines that have high chamber pressures. For example, a chamber pressure of 90 psi. requires a relief valve set at 130 psi. This calls for large pumps, with the associated difficulty in matching the oil flow at lower speeds.

The full flow system achieves the desired oil flow rates throughout the complete engine speed range. To achieve this, the pressure relief valve is dispensed with and the pump output directly supplies the oil feed jets.

The example below shows a turbo-fan engine in which the size of the pressure pump is dictated by the flow at maximum engine speed. Using this method allows smaller pressure and scavenge pumps, due to there being no continuous loss of oil spilling back to the tank, that occurs at high engine speeds, with the other system.

Typical Full Flow Oil SystemFigure 10.2


`To prevent high oil pressure from damaging filters or coolers, relief valves are fitted to by-pass these units. They will normally only open during a cold start or in the event of a blockage of the internal parts.

A differential switch gives advanced warning of a potential filter blockage by sensing the difference between inlet and outlet pressures.

TOTAL LOSS (Expendable) SYSTEM

For engines that run for short periods, such as booster and lift engines, the total loss oil system is sometimes used. This system is simple and incurs a low weight penalty because it does not require an oil cooler, scavenge pump or filters. On some engines oil is delivered in a continuous flow to the bearings via a plunger-type pump driven by the compressor. Some pumps are driven by fuel pressure and, when opening the HP cock during starting, it directs a shot of oil to top and bottom, (front and rear), bearings. The oil is then either ejected into the exhaust or retained and either drained or dumped overboard.

The basic oil system illustrated below, is typical of many gas turbine engines, and contains all of the components mentioned earlier.

Basic Oil SystemFigure 10.3

The previous system is actually from a turbo-propeller engine, hence the additional oil supply to both the reduction gearbox and the torque meter system. Other items of note are:


The measurement of the oil temperature immediately after leaving the tank,

The location of the oil pressure transmitter at the furthest point from the pressure pump.

The stack pipe, (stand pipe), which reserves some oil for propeller feathering, even if the lubricating oil is totally lost due to a leak.

The ‘thread type’ oil filters, which protect the bearings from the finest pieces of swarf.

A de-aerator tray that removes air bubbles from the returned oil.

SYSTEM COMPONENTS

10.1 OIL TANKS

The oil tank is usually mounted on the engine but is normally a separate unit, although it can be an integral part of an external gearbox. It normally contains a number of items, such as draining and filling facilities, methods of checking the oil quantity via a dipstick and sight glass, provision to remove air bubbles from the returned oil and, on some non-commercial aircraft, facilities to allow agile and inverted flight.

10.2 PUMPS

The pumps can be divided into two groups. The pressure pump takes the oil from the tank and pushes it throughout the engine to the bearings, gears and accessories. The scavenge pumps collect the oil after it has served its purpose and return it to the tank, possibly via a cooler and de-aerator tray.

The simple gear type pump is the commonest in use for both of the above purposes. The oil is drawn into the low-pressure inlet and passed around the outer chamber and out into the system, as can be seen in the illustration over-leaf.

10.3 FILTERS

Solid contaminants pumped through an aircraft engine lubricating system can clog the oil passages and damage the bearings. Provisions must be made to remove as much of these as possible.

This is achieved by one of two methods, full-flow filtration and by-pass filtration. All of the oil circulating through the system passes through a full-flow unit, whilst only part of the oil is filtered during each circulation in a by-pass system, although all of the oil is filtered eventually. By-pass filters can be much finer because if they clog, the oil can continue to flow around the filter, retaining the essential lubrication.

10.4 RELIEF VALVES


`Almost all pumps used in aircraft engines produce an excess of oil pressure, which must be controlled. Pressure relief valves, which can be a simple spring loaded plate or valve, remain closed until the oil pressure rises excessively. The valve will then open and spill a percentage of the oil back to the tank, lowering the pressure to the relief valve spring value.

Many relief valves have an adjusting screw, which can be unlocked and turned, to change the value at which the valve opens.

10.5 OIL PRESSURE GAUGE

Pressure is measured at crucial points around the system, depending on the design of the engine. Normally, the system pressure, if measured at a single point, will be sensed far from the pump, so that any system leaks are detected as well as any failure of the supply itself. A simple transducer will be connected to the system, which will send a signal to a gauge unit on the instrument panel.

Often a pressure switch will be located at the same place as the transducer, this will illuminate a warning light or caption as a back-up to the gauge unit, if the system pressure falls below a pre-determined figure.

10.6 OIL TEMPERATURE GAUGE

The oil temperature is usually measured at the inlet to the pressure pump, (hence its name, Oil Inlet Temperature, (OIT.). This ensures that if there is a rise in temperature due to a low oil quantity or a blocked oil cooler it will be detected rapidly.


The illustrations below show examples of some of the components mentioned earlier. Each component on each engine will have been designed for its specific purpose and will therefore, be different from similar components fitted to other engines.

OIL TANK ASSEMBLY OIL PUMP

OIL FILTER RELIEF VALVE

Oil System Components Figure 10.4


`10.7 OIL COOLER

All engines transfer heat to the oil by friction, churning and windage within a bearing chamber or gearbox. To remove this heat an oil cooler is fitted, normally in the return line, to transfer the heat from the oil to either the atmosphere and/or the fuel system.

It is possible to find, on either type of cooler, a temperature sensitive valve that will be set into the inlet. This valve will by-pass the cooler when the oil is still at ambient temperature, i.e. at starting, and it will open as the oil temperature rises.

10.8 MAGNETIC PLUGS

These items are also known as Chip Detectors and are fitted into the scavenge,(return),line to collect ferric debris from each bearing chamber. They are basically permanent magnets inserted in the oil flow and are retained in self-sealing valve housings.

Upon examination, they can provide a warning of impending failure without having to remove and inspect filters and without having to carry out other troubleshooting operations on the system. They are usually removed and inspected during scheduled maintenance inspections for condition monitoring purposes.

Oil Cooler (Left) And Magnetic PlugFigure 10.5


OIL SYSTEM MAINTENANCE PRACTICES

The maintenance of gas turbine engine lubrication systems consists mainly of adjusting, removing and replacing various components.

OIL CHANGE

The first action should always to ensure that the type of oil in the engine is identified by reference to the maintenance manual. There are many different makes of synthetic oil, both type 1 and type 2, and great care must be taken to ensure that the oil put into the engine is of the correct specification.

Oil is usually supplied in quart containers and care must be taken to ensure that both the container and the replenishing point of the engine are clean. If bulk replenishing rigs are used, the correct filtration must be serviceable, usually 10 micron or smaller is normal.

If the engine is being drained and re-filled with different oil, it may be drained from the oil tank, the accessory gearbox sump, the main oil filter and other low points of the oil system. The engine will be flushed first, by refilling the engine with flushing oil and then motoring it over using only the starter motor. Once this has been accomplished, the flushing oil should be drained and the engine refilled with the new oil.

Another important consideration when servicing the oil system is to ensure that servicing is accomplished within a short time after shutdown. This is normally called for by the engine manufacturers to assist in drain down, during an oil change, with the hot and thin oil. It is also to prevent over filling because oil will, over time, drain into the engine causing the tank to show a lower level than the correct value. The manufacturer will often say, for example, oil levels must be checked between 15 and 30 minutes after shut down.

Another consideration is the recording of replenishments. A careful record of all oil put into the engine must be kept in the technical log and, whilst a small, regular consumption is acceptable, a slowly increasing quantity required for replenishment requires investigation.


`OIL FILTER MAINTENANCE

Depending upon the aircraft manufacturer’s published instructions, filters may be replaced at a published life or, on occasions, if the design of the filter is different, it may be cleaned either by flushing or by ultrasonic cleaning. The latter vibrates the filter at very high frequency, in a solvent, which effectively shakes all the particles out of the filter, leaving it clean and ready for re-installation.

An example of one of these ultrasonic cleaners is shown below.

Ultrasonic Filter CleanerFigure 10.6

SCAVENGE SYSTEM

The scavenge systems remove oil from the bearings and gearboxes, by scavenge pump suction, and returns it to the oil tank.

The systems will normally contain scavenge filters, usually a coarse metallic grid, to remove any metallic particles returning to the tank from bearings, gearboxes, etc. In critical sub-systems, magnetic chip detectors may also be fitted in the return scavenge lines, to collect ferrous particles.

These ‘plugs’ will be removed and inspected at regular intervals in accordance with the engine manufacturer’s manuals. Some chip detectors, in addition to being magnetic, can have electrical contacts in them, which will give a flight deck warning if particles of metal are attracted to the magnet, giving the crew the option to close down the affected engine.




`15.11 FUEL SYSTEMS

The functions of the engine fuel system are to provide the engine with fuel in a form suitable for combustion and to control the flow to the required quantity necessary for easy starting, acceleration and stable running, at all engine operating conditions. To do this, one or more fuel pumps are used to deliver the fuel to the fuel spray nozzles, which inject it into the combustion chambers in an atomised spray.

Because the flow rate must vary according to the amount of air passing through the engine, to maintain a constant selected engine speed or pressure ratio, the controlling devices, whether mechanical or electronic, are fully automatic. This is excepting the engine power selection from the flight deck, which is achieved via a manual throttle or power lever.

A fuel shut-off valve or ‘cock’, lever is also normally used to stop the engine. In some cases, these two levers, power and shut-off, are combined into one lever with both functions, (usually protected against accidental closure by a ‘gate’).

There will also be protection devices fitted that prevent excess compressor delivery pressure, (choking) and rotating assembly overspeeding.

With turbo-propeller engines, changes in propeller speed and pitch have to be taken into consideration, as they affect the power output of the engine. Usually, the throttle lever and propeller control unit will have to be inter-connected, thus the correct relationship between the fuel and air flow is maintained at all engine speeds. This arrangement also gives the pilot the advantage of single lever control. Although the propeller will control the speed of the complete assembly, an additional governor in the fuel system acts as a back-up Overspeed protection.

MANUAL AND AUTOMATIC CONTROLS

The control of the power or thrust of a gas turbine engine is achieved by the quantity of fuel injected into the combustion system. When a higher thrust is required, the throttle/power lever is advanced and the pressure to the fuel spray nozzles increases giving a greater fuel flow. This has the effect of increasing the gas temperature, which, in turn, accelerates the gases through the turbine assembly, giving higher engine speeds and a greater airflow. The result of all this is to produce more engine thrust.

Another variable which has to be taken into consideration by any fuel control device, is the change of air density due to altitude, air temperature and aircraft speed. All of these influence the density of the air, and hence the mass air flow, entering the intake of the engine.

If these variables were not taken into consideration, the fuel would not be reduced in proportion to the falling air density, as the aircraft climbed. The result would be a steep rise in the exhaust temperature and the risk of overheating or possibly destruction of the turbine assembly.

Many engines are fitted with an electronic system of control and this generally involves the use of integrated circuits, (ICs), to measure and translate changing engine


conditions to automatically adjust the fuel pump output. Some helicopters also have electronic engine control which, in this case, have an additional free-turbine and hence rotor speed control.

FUEL CONTROL SYSTEMS

Typical high pressure fuel control systems for both turbo-propeller and turbo-jet engines, consist of basically the following components:

HP fuel pump Throttle control Fuel spray nozzles Sensing devices for flow and pressure

The usual method of varying the fuel flow to the spray nozzles is by adjusting the output of the HP fuel pump, which is signalled through a servo system in response to some or all of the following inputs:

Throttle movement. Air temperature and pressure. Rapid acceleration and deceleration. Signals of engine speed, exhaust gas temperature and compressor delivery

pressure.

Simple Fuel System Figure 11.1


`11.1 PRESSURE CONTROLS (Turbo-prop engine)

The pressure control system shown below contains just the components mentioned previously. The fuel pump (top left) output, is controlled by the spill valve in the fuel/flow control unit (right) and the engine speed governor, (left centre). These valves vary the servo pressure and hence correct the pump output.

Changes to the aircraft speed, altitude or the outside air temperature, (OAT), will result in the changes being sensed by the bellows units in the Fuel Control Unit, FCU, (right) and therefore, further alterations to the servo pressure, altering the pump output to the correct amount.

Turbo-Propeller Fuel SystemFigure 11.2

From the above diagram, the basic operations are:

Movement of the throttle lever. Aircraft climbing and descending. Aircraft moving faster or slower. Opening and closing of the High Pressure Cock, (HPC).


All of the first three operations occur during each flight and involve signals being sent to the fuel pump. The last operation occurs on the ground at the start and finish of each flight or period of engine running.

The throttle valve moves as the pilot moves his throttle lever. This causes changes in the differential pressure around the valve and signals the pump to increase or decrease its output to match the demand.

Any change in either outside ambient or intake pressure will result in more or less air entering the engine. This will require a change in fuel flow to match it. The bellows, in the capsule assembly, will expand or contract, depending on circumstances, again altering the fuel pump output to match the airflow.

Finally, to stop the engine, a separate lever is normally operated. The HP cock, cuts off the fuel to the burners. However, the mechanical fuel pump continues to run as the engine ‘winds down’, so the fuel still being pumped has to be recirculated back to the LP side of the supply.

11.2 PRESSURE CONTROLS (turbo-jet engine)

Whilst the propeller controls the speed of the engine on a turbo-propeller installation, the acceleration and deceleration of a turbo-jet is in direct proportion to the change of fuel flow. These fuel control units control both the amount of fuel being fed to the burners and the rate at which the fuel is increased and decreased.

The rate is very important because if the fuel is increased at too fast a rate, the engine can quickly overheat and burn out, before it has time to accelerate and draw in sufficient air to mix with the extra fuel. During throttle closing, however, if the fuel flow is decreased too quickly, the combustion flame can ‘blow out’ due to there being too little fuel for the mass of air still coming through the engine.

Otherwise, the detection of throttle opening, the changes in atmospheric pressure and airspeed are catered for by much the same way as the previous example.

11.3 FLOW CONTROL SYSTEM

A flow control system is generally more compact than a pressure control system and it is not sensitive to the flow effect of variations downstream of the throttle. The fuel pump delivery pressure is related to engine speed, thus, at low engine speeds pump delivery pressure is quite low. The fuel pump output is controlled to give a constant pressure difference across the throttle valve at constant air intake conditions.

Various devices are also used to adjust the fuel flow for intake pressure variations, idling and acceleration control, gas temperature and compressor delivery pressure control.


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11.4 COMBINED ACCELLERATION AND SPEED CONRTOL (CASC)

This is a mechanical system without small restrictors or spill valves. It is also an all- speed governor system and, therefore, needs no separate governor unit for controlling the maximum RPM The controlling mechanism is contained in one unit, normally referred to as the fuel flow regulator, FFR.

An HP fuel pump is used, with the pump servo piston being operated by HP fuel on one side and main nozzle (servo) pressure on the other side.

The fuel flow regulator, shown below, contains inputs from the HP fuel supply, (HP pump), and air pressure tappings from P2.6 and P3. In addition, the drive shaft is driven from the accessory gearbox, running at a speed that is proportional to engine speed.

The outputs are the primary and main flows leading to the spray nozzles.

Combined Acceleration And Speed Control UnitFigure 11.3


11.5 FUEL FLOW REGULATOR

The fuel flow regulator, driven by the engine, has two centrifugal governors, (speed control and pressure drop control) and two sliding valves, which also rotate.

The first valve, known as the variable metering sleeve, has a triangular orifice known as a variable metering orifice, VMO. This sleeve is moved by the capsule assembly. The governor sleeve, sliding over the VMO sleeve, is moved by the speed control governor and the stirrup arm, both of which are controlled by the throttle lever in the cockpit.

The other, pressure drop valve, is also controlled by a governor and forms a piston. It has a triangular, variable orifice and a rectangular, fixed-area orifice. Primary fuel flow comes from the fixed area orifice and, on its own, will provide a satisfactory fuel flow for idling at all altitudes. The triangular, variable orifice allocates the amount of extra fuel to the Main fuel flow in proportion to throttle opening, HP fuel pressure, engine RPM and VMO pressure.

11.6 ELECTRONIC ENGINE CONRTOL

Some engines utilise a system of electronic control to monitor engine performance and make necessary control inputs to maintain certain engine parameters within pre-determined limits. The main areas of control are engine shaft speeds, (N1, N2 and N3), and exhaust gas temperature, EGT, which are continuously monitored during engine operation. Some types of electronic control function only as a limiter, that is, if the shaft speeds or the EGT approach dangerous levels, an input is made to the fuel flow regulator, FFR, to reduce the fuel flow thus maintaining shaft speed or EGT at safe levels.

Supervisory systems, such as those mentioned previously, may contain a limiter function but, basically, by using aircraft generated data, the system enables a more appropriate thrust setting to be selected quickly and accurately by the pilot. The control system then makes small control adjustments to maintain engine thrust consistent with that pre-set by the pilot, regardless of changing atmospheric conditions.

Full Authority Digital Engine Control, (FADEC), takes over virtually all of the steady state and transient control intelligence and replaces most of the hydro-mechanical and pneumatic elements of the fuel system. The fuel system is thus reduced to a pump and control valve, an independent shut-off cock and a minimum of additional features necessary to keep the engine safe in the event of extensive electronic failure.


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As an example, the Rolls Royce RB211 engine FADEC system shown below, demonstrates how the computer known as the Control Amplifier, has inputs of engine temperature and shaft speeds and control output to the differential pressure regulator.

At high power settings, when there is a risk of shaft Overspeed or excess temperatures, the pressure regulator returns excess fuel to the pump inlet. The fuel flow regulator acts as a hydro-mechanical control, with inputs from the high-speed compressor, the gas path pressures and the power lever position.

Typical FADEC SystemFigure 11.4

SYSTEM LAYOUTS

Engine fuel systems vary in detail, but generally they contain similar components which do the same jobs. It can be seen from the two illustrations below, showing typical turbo-propeller and turbo-jet installations, that components like HP pumps; throttle unit/fuel flow regulators; shaft governors and HP shut-off cocks are common to both installations. It will also be seen that the sensors measuring such items as intake temperature; internal pressures; exhaust gas temperature and shaft speeds are, again, common to both designs.


Engine Fuel System LayoutsFigure 11.5


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There are a number of components, fitted to gas turbine engines, which require either some limited maintenance or adjustment during their installed lives. The short explanation and illustrations that follow will cover the most common of these.

ENGINE DRIVEN FUEL PUMP (EDP)

This pump delivers more fuel as the engine speeds up. It is designed to deliver a continuous supply of fuel to the fuel control at a quantity in excess of the engine needs. After metering the required amount of fuel to the combustor, the fuel control unit returns the surplus fuel to the pump inlet.

Main pumps can be either spur gear types, (positive displacement – output speed), or plunger types which have their output dependent on a servo signal from the throttle position as well as the engine speed. The pump illustrated below is a single unit, variable stroke, plunger type. The output of this type of pump can vary from about 100 to 2000 gallons per hour, at 2000 psi. depending on demand.

Variable Output Fuel PumpFigure 11.6


FUEL SPRAY NOZZLES

The three major types of atomising fuel spray nozzles are Simplex, Duplex and Airspray. These are the final component in the engine fuel system and the engine performance depends upon their correct operation.

Their task is to either atomise or to vaporise the fuel to ensure its rapid burning. The difficulties involved in this process can be appreciated when the velocity of the compressor outlet air and the short burning length available is considered.

The basic simplex burner imparts swirl to the fuel and then, after straightening out the swirl, it atomises the fuel from when it can be ignited to produce energy. This type of burner was first used on early jet engines and contained a chamber that induced the swirl to the fuel. Its main limitation was that the pressure required to achieve very high flows was difficult to achieve with the pumps available in those days.

A cutaway of a simplex burner, together with a typical spray pattern is illustrated below.

Simplex nozzleFigure 11.7

The duplex spray nozzle, (overleaf), requires a primary and a main fuel manifold and has two independent orifices, one much smaller than the other. The smaller orifice handles the lower flows, whilst the larger orifice deals with the higher flows as the fuel pressure increases. A pressurising valve or flow divider may be employed to apportion the fuel to each manifold, depending on the demand.


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Duplex burnerFigure 11.8

In this system, the duplex nozzles are able to give effective atomising over a wider flow range than the simplex nozzle for the same fuel pressure. These nozzles are also more effective in atomising the fuel at the low flows required at high altitudes.

The Airspray nozzle (right) carries a proportion of the primary combustion air with the injected fuel. By aerating the spray, the local fuel rich concentrations, produced by the other types of spray nozzle, are avoided. This gives a reduction in both carbon formation and exhaust smoke.

Another advantage of this type of nozzle is that the low pressures required for atomisation of the fuel permits the use of the comparatively lighter gear-type pump.

Airspray NozzleFigure 11.9


OVER-SPEED PROTECTION

Some engines require additional protection against overspeeding, other than the normal fuel control unit with engine speed input. A device similar to that illustrated (right) might be found on some engines, in this case the Rolls Royce Tay Turbo-fan engine, to limit the rotational speed of the LP fan. It achieves this, in a potential overspeeding condition, by directly restricting the fuel flow to the burners, via the metering plunger, which is lifted by the centrifugal governor.

Overspeed RPM GovernorFigure 11.10

Other components found within some engine fuel systems would include Filters, which remove any foreign particles not removed by earlier filtration and, fuel heaters, which ensure any ice crystals entrained in the fuel are melted before they can block any filters.


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15.12 AIR SYSTEMS

A definition of the engine internal air system is “Those airflow’s which do not directly contribute to the engine thrust”. The system has several important functions to perform for the safe and efficient operation of the engine. These functions include internal engine and accessory unit cooling, bearing chamber sealing, prevention of hot gas ingestion into the turbine disc cavities, control of bearing axial loads, control of turbine blade tip clearance and engine anti-icing.The internal air system also supplies air for aircraft services. Up to one fifth of the total engine core mass airflow may be used for these functions.

An increasing amount of work is done on the air, as it progresses through the compressor, to raise its pressure and temperature. To reduce performance losses, the air is taken as early as possible from the compressor, relative to the requirement of each particular function. The cooling air is usually expelled overboard, once it is of no further use. It may, however, be fed into the main gas stream, at the highest possible pressure, where a small performance recovery is achieved.

COOLING AIR

At the design stage of a gas turbine engine, it must be designed to ensure that certain parts of the engine, and in some instances certain accessories, do not absorb heat to the extent that it is detrimental to their safe operation. The principal areas that require air-cooling are the combustor and turbine.

Some internal airflow is used to control the temperature of the compressor shafts and discs by either cooling or heating them. This ensures an even temperature distribution and therefore improves engine efficiency by controlling thermal growth and thus maintaining minimum blade tip and seal clearances. The illustration below shows how the L.P, HP intermediate and HP air are used to cool the various internal parts of a twin-spool by-pass engine.

Cooling Airflow’sFigure 12.1


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TURBINE COOLING

A high thermal efficiency depends on a high turbine inlet temperature, which is limited by the turbine and nozzle guide vane materials. Continuous cooling of these components allows their environmental operating temperature to exceed the melting point of the material, without affecting the blade and vane integrity.

Heat conduction from the turbine blades to the turbine disc requires the discs to be cooled also and thus prevent thermal fatigue and uncontrolled expansion and contraction rates.

The life of turbine blades and vanes depends not only upon their form, but also on the method of cooling used therefore, the flow design of the internal passages is very important. There have been numerous methods of cooling used over the history of gas turbines. Single pass internal cooling was the most basic method used in the early designs, whilst later designs have multi-pass internal cooling with external air film cooling as well. Examples of three turbine blade design, from different eras, are illustrated below.

Turbine Blade CoolingFigure 12.2


It will be noticed that the later blades use both LP and HP cooling air to control the operating temperature of the blades. These differing air supplies are taken from tappings at different places along the compressor stage.

BEARING CHAMBER COOLING

Air cooling of the engine bearing chambers is not normally necessary since the lubrication system is adequate for cooling purposes. The bearings are also normally located in the cooler regions of the engine.

ACCESSORY COOLING

A considerable amount of heat is produced by some of the engine accessories; the electrical generators for example. Separate cooling sub-systems may be necessary for some components like these when the aircraft is on the ground, which may use engine bleed air or atmospheric air, which is ducted from outside the cowling.

Whilst the bleed air is supplied from a pressure tapping, external air must be induced to pass through the cooling system ducting. This is achieved by using compressor delivery air passing through nozzles in the outlet duct, creating a low-pressure area.This induces a flow though the cooling system from the inlet louvers to the outlet duct, as shown below.

Accessory CoolingFigure 12.3


`SEALING AIR

Seals are used to prevent oil leakage from the engine bearing chambers, to control cooling airflow’s and to prevent ingress of the mainstream gas into the turbine disc cavities. Various sealing methods are used on gas turbine engines. The choice of which method is dependent upon the surrounding atmosphere and pressure, resistance to wear, heat generation, weight, space available, ease of manufacture and ease of installation and removal.

The typical turbine assembly illustrating hypothetical cooling and sealing arrangements below, shows the usage of most of the common methods of sealing, which will be described later.

Sealing Air ExamplesFigure 12.4


LABRYNTH SEALS

This type of seal is widely used to retain oil in bearing chambers and as a metering device to control internal airflow’s.

The seal comprises a finned rotating member with a static bore, which is lined with either a soft abradable material or a high temperature honeycomb structure. On initial running the fins lightly cut small grooves into the static bore material, leaving a minimal clearance. This clearance varies throughout the flight cycle, depending on the thermal growth of the parts and the natural flexing of the rotating members.

Across each seal fin there is a pressure drop resulting in a restricted flow of sealing air from one side of the seal to the other. When used for bearing chamber sealing, it prevents oil leakage by allowing air to flow from the outside to the inside of the chamber, which also induces a positive pressure to assist the oil return system. The illustrations below show two different uses of labyrinth seals, an oil and air seal and a simple airflow control seal.

Labyrinth SealsFigure 12.5

RING SEALS

A ring seal comprises a metal ring housed in a close fitting groove in the static housing.

The clearance between the ring and the shaft is smaller than with a labyrinth seal, due to the ring being able to move in the groove, when it contacts with the shaft.

Ring seals are used for bearing chamber sealing, except in the hot areas where oil degradation due to heat would lead to ring seizure within its housing.


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Ring Type Oil SealFigure 12.6

HYDRAULIC SEALS

This method of sealing is often used between two rotating members to seal a bearing chamber. Unlike the labyrinth and ring seals, it does not allow a controlled flow of air to traverse across the sea.

These seals are formed by a seal fin immersed in an annulus of oil, which has been created by centrifugal force. Any difference in air pressure inside and outside of the bearing chamber is compensated by a difference in oil level either side of the fin.

CARBON SEALS

Carbon seals consist of a static ring of carbon, which constantly rubs against a collar on a rotating shaft. Several springs are used to maintain contact between the carbon and the collar. This type of seal relies upon a high degree of contact and does not allow oil or air leakage across it. The heat caused by friction is dissipated by the oil system.


BRUSH SEALS

These seals comprise a static ring of fine wire bristles. They are in continuous contact with a rotating shaft, rubbing against a hard ceramic coating. This type of seal has the advantage of withstanding radial rubs without increasing leakage.

Carbon (Left) And Ceramic SealsFigure 12.7

Aircraft Services

To provide a wide range of air services, such as cabin pressurisation/air conditioning, airframe/engine anti-icing, cross starting and, finally pressurisation of water and hydraulic tanks, substantial quantities of air are required from the compressor. It is desirable to bleed the air as early as possible from the compressor to minimise the loss on engine performance.

However, during some phases of the flight cycle, such as a low engine speed during the descent and approach, it may be necessary to switch the bleed source from an earlier to a later stage tapping. This allows the higher pressure and temperatures, required for the services, to be maintained. As an example, the Rolls Royce Tay engine has tappings at the 7th and 12th stages of the HP compressor and the changeover from one to the other occurs automatically when the engine speed falls below, or rises above 80%.

The bleed air, tapped from the compressor of the engine, is distributed to numerous services. The diagram below shows the air distribution on a typical twin-engined, 100-seat airliner. Of note are the additional inputs from the Auxiliary Power Unit (APU) and the ground supply unit, as well as the two power plants.


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Pneumatic ServicesFigure 12.8

The more complex system diagram overleaf only shows how the two tappings from theengine, at stages 7 and 12 are fed into the common manifold. From there it is used to anti-ice the intakes, pressurise the cabin, anti-ice the wing and tail, pressurise the hydraulic and water tanks, as well as cross-starting the other engine in flight, if required. It will also be seen that the Shut-off and Temperature Modulating Valve is the valve that switches between the bleed air stages 7 and 12, depending on the engine speed.


Typical Pneumatic SystemFigure 12.9

It will also be seen from the illustration on the previous page, that there are a number of shut-off valves used to activate and de-activate the various engine bleed air supply systems. The crew activates all these shut-off valves, from the flight deck, when required.

Due to the large demand for bleed air from some systems, inhibitions are applied at certain times such as take-off, to prevent a reduction in power, due to excessive demand, at critical times. There are also overpressure valves and temperature sensors, within the bleed air supply, that protect both the ducts and the services that they supply.


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15.13 STARTING AND IGNITION SYSTEMS

INTRODUCTION

Two main systems are required to ensure that a gas turbine engine will start satisfactorily. Firstly, provision must be made for the compressor and turbine to be rotated up to a speed at which adequate air passes into the combustion system to mix with the fuel from the fuel spray nozzles. Secondly, provision must be made for ignition of the air/fuel mixture in the combustion system.

During normal engine starting, the two systems must operate simultaneously. It must also be possible to motor the engine over without ignition for maintenance checks and to blow out residual fuel after a failed start. In addition, it must be possible to operate the ignition system for relighting the engine during flight.

The functioning of both systems is co-ordinated during a starting cycle and their operation is automatically controlled after the initiation of the cycle by an electrical circuit. A typical sequence might be as follows:

Start button pressed Ignition ‘ON’ HP Fuel ‘ON’ Light-Up Self Sustaining Starter Circuit ‘OFF’ Idle RPM

Methods of Starting

The starting procedure for all jet engines is basically the same, but can be achieved by various methods. The type and power source for the starter varies with engine and aircraft requirements. The power sources can be electrical, gas, air or hydraulic and each method has its merits.

The requirements for a military aircraft, for example, are totally different to those for a commercial airliner. The starter motor must, however, always produce a high torque and then transmit this torque to the engine in a smooth manner to accelerate it to self-sustaining speed.


13.1.1 ELECTRIC STARTING

The electric starter is usually a direct current, (D.C.), electric motor coupled to the engine through a reduction gear and ratchet mechanism, or clutch, which will automatically disengage once the engine is self-sustaining.

The electrical supply voltage can be progressively increased by the removal of resistances in the circuit as the engine increases in speed. The ignition system is also actuated and supplied at the same time as the start is initiated.

Once the engine is running, the starter supply is cancelled by the drop in supply current or by the action of a timer mechanism. Either way, the starter slows down and the clutch or ratchet mechanism ensures that the engine can accelerate free from the starter drive shaft.

The diagram below shows a simplified electric starter circuit. It contains most of the components found in many starter circuits such as master switch, start button and main starter relay. Overspeed relays usually disconnect the starter motor electrically, once the amount of current being drawn falls below a value which can only be reached if the engine is self-sustaining.

Electric Starting CircuitFigure 13.1


`13.1.2 AIR STARTING

Air starting is used on most commercial and some military jet engines. It ahs many advantages over other starting systems and is comparatively light, simple and economical to operate.

An air starter motor transmits power through a reduction gear and clutch to the starter output shaft, which is connected to the engine. A typical air starter is a basic air turbine that rotates at high RPM when HP air is passed through it from the on board Auxiliary Power Unit, (APU), a cross-feed from a running engine, or an external air supply.

The air supply, from whichever source, is controlled by an electrically operated control and pressure-reducing valve that is opened when an engine start is selected. It is automatically closed at a pre-determined starter speed. The clutch automatically disengages as the engine accelerates up to idling RPM and the rotation of the starter ceases.

The illustration below shows a typical air starting system, with a cut-away of the actual starter motor showing its rotor.

Starter system(left)and starter motorFigure 13.2


13.1.3 IMPINGEMENT AIR STARTERS

Some turbo-jet engines are not fitted with starter motors, but use air impingement onto the turbine blades as a means of rotating the engine. The air is obtained from an external source, or a running engine, and is directed on to the turbine blades.

13.1.4 GAS TURBINE STARTERS

On a few turbo-jet engines, a small self-contained gas turbine is used to start the engine. It is completely independent of the aircraft systems, excluding the electric starter. Once the small engine has started, its exhaust is directed, through nozzle guide vanes on to the turbine of the main engine, which will rotate through its own starting cycle, until it reaches self-sustaining speed.

13.1.5 HYDRAULIC STARTING

This form of starting is found, on occasions, fitted on to small gas turbine engines. In most applications, one of the engine mounted hydraulic pumps is utilised and is known as a combined pump/starter. This unit is coupled to the engine through the accessory gearbox and a reduction gearing. The hydraulic power, which will drive the unit in its ‘starter mode’ can come from external sources or on-board accumulators.

Once the starter has powered the gas turbine engine to self-sustaining speed, the unit changes from being a starter and becomes a normal hydraulic pump. In this form it acts as a normal pump throughout the remainder of the flight.


`13.2 HIGH ENERGY IGNITION SYSTEMS

High-energy (HE) ignition is used for starting all jet engines and, excluding APU’s, all have dual systems fitted. Each system has an ignition unit connected to its own Ignitor plug, the two plugs being fitted in different positions, (or Combustors), in the engine.

Each HE ignition unit receives a low voltage supply, controlled by the starting system circuit, from the aircraft’s electrical system. The electrical energy is stored in the unit until, at a pre-determined value, the energy is dissipated as a high voltage, high current discharge across the plug.

These ignition units are rated in ‘Joules’. Each Joule is equal to one Watt/ Second, with a value of 12 Joules being typical for a high output and 3 to 6 Joules for a low output. A high output would be required for re-lighting at altitude and certain ground starts, whilst a low output would only be required during continuous operation in icing or wet weather, giving longer Ignitor and ignition unit life.

To be able to operate at both levels, combined systems, giving high and low level outputs are most popular. Such a system would consist of one unit emitting a high output to one Ignitor plug and a second unit giving a low output to a second Ignitor plug. Some Ignitor units have been manufactured which contain both high and low outputs, which means that two igniters can be operating at either level depending on the conditions and the relevant cockpit selection.

A typical, simple ignition system is illustrated below and shows how the inputs are modified, through several stages, to give a high voltage, direct current to the HT terminal of the Ignitor.


LETHAL WARNING

The electrical energy stored in the HE ignition unit, (HEIU), is potentially lethal. Before handling the component, the associated circuit breaker should be tripped or the relevant fuse removed. Allow at least one minute to elapse, after isolating the unit, before touching the unit itself, the HT lead or the Ignitor plug.

Typical Ignition SystemFigure 13.3

The Ignitor plugs operate in the same way as sparking plugs, except that they are only required to start the engines, they are then switched off until the next start. There are two basic types of Ignitor plug, the air gap type and the surface discharge type. The air gap type require a potential difference in the region of 20,000 volts, whist the surface discharge type only requires a voltage in the region of 2,000 volts.

As igniters are used for both low-tension D.C. systems and high- tension A.C. systems and are NOT interchangeable, care must be taken to use the correct item, as recommended by the manufacturer, in their overhaul/maintenance manuals.

The normal spark rate of a typical ignition system is between 60 and 100 sparks per minute. Periodic replacement of the Ignitor plug is necessary due to the progressive erosion of the electrodes caused by each discharge.


`The two illustrations below show one unit of a dual ignition system (left) and a surface discharge Ignitor (right).

Ignition System(Left)And Ignitor PlugFigure 13.4

RELIGHTING

The jet engine requires the facility for relighting should the flame in the combustion chamber become extinguished during flight. This ‘relighting’ can only be safely accomplished if the aircraft is at the correct speed and below a certain altitude. The chart, (below) illustrates the relighting envelope for a specific aircraft. If the aircraft is too slow or too fast, or if it is above about 25,000 feet, there is little chance for the engine to relight. Within this envelope the airflow will rotate the compressor at a speed satisfactory for relighting.


Relighting EnvelopeFigure 13.5


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MAINTENANCE SAFETY

It has already been mentioned that the HEIUs can deliver a fatal shock if they are handled whilst still live. It is also possible to be shocked if either the Ignitor leads or the igniters themselves are handled before 1 minute has elapsed after removing all power from the system.

DO NOT depend on just the starter master switch being placed into the ‘OFF’ position as it is possible someone may switch it to ‘ON’ whilst you are working some way from the cockpit, on aft mounted engines for example. At least pull AND LABEL AS ‘INOP’, any circuit breakers applicable to the HEIUs. Also, disconnect the Low-Tension connectors on the Ignitor box itself to be doubly sure.

Exercise great care when handling some types of ignition transformer units if they are damaged. They can contain radioactive material on their air gap points.

Some Ignitor plugs are manufactured from exotic materials, which require special disposal arrangements. Check to see whether the items you are removing for disposal at life expiry are of this type.




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15.14 ENGINE INDICATING SYSTEMS

All engine parameters require monitoring and, therefore, instrumentation is provided to inform the pilot of the correct functioning of the various engine systems and to warn of any impending failure.

Should any of the automatic controls fitted to the engine fail, the engine can be manually controlled, by the pilot, who can select the required thrust setting by monitoring the instruments to maintain the engine within the relevant operating limitations.

The multitude of dials and gauges on the instrument panels may be replaced by one or more cathode ray tubes, (CRTs), to display engine parameters. These screens are often integrated into a complete set of flight and engine instrumentation displays.

As an example, the first illustration below shows a typical analogue engine parameter display for a twin engined aircraft. There are displays for the following engine parameters:

RPMT.G.T.E.P.R.OIL TEMPERATUREOIL PRESSUREVIBRATIONFUEL TEMPERATUREFUEL QUANTITY FUEL FLOW

Engine IndicationsFigure 14.1


Whilst this display is from a typical twin-jet aircraft, the addition of one or two further engines would complicate the display by adding at least 12 more instruments to the display. Electronic indicating systems, however, consolidate engine indications, system monitoring and crew alerting functions on to one or more CRTs mounted on the instrument panel.

The following illustrations show two forms of engine instrumentation. One form duplicates the analogue instruments so that they display the readings much as the older analogue instruments would have done. This example displays the following parameters:

E.P.R.ENGINE SPEED (N1)E.G.T.OIL PRESSUREOIL TEMPERATUREOIL QUANTITYVIBRATIONENGINE SPEED (N2)ENGINE SPEED (N3)FUEL FLOW

CRT Analogue DisplaysFigure 14.2


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The other example,( below)displays its parameters in the form of ‘ribbons’ or ‘tapes’, which climb up the display as the quantities being represented increase. On this display the parameters shown are as follows:

E.P.R.T.G.T.ENGINE SPEED (N1)ENGINE SPEED (N2)FUEL FLOWOIL PRESSUREOIL TEMPERATUREOIL QUANTITYVIBRATIONFUEL TEMPERATUREFUEL USED

Tape DisplaysFigure 14.3


TURBINE GAS TEMPERATURE MEASUREMENT

The temperature of the exhaust gasses is always indicated, thus ensuring that the temperature of the turbine assembly can be checked at any specific operating condition. In addition, an automatic gas temperature control system can be fitted to some gas turbine engines, to ensure that the maximum gas temperature is not exceeded.

The turbine gas temperature, (T.G.T.), can sometimes be referred to as exhaust gas temperature, (E.G.T.), or jet pipe temperature, (J.P.T.). It is a critical variable of engine operation and it is essential to provide an indication of this temperature.

Ideally, turbine entry temperature, (T.E.T.), should be measured; however, because of the high temperatures involved, this is not practical. As the temperature drop across the turbine varies in a known manner, the temperature at the outlet from the turbine is usually measured by suitably positioned thermocouples. The temperature can also be measured at an intermediate stage of the turbine assembly.

The thermocouple probes use to transmit the temperature signal to the indicator consist of wires of dissimilar metals that are joined together inside a metal guard tube. Transfer holes in the tube allow the exhaust gasses to flow across the junction. These wires are usually nickel-chromium and nickel-aluminium alloys. The probes are connected in parallel and their output is transmitted to a milli-voltmeter calibrated to read in degrees centigrade, (0C). For fine adjustments, there usually is a trimmer resistor in the circuit.

A basic system diagram is illustrated below, showing just a single temperature probe.

Basic Thermocouple SystemFigure 14.4


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ENGINE THRUST INDICATION

The thrust of an engine is normally shown on an instrument system that will be one of two types. The first type measures the turbine discharge or jet-pipe pressure and the second, known as an engine pressure ratio, (E.P.R.) gauge, measures the ratio of two (normally) or three parameters. (The straight jet pipe pressure measurement system is rarely used these days, due to the E.P.R. system being much more accurate).

When E.P.R. is measured, the ratio is usually that of jet-pipe pressure to compressor inlet pressure. On a fan engine, the ratio can be either between an integrated turbine discharge/fan outlet pressure to compressor inlet pressure or, as on the Rolls Royce Tay fan engine, between the by-pass duct pressure, (fan outlet), and the compressor inlet pressure.

In each of the above examples, it will be necessary to have a correction figure for the current ambient conditions. This figure will form a standard against which the jet pipe or duct pressures can be compared. This comparison is normally carried out automatically.

To measure the compressor inlet pressure a pitot type tube is normally used. The pressure that is read by this is either connected directly to the indicator or to a pressure transmitter that is electrically connected to the transmitter. Measurements taken in the fan duct, or the jet-pipe itself, are also taken by probes that have to be very sturdy to stand the air and temperature loads exerted on them.

The display shown to the pilot can be in either analogue or digital form. It will simply be a ratio above 1.0 and normally below 2.0. The ratio for take-off on a specific day, can be obtained from either the flight manual or, on modern aircraft, from the flight management system, (FMS), computer. This ratio will be aimed for when the pilot moves the throttle levers forwards prior to taking off.


TORQUE INDICATION

Engine torque is used to indicate the power that is developed by a turbo-propeller engine and its indicator is known as a torquemeter. The engine torque, or turning moment is proportional to the horse-power and is transmitted through the propeller reduction gear.

A basic torquemeter system is shown in the illustration overleaf. In this system, theaxial thrust produced by the helical gears is opposed by the oil pressure acting on the pistons. This pressure which is required to resist the axial thrust is transmitted to the indicator.

In addition to providing an indication of engine power, the torquemeter system may also be used to automatically actuate the propeller feathering system, if the power fails at a critical time. It can also actuate a water injection system when hot or high.

Another method of measuring torque is to measure the ‘twist’ on the main drive shaft, in the reduction gear-box, whilst the engine is producing high power. A phonic wheel is installed at each end of the shaft, each wheel having a sensor pick-up. When the shaft is rotating without carrying power, the two phonic wheel outputs are in synchronisation. When the shaft is carrying high torque, it twists, putting the outputs out-of-phase. This phase difference is measured electronically and displayed as torque, usually in foot/pounds.

OIL PRESSURE AND TEMPERATURE

It is essential for correct and safe operation of the engine that accurate indication is obtained of both the temperature and the pressure of the oil in the system.

OIL TEMPERATURE

This is sensed by a detector fitted in the oil system. On dry-sump engines, (those with separate oil tanks), the temperature sensor is often located in a special fitting between the oil tank and the pump. On wet-sump engines, (those whose oil is contained within the lower section of the engine itself, together with the pumps etc.), it is usually installed inside the oil screen immediately after the pump.

Early systems measure the oil temperature mechanically, by measuring the pressure of a gas sealed inside a bulb located in the oil stream. The pressure of the gas varies in proportion to its temperature and is displayed on the flight deck as oil temperature.

On modern systems, an electrical temperature sensitive element is fitted. A change in temperature causes a change in the resistance value and, consequently, a corresponding change in the current flow at the indicator. The indicator pointer is deflected by an amount equivalent to the temperature change. This is displayed on the gauge in degrees centigrade.


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OIL PRESSURE

Oil pressure is usually measured at the outlet of the engine driven pump and is indicated on the instrument panel. It will give warning of impending engine failure caused by a depleted oil supply, pump failure, damaged bearings or a ruptured oil supply line, all of which will be indicated by a fluctuation or fall in oil pressure.

Again, early designs used a Bourdon Tube mechanism that measures the difference between the oil pressure and the ambient air pressure. The rising oil pressure tends to straighten the flattened and curved tube in the gauge. This mechanical movement was transmitted to the needle, which displayed the oil pressure in pounds per square inch.

Modern designs use electrical systems to indicate to the flight deck. Simple systems use a ‘flag’ method that simply indicates if the pressure is high, normal or low. Normally though, the pressure sensed by a transducer in the oil supply line is transmitted, through wires, to the cockpit gauge, with display changes being in proportion to the oil pressure changes. The display may be in pounds per square inch or bars.

In addition to a pressure gauge operated by a transmitter, an oil low-pressure warning switch may be provided to indicate that a minimum pressure is available for continued safe running of the engine. This switch can be set to operate a warning ‘caption’ on the instrument panel and/or an audible warning for the crew.

This type of warning allows the crew to immediately shut an engine down in flight, allowing the fault to be found later, instead of the engine continuing to run dry and being damaged, perhaps permanently.


FUEL TEMPERATURE AND PRESSURE

The temperature and pressure of the low-pressure fuel supply are electrically transmitted to their respective indicators. These show if the low-pressure system is providing an adequate supply of fuel without cavitation and at a temperature to suit the operating conditions.

Basically, a low fuel pressure could indicate the failure of a booster pump or other flow problems. The gauge may also be supported by a warning light, (caption), or even an audio warning. There are nearly always two booster pumps fitted in each tank that cater for the pump failure situation.

A low fuel temperature indication might require the flight crew to actuate the fuel heating facility to prevent the filter becoming clogged by ice crystals. This heat is fed through a heat exchanger and can come from air, bled from the engine or oil from the lubrication system.

On some engines, a fuel differential pressure switch, fitted to the low-pressure fuel filter, senses the pressure differential across the filter element. The switch is connected to a warning caption or an audio warning to indicate a partial filter blockage, with the risk of fuel starvation.

FUEL FLOW

Although the amount of fuel consumed during a given flight may vary slightly between engines of the same type, fuel flow does provide a useful indication of the satisfactory operation of the engine and the amount of fuel being consumed during the flight.

It is also useful to sum the fuel flow, over time, to give a ‘total fuel used’ figure for each engine, (or for the aircraft), so that the amount used at differing times of the flight, such as climb, cruise, etc., can be logged.

A typical system consists of a fuel flow transmitter, which is usually some form of turbine or impeller, fitted in the low-pressure fuel system and an indicator. The indicator shows the rate of fuel flow, in whichever units the aircraft tanks are calibrated, (gallons, pounds, kilograms or U.S. gallons per hour).

The transmitter measures the fuel flow electrically and an associated electronic unit gives a signal to the indicator, proportional to the fuel flow.


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ENGINE SPEED

All powered aircraft have a display that shows the engine speed. On multiple spool engines, the high-pressure spool speed is always shown and, in most cases the intermediate, (if fitted), and the low-pressure spool speeds are also displayed.

The engine speed indication is usually electrically transmitted from a small generator, driven by the engine, to an indicator, which displays the actual revolutions per minute, (RPM) or a percentage of maximum engine speed. Whilst the engine speed is sometimes used to assess the engine thrust, it does not give an absolute and accurate figure, due to the atmospheric circumstances of the day. (Pressure and temperature), which can change the thrust available at a specific time.

Engine Speed Indicator And TransmitterFigure 14.5

The engine speed transmitter is a simple 3-phase generator in which the frequency of the output is read by the indicator and displayed either as RPM or as a percentage between 0% and 100%.


Where there is no provision for driving a generator, a speed probe in conjunction with a phonic wheel may be used, (right). This will induce an electric current that is amplified and then transmitted to an indicator. This method can be used to provide an indication of RPM without the need for a separately driven generator, with its associated drives; thus reducing the number of components and moving parts in the engine.

Phonic Wheel SystemFigure 14.6

Some speed probe/phonic wheel assemblies have the additional facility of being able to illuminate a light on the flight deck when the relevant shaft begins to turn. This is used to assist the pilot by telling him when to open the HP fuel cock during the start cycle. It is only in operation during the starting cycle.


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VIBRATION MEASUREMENT AND INDICATION

A turbo-jet engine has an extremely low vibration level, especially when compared to reciprocating engines, even fully balanced ones. Because of this, a vibration change or the appearance of a new vibration due to an impending or partial failure, may pass without being noticed. Many engines are therefore fitted with vibration indicators that continually monitor the vibration level(s) of the engine. The cockpit indicator is usually a milliammeter that receives signals through an amplifier from engine mounted transmitters.

A vibration transmitter is mounted on the engine casing and electrically connected to the amplifier and the indicator. The vibration sensing element is usually an electro-magnetic transducer that converts the rate of vibration into electrical signals which cause the indicator to show the amplitude of the vibrations being sensed, usually in inches per second, (I.P.S.).

Because of the rarity of an excessive engine vibration reading, there will be a warning light/caption or audio warning to warn the flight crew that there is an abnormal reading on the indicator, enabling them to shut the engine down and so reduce the risk of damage.

On advanced, three-spool engines, there is an additional facility that allows the vibrations from the three spools to be differentiated and, therefore, each spool can be monitored separately. It is also possible for the flight crew to select a specific area such as the accessory gearbox, for vibration monitoring.




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15.15 POWER AUGMENTATION SYSTEMS (level B1 only)

The maximum power output of a gas turbine engine depends to a large extent upon the density, or mass, of air passing through the engine. There is, therefore, a reduction in thrust, (or shaft horsepower on turbo-propeller engines), as the atmospheric pressure decreases with altitude and/or the ambient air temperature increases.

Under these conditions of low power, the power output can be either restored or boosted for take-off, depending on the engine design, by cooling the intake airflow with water or water-methanol mixture. When methanol is added to the water, it gives both anti-freezing properties and also provides an additional source of fuel.

The graphs below show how a gas-turbine engine can have its thrust restored when the ambient temperature reaches high levels, (left) and how a turbo-propeller engine can have its power either restored or boosted at higher temperatures.

Power restored (left )and restored/boostedFigure 15.1


INTAKE INJECTION

Coolant is normally injected directly into the compressor inlet, although some have it injected into the combustion chambers themselves. This is more efficient because it gives both a more even distribution and a greater quantity of the liquid into the engine. When water/water methanol mixture is sprayed into the compressor inlet, the temperature of the compressor inlet air is reduced and consequently the air density and, therefore, the thrust are increased. If water only was injected, it would reduce the turbine inlet temperature. With the addition of methanol, however, the turbine inlet temperature is restored by the burning of the methanol in the combustion chamber. Thus, the power is restored without having to adjust the fuel flow.

COMBUSTION CHAMBER INJECTION

The injection of coolant into the combustion chamber inlet increases the mass flow through the turbine, relative to that through the compressor. The pressure and temperature drop across the turbine is thus reduced, which results in an increased jet pipe pressure, giving additional thrust. The fuel system is now able to schedule more fuel than before the water was injected, giving even more thrust. If water methanol is injected, there is no need for extra fuel to be scheduled as the burning methanol provides the extra power.

CARE MUST BE TAKEN WHEN REPLENISHING THESE SYSTEMS –

WATER/METHANOL IN A WATER SYSTEM CAN DESTROY AN ENGINE.


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15.16 TURBO-PROPELLER ENGINES

Whenever a gas turbine is used to power a propeller, by whatever means, it us normally referred to as a turbopropeller engine. Other terms used are turboprop and, rarely, propjet. The exhaust gasses from the basic part of the turbojet, often called the gas generator, are used to rotate an additional turbine that drives the propeller through a speed-reducing gearbox.

In some turbo-props, additional turbine stages are incorporated into the existing turbine assembly that rotates the compressor. The additional power that is removed from the gas stream achieves two things; firstly it reduces the thrust from the exhaust of the engine and, secondly, the energy removed from the exhaust is used to drive the propeller reduction gearing, directly from the compressor drive shaft. Engines of this type are known as direct-drive turbopropeller engines.

More common these days are free-turbine turbopropeller engines. This design has the additional turbine stage independent of the compressor drive turbines, which is free to rotate by itself in the engine exhaust gas stream. The shaft on which the free turbine is mounted drives the propeller through the propeller reduction gearbox.

The simplified illustrations below show (top) a turboprop with the propeller driven directly from the compressor shaft through reduction gears and (bottom) a turboprop with the propeller driven by a free turbine, also through reduction gearing.

Direct Coupled(Upper)And Free Turbine Turbo-PropFigure 16.1


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Another configuration of the free turbine turbopropeller engine has a rather unconventional rear-to-front air and gas flow direction. This configuration provides greater flexibility in the design of nacelle installations. The space behind the engine, which is normally occupied by the exhaust jet pipe, can be used for wheel wells, fuel tanks or a baggage compartment.

Another advantage of this ‘reverse design’ type of engine is the ease of hot end replacement, which occurs when the turbine of the engine, (which is the most thermally stressed part), has to be replaced partway through the overhaul life of the engine. This often requires an engine change and even, occasionally, an overhaul at the factory.

With this design, the turbine is at the front of the engine and can be accessed by simply removing both the propeller and the front engine casing. The engine remains fitted in the airframe.

A typical engine of this type is the Pratt and Whitney PT6. A cutaway of one model of this engine is shown below, where it can also be seen that the compressor is one of the axial/centrifugal types popular with these small turboprop/turboshaft engines.

Reverse flow turbo-prop engineFigure 16.2

This type of engine will also be found installed in helicopters, marine craft and other installations, where it is known as a turboshaft engine. This will be covered in the next section.


REDUCTION GEARBOXES

Due to the fact that propellers are most efficient at relatively slow rotational speeds, they will usually be found rotating at a speed of between 1,500 to 2,500 RPM, the shafts of turbopropeller engines can be rotating at many times more than this. As an example, the PT6 engine mentioned above has a power turbine rotational speed of about 30,000 RPM, whilst the propeller rotates at only 1,700 RPM

This means that the shaft speed must be reduced by a factor of about 1:0.057. This is achieved by the use of an epicyclic gearbox which ‘steps down’ the speed of the power shaft in two stages. This reduction will be explained later.

The illustration below shows where the reduction gearbox, which will be described later, is located relative to the other parts of a typical turbopropeller engine.

Reduction GearboxFigure 16.3

The location of the gearbox at the front, immediately behind the propeller, is by far the most common location and the explanation which follows will cover a single stage reduction between an input shaft to an output propeller shaft


`The high-speed pinion rotates at the same speed as the LP (or power turbine), shaft and meshes with the large planet wheels. The large and small planet wheels are on common shafts and, hence, rotate together, they are also all mounted on the planet carrier which is rigidly fixed to the propeller shaft.

The annulus, or ring, gear is the only fixed item in the whole assembly, hence the small planet gears, which mesh with it, pull themselves around as they rotate.

Epicyclic reduction gearsFigure 16.4


INTEGRATED CONTROLS

A motor car has a simple accelerator to increase and decrease the speed of its engine, a turbojet has a single throttle lever to carry out the same operations. The turbopropeller engine, however, requires either a single lever with two different but integrated controls, or two levers that have to be operated separately to control different functions.The reasons for this is that the propeller has to be given instructions separate to the engine itself. For example, the engine might need to be run on the ground with nil thrust at some times and producing power at others. Normally this involves the lever controlling the speed of the propeller, or that of the free-turbine and it usually governs the pitch angle of the propeller blades.

In a free-turbine engine such as the PT6 mentioned earlier, the engine acts as a gas generator furnishing high velocity gasses to drive the free turbine. It is normally controlled by a power control lever and a propeller control lever. In addition a third, start control lever, (also sometimes known as the fuel condition lever), is used to select a ‘high’ or ‘low’ speed range for engine operations and to cut off fuel to shut the engine down.

The power control lever modulates engine output from maximum take-off power to full reverse power. It also controls the propeller blade angle when in reverse and in what is known as the “Beta” range, (approach and landing).

The propeller control lever is in control of the propeller control unit, (P.C.U.), above “Beta”, which is generally the normal flight regime. Here it operates as a normal single lever ‘throttle’, when the thrust being produced is a function of lever angle and the propeller RPM is governed to a specific speed, again for each lever angle.

The illustrations overleaf show how complex the three different lever positions are and, also how difficult it is for the flight crew to know which lever has to be in which position at different times.


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Power lever anglesFigure 16.5

Some engines can have a different control layout when installed in different airframes. For example, the Pratt and Whitney 120 series engines can be installed with single or double lever control depending whether they are installed into the Fokker 50 or the ATR 72 aircraft. The illustration, (below)shows the ATR72 layout, which has power levers, (left) and condition levers, (right).

Pedestal LeversFigure 16.6


It can be seen that whilst the condition levers control the fuel shut-off and feathering controls, they also set the engine to automatic engine power and a maximum 100% override position.

The power levers are used to change the engine thrust when taxiing, select ground and flight idle, actuate reverse thrust and the automatic thrust position.

Finally, the installation of the early Rolls Royce Dart engine in the 1950s airframe design, the Fokker 27, has basically a single lever operation. The illustration below shows how the central pair of levers control ‘RPM’, e.g. propeller pitch angle and fuel control, whilst the outer pair of levers, (H.P.C.), have no input to the power of the engines, purely fuel shut-off and propeller feathering operations.

Single Lever ControlsFigure 16.7


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OVERSPEED PROTECTION

A propeller that overspeeds, even by the small amount of, say 5 or 10% increases the centrifugal forces on the hub by a huge amount. This could cause the blades to separate from the hub with catastrophic results to the aircraft.

A gas turbine engine has its own fuel control system, which maintains the engine within its operating speed range. With a turbopropeller engine it is normally the propeller which acts as a governor by increasing or decreasing its pitch angle to add or remove the loading on the rotating parts of the engine.

If a turbopropeller overspeeds, it is usually due to the fact that the propeller controls have allowed the pitch angle of the propeller to decrease, so that the reduction of load on the engine has caused it to Overspeed. This reduction of pitch is as a result of aerodynamic and centrifugal forces acting on the rotating propeller which will be covered later.

If the reduction of the propeller pitch has been caused by failure of the propeller control unit, there may be a back-up method, built in to the control system, to drive the propeller back to a coarser angle, thereby slowing it down to a safe value. These back-up systems usually involve the use of centrifugal governors which sense the slightest Overspeed.

If the propeller control system is damaged or it cannot drive the propeller to a safe, coarser, blade angle, the fuel control of the engine reduces the flow of fuel to the engine, effectively acting as if the pilot had retarded the throttle. This should bring the hub loading within a safe value.


As an example, the system shown is that fitted to the Pratt & Whitney 124 engines on the ATR72 aircraft, which has a combined hydraulic/pneumatic Overspeed protection. If the propeller overspeeds above 102.5% NP, (NP = propeller speed), The flyweights move outwards, opening the pilot valve and allowing metered oil pressure to drive the propeller towards coarse.

In the event that the above system fails to operate, (propeller continues to accelerate), the air bleed orifice opens at a slightly higher NP. This bleed biases the fuel control system, (H.M.U.) to decrease the fuel flow, reducing the engine speed.

Propeller Control UnitFigure 16.8


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Whilst the system previously described is rather complex, the engine is of a modern, ’free power turbine’ design and has to have sophisticated protective measures fitted. By comparison, the Overspeed protection installed on the Rolls Royce Dart, a ‘direct coupled’ drive engine designed in the 1940s, is a relatively simple system.

The pump case pressure is fed with fuel from radial tappings in the rotating pump assembly. If the engine overspeeds, the fuel is ‘centrifuged’ into the pump case at a higher pressure. This pressure is fed to a diaphragm in the Overspeed governor, which spills the servo pressure and reduces the fuel supply to the engine. This limits the engine, which normally has a governed maximum of 15,000 RPM, to an Overspeed maximum of 16,400 RPM The illustration below shows the basic system showing how spilling the servo pressure reduces the pump output.

Overspeed protectionFigure 16.9

Apart from the protection mechanisms already mentioned, which have to react extremely fast to prevent accidents, there are a number of flight deck indications which may be in place of, or in addition to the automatic systems.

The simplest is the ‘red line’ on the tachometer, (revolution counter), or power, (percentage), instrument, which must not be exceeded at any time. If the aircraft has an electronic flight warning system, (F.W.S.) however, then warning lights, captions and audio warnings may be used to get the attention of the flight crew.


15.17 TURBO-SHAFT ENGINES

Gas turbine that delivers power through a shaft to operate something other than a propeller is referred to as a turboshaft engine. These are widely used in such industrial applications as electrical power generating plants and surface transportation systems. In most cases, the output shaft, (power take-off), is driven by its own power turbine which extracts about two thirds of the total output power from the gas generator.

In aviation, turboshaft engines are used, to power many modern helicopters of all sizes, from under 1 tonne up to a Russian giant of 100 tonnes maximum weight. To produce the power required for the previously mentioned helicopter applications, the turboshaft engine have been designed to produce from as little as 400 shaft horse power, (S.H.P.), up to 11,500 S.H.P. from each engine.

An illustration of how efficient modern turboshaft engines are, can be seen from the weight of the two, (extreme), cases mentioned. The Allison 250 produces the 400 S.H.P. mentioned with a total weight of just 158 lbs. (72 kg.), whilst one giant Lotarev D-136 engine, which produces that huge 11,500 S.H.P., weighs 2,300 lbs. (1,050 kg.), more than the weight of the smaller helicopter.

APPLICATIONS

In many installations, possibly reflecting the greater emphasis on safety at times such as over-water and over hostile lands, etc. they are installed in pairs, threes and in on extreme case in two pairs to make a total of four engines on one craft. Because of the need to drive the rotor systems, consisting of many different types of gearbox, the turboshaft engine has to be able to drive from a variety of different places.


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This illustration, taken from the Rolls Royce Gem brochure, shows how the designer could offer a power take-off from the front, back and side of the basic engine, to suit different aircraft designs. It can be seen that the engines could also be joined together by a combining gearbox to double the power output.

R.R. Gem InstallationsFigure 17.1


A few actual examples of turboshaft installations show how the engines can be located ahead or behind of the main transmission gearbox. The Westland Lynx has two Rolls Royce Gem engines mounted aft of the gearbox driving through couplings at the front face of the engines. It can be seen from the illustration below how the engine/gearbox unit is quite compact.

Turboshaft Installation(1)Figure 17.2


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Another twin-engined installation is that which can be found fitted to numerous Sikorsky and Westland helicopters. These are fitted ahead of the main gearbox, so that the output shaft and coupling projects from the rear of each engine. The location of all of the previously mentioned layouts permits very easy maintenance and engine changes due to the unobstructed access to the engines. The illustration is of the S-61N model which has two 1,400 S.H.P. turboshaft engines.


DE-RATING

A common method of being able to maintain normal flight in the event of losing one of a pair of engines is de-rating. The engines might be advertised as having, say, 1000 S.H.P. each but, in the event of a single failure the second, good engine, can be called upon to produce 1,500 S.H.P. for a limited time. This system of rating also allows the engines to be operated in ‘Hot and High’ situations without power loss.

ARRANGEMENTS

Finally, there are a few other installations on helicopters, using turboshaft engines, that show the flexibility in the way these engines can be mounted to suit the designer’s needs. The little Hughes 500 series, (upper illustration), has a small 400+ S.H.P. engine, installed at an angle, driving upwards at 450 to the main gearbox. The large E.H. 101 helicopter, however, has not only three engines, each of 2,000 S.H.P., installed above the decking and all feeding into the main gearbox, but the Auxiliary Power Unit is installed up there as well, (lower illustration).




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One major difference between turbopropeller and turboshaft engines is their operating regime. Whilst turbopropeller engines are called upon to vary their speed as the pilot varies the throttle position, turboshaft engines, as a general rule, operate at a constant speed. This is because helicopter rotors have to operate at an almost constant speed so that the engines, especially the non-free turbine designs, also have to remain at the same speed also.

The free turbine designs will also remain at a constant speed for most of the flight, only altering slightly up or down as the demand for climb or descent are initiated, requiring a little more or a little less lift from the rotors.

DRIVE SYSTEMS

Because of the way that both the engines themselves and their attached main gearbox are normally mounted to the airframe, there has to be a small amount of permitted movement between the components. This poses problems when they have to be connected together, with the need for both accurate alignment and the ability to transmit many thousands of brake horsepower. To this end there will be found between the engines and their respective gearboxes, some form of flexible drive couplings with a short, (rigid) drive shaft between them. This arrangement allows the two components to move small amounts, independently of each other.

One of the most common couplings is the ‘Thomas Coupling’ that consists of a number of thin, steel laminations which are attached to the engine output flange and either the gearbox flange or the drive shaft flange. They are attached at different radial points, e.g. at three places, 1200 apart and displaced 600 from the opposite flange.


In some cases, including the illustration below, there is an additional outer structural sleeve, which is part of the actual mechanical connection between the engine and the gearbox. This also will have a mechanism to allow slight, controlled movement between the two components. It will be seen that, in the example shown, the drive shaft has a Thomas Coupling on the left-hand connection and a rigid flange-to-flange connection on the right. The outer support tube, however, is rigidly mounted on the left and flexibly mounted on the right.

Coupling shaftFigure 17.5


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When there are two engines installed, driving into a common main gearbox, there are two main design choices. The engines can be connected direct to the gearbox at two separate points, with the drives being connected within the gearbox casing, through their respective freewheels. Alternately, the engines can drive into a combiner or coupling gearbox, separate from the main gearbox with two input connections and one, larger, output connection to the main gear box.

These boxes allow the main box to remain smaller, with the added advantage that either can be changed for fault rectification or at overhaul without disturbing the other.

As mentioned previously, there are a number of combining gearboxes used on many twin engined helicopters.

The first example,(below), called in this case a coupling gearbox, has the inputs low on the left hand side and the single output in the centre right of the box.

Coupling gearboxFigure 17.6

This box also contains mechanisms to allow separate connection and dis-connection of each input drive. Also, it has drives to a number of accessories such as pumps and generators. It also has its own self-contained oil system, with pump, filter etc.


The second example, seen below, is installed in one of the larger, 15 seat Sikorsky helicopters. It is also connected to the main gearbox via a short drive shaft, again allowing for movement between the components.

Transmission DetailsFigure 17.7

Because the engines in this class rotate at extremely high rates and the outputs need to be considerably less, a ‘train’ of reduction gears needs to be built into the system. (As an example, the Garrett TPE 331-11 has a gas generator shaft speed of 41,730 RPM, whilst its output shaft rotates at only 1,600 RPM).


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In some installations the reduction gearing is installed in the front of the engine itself, so that its shaft output is already at a useable rotational speed. The Gem engine mentioned earlier has the option of being built without a reduction gear installed; output is the same as the gas generator, 27,000 RPM or with the reduction gears installed with a 4.5:1 ratio, giving an output of about 6,000 RPM

The illustration below is of the, (optional), reduction gearbox fitted to the front of the R.R. Gem turboshaft engine. This will take the 27,000 RPM and, through the two-stage epicyclic gear train, reduce it to around 6,000 RPM At this speed, it can be directly connected to the main rotor gearbox, which will reduce it further to around 250 – 300 R.R.P.M. (Rotor RPM). This reduction mechanism allows the engine to be installed directly in a number of different situations such as powering marine craft, power-generating stations and pumping stations. This use of turbo shaft engines is very common and even engines as large as the Rolls Royce RB211 are used for such purposes, providing all the power for complete oil exploration and production platforms.

Reduction gearboxFigure 17.8


CONTROL SYSTEMS

The controls of turboshaft engines are, generally, the same as turbopropeller engines. There has to be some form of fuel shut-off, to permit starting and closing down of the engine and there has to be some form of speed control, often called the ‘speed select lever’.

The fuel shut-off is simply a high-pressure fuel cock that is opened at start-up of the engine and closed to stop the engine at the end of the flight. The speed selector might have a number of positions, including ‘ground idle’, ‘flight idle’ and ‘flight’ positions. Sometimes these points are indicated, by feel, using detents in the quadrant.

Because of the need for rotors, (and electrical generators, pumping units, etc.), to be driven at more or less constant speeds, there is often a governing mechanism, either mechanical or electronic which is engaged when the speed select levers are placed into the ‘flight’ or ‘constant speed’ position. Normally the pilot does not have to touch the control again until he retards the speed selector to the ‘ground idle’ position.

Once the speed select levers have been advanced to their normal operating position, the engines will be maintained at their governed speed regardless of the load being applied by their governor or fuel computer. This will simply increase the fuel flow when there is an application of load, (up to the maximum limits of speed and temperature), and reduce the flow under low demand circumstances.


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The illustration below shows the lever quadrant from a Sikorsky S-76. It can be seen that whilst the fuel levers, (outboard), have additional positions of ‘D/R’(Dry Run) and ‘PRIME’ peculiar to the Allison 250 engine, the power levers simply have ‘OFF’, ‘IDLE’ and ‘FLY’ positions.

Lever QuadrantFigure 17.9


The complete diagrammatic representation of the S-76 engine control system below, shows the pilot’s control inputs to the engines. They are simply the Off-Idle-Fly speed lever which control the gas generator, (N1), part of the engine, and the collective, (lift), lever which simply increases the fuel flow to the gas generator part of the engine when more lift is demanded. The ‘beeper’ is simply a fine trimming switch that allows the pilot to adjust the rotor/engine speeds to an exact figure.

System DiagramFigure 17.2


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15.18 AUXILIARY POWER UNITS (APU’s)

The Auxiliary Power Unit, (A.P.U.), is fitted in many aircraft to provide a supply of electrical, pneumatic and sometimes, hydraulic power when the aircraft is on the ground, parked, without the engines running. The A.P.U. can, in some circumstances, be used to provide limited services usually electrical power, in flight, during an emergency such as an engine failure on a twin-jet. The A.P.U. engine provides shaft power for electrical generation and hydraulic power, (if applicable), and bleed air for pneumatic supplies.

Pneumatic power is supplied for starting aircraft main engines, aircraft cabin air conditioning and pressurisation, air for A.P.U. oil cooler A.P.U. enclosure and fuel heating, (in some cases).

Electrical power can be provided independently of, or together with pneumatic power and, with some installations, the A.P.U. generators can provide power throughout the flight, in the case of a main generator having failed. (In this case, there may be a limit on the maximum altitude that the aircraft may fly). Note: Electrical power has priority at all times

For the majority of the notes that follow, reference will be made to the Allied Signal GTCP 36-150, which is a typical, medium sized A.P.U. fitted to a range of aircraft in the 100+ seat category.

DESCRIPTION

The basic A.P.U. consists of three main sections; power section, gearbox assembly and the controls and accessories. Operation of the engine is controlled by four systems; fuel, engine lubrication, electrical and pneumatic.

The rotating group of the power section consists of a single-stage centrifugal compressor with a single stage turbine, both mounted on a single shaft, which is mounted on ball and roller bearings.

The gearbox is attached to the inlet housing and converts the power section input, which is high speed / low torque to an output of low speed / high torque suitable for the accessories fitted to the gearbox. These include an alternator, fuel control unit, cooling fan and starter motor.


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An Electronic Control Unit, (E.C.U.), provides a fully automatic control system, including starting sequencing, normal speed control and Overspeed protection. Other protective shutdown circuits, monitored by the E.C.U., include the following:

Over Temperature Low Oil Pressure Over Current Loss of Thermocouple Loss of Speed Signal Hot Oil Temperature.

All are inhibited in flight except fire, loss of speed signal and too long start time.

CONSTUCTION

The construction of the power unit is simple and conventional. The simplified diagram below shows the gearbox / engine assembly with single stage compressor and turbines, fuel injection, single Ignitor and a drive to the accessory gearbox.

A.P.U ConstructionFigure 18.1


INSTALLATION

The A.P.U. is usually mounted in a non-pressurised area aft of the rear pressure bulkhead. Space available and the aircraft centre of gravity considerations usually dictate the exact mounting, which is why the Boeing 727 and the Bae ATP have the A.P.U. installed in the wing root and the Fokker 27 / 50 aircraft have it installed in the rear of the starboard engine nacelle.

The illustration below shows a typical rear fuselage installation of an A.P.U. in an Embraer EMB 120.

A.P.U InstallationFigure 18.2


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BASIC OPERATING CYCLE

The block diagram below shows the major sections of the A.P.U. in question and they are: The compressor, combustor, turbine and gearbox. The operating cycle, similar to a normal jet engine, consists of the compressor, when driven, drawing large volumes of ambient air and delivering it under pressure to the combustor. Fuel and ignition are added in the combustor and the resultant added energy delivers high velocity hot gasses to the turbine section. The turbine captures most of the energy of these high velocity gasses and converts it to the mechanical energy to drive both the compressor and the gearbox. The gearbox drives the required fuel and oil system components, as well as providing a means for mounting the airframe furnished components, such as generator; alternator, (a.c. generator); etc.

A.P.U Major SectionsFigure 18.3


A.P.U. operation is usually fully automatic from start to shut down, thus the flight compartment controls are kept to a minimum. Operational control is through an A.P.U. control panel like the one illustrated on the below.

Included on this panel are a three position rotary switch, an ‘available’ light and a ‘fault’ indicator. A more complex and powerful A.P.U., from a much larger aircraft, might, additionally, have an RPM percentage indicator as well as gauges showing E.G.T. and duct pressure. It might also have warning caption lights for ‘fire’, ‘low oil pressure’, ‘starting’,etc.

A.P.U. Control PanelFigure 18.4

The selector switch functions are as follows;

OFF Shuts down the A.P.U.

START Arms the start circuit, sprung loaded to the RUN/ON position

ON Normal position during running. Opens inlet doors prior to start.


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BLEED AIR CONTROL

The A.P.U. provides the required bleed air for aircraft use when this air is not available from the main engines.

Bleed air can be taken from the A.P.U. after the unit has accelerated to its normal operating speed and has stabilised for a period of time. Providing the ground/flight system is in the ‘ground’ position and the ‘bleed’ push switch on the control panel is pushed, air will be supplied to whichever services is selected.

The panel illustrated is from the Fokker 100 airliner. It shows both the ‘APU BLEED’ push switch, which selects the A.P.U. supply to the services required and the bleed manifold pressure on the display.

Bleed Air Control PanelFigure 18.5

The push switch will show one of three indications:

BLANK System selected and operating

OFF (White) System switched off

FAULT (Amber) A fault has occurred and the A.P.U. has shut itself down.


There is a similar system for using the electrical generating unit. There is usually a similar push-button switch on the electrical supply panel with ON (Blank) - OFF (White) – FAULT (Amber) selections. Pushing the button whilst the A.P.U. is running will supply any of the aircraft electrical services both on the ground and in flight. It is fairly common for the ac. Generator, fitted to the A.P.U., to be identical to that fitted to the engines. This ensures that the A.P.U. generator can take over the generating work for either of the engine-driven generators in the event of a total failure.

MAINTENENCE

There is little line maintenance carried out on the A.P.U., apart from a daily check of the oil level, (with a replenishment if required); the ‘filter by-pass’ indicators and an overall check for signs of damage or distress. The only other maintenance that might be carried out on a unit, whilst still installed in the aircraft, could include an oil and oil filter change or a check of the magnetic drain plug, looking for signs of metallic debris.

There are a large number of Line Replaceable Units, (LRU), that can be replaced, mostly in situ, if a fault indicator has shown that they require replacement. Whilst the purpose of some units shown have not been mentioned in this chapter, the thermocouple, ignition unit, the electronic control unit, fuel control unit, etc should be self-explanatory.


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Line Replaceable UnitsFigure 18.6


PROTECTIVE SYSTEMS

Because the A.P.U. operates automatically once it has stabilised at its normal 100% running speed, which can be around 60,000 RPM, there is little if anything that an engineer needs to do with regards to checking the condition of the unit.

As everything is under the control of the Electronic Control Unit, (E.C.U.), all faults come through this unit, which then either notifies the flight deck crew, (if it is an in- flight notifiable fault), or keeps it in its internal memory for later interrogation.

The E.C.U. can be regarded as the ‘brain’ of the control system and will usually be found in an avionics bay or some similar location with conditioned air. It receives its input signals from a speed sensor and a thermocouple, both located in the A.P.U. These allow the E.C.U. to control the actuation of the start, ignition, fuel, pneumatic and protective shutdown circuits.

The unit also provides electronic control of all the parameters required to start, accelerate, and safely operate the A.P.U. during ‘start’, ‘acceleration’, ‘idle’ and ‘load’ conditions.

Different A.P.U.s have different methods and parameters when giving warnings. In fact, in some cases, the unit will close itself down if the risk is high and, it will have a different close down schedule depending whether the aircraft is in flight or on the ground.

For example, in the GTCP 36-150, a low oil pressure will be an alert if the aircraft is in flight and a shutdown if the aircraft is on the ground.

Finally, the E.C.U. has a set of three ‘doll’s eye’ indicators which will indicate to the engineer what fault occurred, no matter if this was an airborne alert or an on ground shutdown. The code for the trouble shooting of the faults is shown in the table below.


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BITE NUMBER DECODED BITE INFORMATION

1 2 3

RESET O OVERSPEED O OVERTEMPERATURE O O LOW OIL PRESSUREO OVERCURRENTO O LOSS OF THERMOCOUPLEO O LOSS OF SPEED SIGNALO O O HOT OIL TEMPERATURE

BITE Code InformationFigure 18.7

To summarise, if a fault is not considered immediately threatening to both the aircraft and its passengers, the E.C.U. will generate an alert to the flight deck crew and leave the decision up to them, whether to shut the unit down or to leave it running. This allows them to continue with the unit output, (electrical power), during flight, if there is a more urgent need for the output than for the continued health of the A.P.U.

If the aircraft is on the ramp, and the A.P.U. is producing both an air supply and electrical power, any fault will result in the unit shutting itself down and indicating what the fault is, via the doll’s eyes. This is because, at times, the aircraft is not manned by either flight deck crew or ground crew, so the decision to shut the unit down is left to the E.C.U.

In the case of a fire, not only does the A.P.U. shut itself down, but it can fire its own fire extinguisher bottle, after all the vent doors have been closed, as well. The unit will then sound an alert outside of the aircraft to attract the attention of the ground engineers.




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15.19 POWERPLANT INSTALLATION

When a gas turbine engine is installed in an aircraft it usually requires a number of accessories to be attached to it. It also requires a number of connections, some of them ‘quick release’ type, to connect the engine to the many aircraft systems. The engine, jet pipe, accessories and, in some installations, a thrust reverser, must be suitably cowled. An intake with minimum losses must also form part of the powerplant installation.

The powerplant location and aircraft configuration are of an integrated design the form of which will be dependant upon the duties the aircraft has to perform. Turbo-jet engine installations can be in the form of pods that are attached to the wings by pylons or pods attached to the sides of the rear fuselage by short stub wings. They can also have a combination of rear fuselage and tail mounted power plants. Combinations of the preceding layouts can also be found on some aircraft.

Turbo-propeller engines are normally limited to installation on the wings or the nose of the aircraft, whilst turbo-shaft engines, almost exclusively installed in helicopters, can be found buried within the fuselage, installed above the cabin structure and, on some larger machines in pods, external to the fuselage structure.

The following illustration shows some typical installations, both of turbojet, turbo-propeller and turboshaft powered aircraft.

Powerplant InstallationsFigure 19.1


ENGINE MOUNTINGS

The position of the powerplant must not affect the efficiency of the air intake and the exhaust gasses must be discharged clear of the airframe and its control surfaces. All installations must produce the lowest drag possible. Installations are always numbered from left to right when viewed from the rear of the aircraft.

The engine is mounted in the aircraft in a manner that allows the thrust forces developed by the engine, (or propeller), to be transmitted to the aircraft main structure, in addition to supporting the engine weight and carrying any flight loads.

Because of the wide variations in the temperature of the engine casings, the engine is mounted so that the casings can expand freely in all directions. Turbojet engines are usually either side mounted or underslung as illustrated previously and below, whilst turbopropeller engines are usually mounted on a tubular framework, often called the engine mounting.

Engine MountingsFigure 19.2


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JET PIPE

The jet pipe is usually attached to the rear of the engine and supported by the engine mountings. On longer jet pipes, rollers on each side of the pipe, run in rails attached to the sides of the jet pipe bay.

ACCESSORIES

An aircraft powerplant installation generally includes a number of accessories that are either electrically operated, mechanically driven or driven by high pressure air. Electrically operated accessories might include actuators, amplifiers, control valves and solenoids.

Mechanically driven units however, might include generators, constant speed drive units, hydraulic pumps, oil pumps, fuel pumps, engine speed signalling components, measuring and governing units.

Air driven accessories are generally driven through bleed air tapped from the engine compressor. They can include the air starter, using the other engine(s) or the A.P.U., and possibly the thrust reverser and water injection pump. The air conditioning and pressurisation air will also be taken from compressor tappings in most cases. The total air drawn from the compressor must be a small percentage of the total airflow to avoid loss of power and higher fuel consumption


FIREWALLS

The firewalls are the metal dividing partitions that are made of fireproof material, often stainless steel, and which divide the complete engine bay into smaller compartments. These smaller compartments are usually to assist both in fire detection and extinguishing. In the majority of installations, each bay will have its own detection system and fire-extinguishing bottle.

The other very important purpose of the firewalls, (or fireproof bulkheads), is to restrict the spread of fire to the bay in which it originated, making extinguishing the fire much easier. The firewalls might divide the bay into different areas where a fire is most likely to occur, such as ones containing heat and combustible materials or fluids.

A typical division of an engine bay might be:Engine power section – includes burners, turbine and jet pipe.Engine compressor and accessory section OrComplete powerplant compartments, in which NO isolation exists between the other two sections mentioned earlier. A twin engined example is shown below.

Single Engine Bay FirewallsFigure 19.3


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COWLINGS

Access to an engine mounted in the wing or fuselage is by hinged doors; on pod and turbopropeller installations, the main cowlings are hinged. Access for minor servicing is by small detachable or hinged panels with most fasteners being of the ‘quick- release’ type.

A turbopropeller engine, or a turbojet engine mounted in a pod, is usually far more accessible than a ‘buried’ engine because of the larger area of hinged cowling that can be provided. The illustrations below, show the clear access to the engine afforded by the ‘petal cowlings’ (left) of a turbopropeller installation, and, (right), the ‘clamshell doors’ fitted to a turbofan engined airliner.

Engine CowlingsFigure 19.4

The method of holding the large cowlings securely closed, usually involve ‘overcentre’ fasteners.(figure 22.5) These fasteners will have a hook-and-clasp system allowing the fastener parts to be hooked together, whilst the cowling is almost closed, and then allowing the two parts to be pulled snugly together by a form of lever. The lever will also be covered by a small access panel.


Depending on the size of the cowling, there can be any number of primary, overcentre latches, backed up by other types of fastener, normally always quick acting.

Cowling doors can be manufactured from either aluminium alloy, Glass Reinforced Plastic, G.R.P., in a honeycomb formation or even Carbon Reinforced Plastic, C.R.P.

Cowling FastenerFigure 19.5


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ACOUSTIC PANELS

Whilst great steps have been taken to silence the exhaust noise by air-mixing and other methods, there are still lower levels of noise to be suppressed. Increasing the size and slowing the fan has lowered the noise from the front of the engine, leaving the highest amount of noise, now being produced by the engine, coming from the internal rotating assemblies especially the turbine. To reduce the noise from these assemblies, optimum use of acoustically absorbent linings is made.

Noise absorbing ‘lining’ material converts acoustic energy into heat. The absorbent linings, shown in the illustration below, normally consist of a porous skin supported by a honeycomb backing, to provide the required separation between the ‘facesheet’ and the solid engine duct.

The acoustic properties of the skin and the liner depth are carefully matched to the characteristics of the noise, for optimum suppression. The disadvantages of this method of suppression are the slight increase in weight and skin friction and, hence, a slight increase in fuel consumption. This method does, however, provide a very powerful suppression technique.

Various materials can be used to produce acoustic linings for jet engines. They fall mainly within two categories, lightweight composite materials are used in the lower temperature regions, and fibrous metallic materials are used in the higher temperature regions. The noise absorbing material consists of a perforate metal or composite facing skin, supported by a honeycomb structure on a solid backing skin which is bonded to the parent metal of the duct or casing.


Acoustic Cowling MaterialsFigure 19.6


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ENGINE MOUNTINGS

Earlier in this chapter, the mounting/suspension of the power units was mentioned. The critical part of the assembly is the mounting itself, which has to transfer all the forces, including both the thrust as well as aerodynamic loads from the engine to the airframe itself. The mount also has to isolate the airframe from any vibrations produced by the engine or its accessories.

Many of these mountings are made of stainless steel and can be located in different patterns around the engine, depending on the designer’s need to manage all of the forces encountered. Some mounts that are not required to carry high loadings can be manufactured from certain hard rubber compounds, which can be formed into circular toroidal form or as flat blocks that work in shear.

The first example, illustrated below, is the mounting of the Rolls Royce Tay engine, which has two metallic vibration isolators and one thrust trunnion to carry all the loads from the engine to the airframe, via the two strong crane beams which are attached to two of the fuselage frames. The vibration isolators are of the stainless steel type, (called Met-L-Flex by the manufacturer), and are shown in the detail drawing.

The trunnion has two sets of Met-L-Flex mounts inside that have to carry, in this example, 14,000 pounds of thrust.

Engine Mountings(1)Figure 19.7


Mountings that are for lighter, less powerful engines, can be manufactured from other materials including the most common, rubber and other elastomers. In the following example, which is the installation of the 1,400 S.H.P. CT-58 engine on the top decking of the Sikorsky S-61N helicopter, the mountings are elastomeric and both support the front of the engine and limit the vibration that travels down from the engine to the cabin beneath. As they are only used to support the front of the engine, they are unlikely to be subjected to much heat generated by the aft end of the unit.



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Finally, the P&W 124 engine mountings are again elastomeric and are, in some cases, in shear and others in compression. The mounting frame, to which these are attached, is shown in the introduction to this topic as the tubular example.



CONNECTION OF SERVICES

The basic engine has many self-contained systems within its configuration. These can include fuel, oil, electrical, fire warning, controls, HP air, etc. However, the actual powerplant installation has to be connected to the airframe interface, (pylon, stub-wing, etc.), via many different forms of connection.

Due to the pressures of time and economics, the airframe and engine manufacturers have to make the removal of the engine, especially the disconnection procedure, as simple as possible. To this end, many powerplants have all of their system and service connections in one place, so that they can all be dis-connected and re-attached at one time. To avoid, as much as possible, the loss of vital fluids from some systems when the engine is being disconnected, (not to mention the ingress of air back into those systems), some form of quick-release and/or self-sealing couplings are required.

QUICK DISCONNECT SHELF

The example below shows the quick disconnect shelf where most of the services of the Lycoming T-55 are joined to the ‘aircraft systems’. Notice that, in this example, there are a number of quick release couplings allowing systems such as the hydraulic fluid lines, to be dis-connected without loss of oil and ingress of air or dirt.

Quick disconnect couplingsFigure 19.10


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Another operation to assist in engine changes is the “handing” of the ancillary equipment and connections. This means that an engine installed on the port stub-wing would have a different layout than one installed on the starboard stub-wing. For example, on a popular turbo-fan engine, it has its hydraulic pumps etc., located on the right hand side of the engine. This means that when installed on the aircraft port side, its connecting hoses are all short, between the fuselage and the engine. When installed on the starboard side, however, there has to be a complete gallery of rigid pipes, fitted to the engine, to connect the pumps to the aircraft systems.

BREAK POINTS

The mechanical controls to the engine, Throttle, Fuel Control (H.P.C.) and, if applicable, Propeller Controls, will all have ‘break points’, where the simple removal of a bolt or pin will separate the engine parts from the aircraft control runs.

The simple removal of a pin or bolt has a further advantage during some disassembling operations. If the relevant system is reconnected again, without other disturbance, it should not require re-rigging.

The illustration below shows how a simple turbopropeller engine has the three connections of its major controls; throttle, H.P. cock and fuel trimmer all grouped together at one point to ease disconnection and/or engine removal.

Engine Control DisconnectsFigure 19.11


ELECTRICAL CONNECTIONS

There is a wide range of electrically powered components on the average gas turbine engine, all of which have to be connected to the aircraft’s own electrical and instrument systems. This can include indications of the R.P.M., fuel flow, T.G.T., vibration, as well as a filter blockage warning, ignition leads and other systems. All of these connections are normally made using ‘multi-pin’ connectors, which not only connect 30, or more, different pairs of wires, but also keep the connections free from oil, moisture and dirt by being sealed inside the plug.

A simple, four-pin connector is illustrated below to show how a plug and receptacle fit together. It should be clear that a keyway and groove always ensure that the plug is connected correctly. This is especially important when angled plugs are used to ensure that cable looms, which approach the connectors at a particular angle, are not put under any strain, which might tend to break the wires in the looms.

Electrical ConnectionsFigure 19.12


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LIFTING POINTS

As mentioned earlier, the engines on commercial aircraft today are designed for the simplest and quickest removal and re-installation. To this end, the engines will often be found with built-in lugs and other attachment points that will allow a crane, of suitable lifting capacity, to lift the engine from the airframe attachments to a suitable support frame or transportation trolley. If some form of lifting sling is used between the crane hook and the engine attachments, it must be checked both for serviceability and that its inspection due date, marked on it, has not been exceeded.

If the engine is to be lifted using a purpose built sling, care must be taken to check the relevant manuals because some lifting slings have more than one attachment point for the crane. This is to allow the engine to remain in balance, (the C of G of the engine under the crane hook, when lifting), in different configurations. For example, if the engine can be lifted with or without the thrust reverser or nose cowling fitted, the centre of gravity will be in a different position for each and will, therefore, require a different lifting point on the sling. The same applies with turbopropeller engines that must be removable both with and without the propeller being fitted.

The engine hoisting sling shown below, from the Rolls Royce Tay engine series, has pin attachments to connect it to the engine and a range of lifting lug positions on the top, which allow a variety of engine configurations to be lifted.

Engine Lifting SlingFigure 19.13


DRAINS

Drains are installed on gas turbine engines of all types, to relieve the collection of different fluids from either within the engine itself or from one or more of its external units. The drains will run into collector tanks, which will be either drained during maintenance or will be self-emptying whilst the engine is operating.

Drains can be found from the fuel and hydraulic systems mainly; although, in most circumstances, the drains are only put to use in situations when something out of the ordinary has occurred.

For instance, if the engine has been turned over to start it, but it has not ‘lit up’ the start must be terminated after a specific time to avoid overheating the starter motor. This means that a quantity of fuel has been spraying into the engine without igniting and, therefore, a second start cannot be initiated until the fuel from the first start has drained down to a collector tank.

With the hydraulic pumps that are fitted to the accessory gearbox on the engine, there is a slight risk that if a seal failed, the hydraulic oil might mix with the gearbox/engine oils. To avoid this, a space is left between the two components and this space is drained to the outside of the engine cowling. A drip of any oil from this drain indicates that one or other of the seals has failed and deeper investigation is required. Drains of this type are often referred to as ’witness drains’ in that they bear witness to a fault.

As a general rule, if drains are directed outside the engine cowlings, they indicate that the leaking fluid is from somewhere it should not be leaking from. If, however, the fluid is directed to a collector, or holding, tank then the fluid is intended, at some time, to be there. (Such as the ‘wet start’ described earlier).

The illustration overleaf, from the Rolls Royce Tay engine again, it can be seen that any leaks from the hydraulic pumps, air starter, internal gearbox, front bearing and I.D.G. will show beneath the engine and alert the engineers of the fault. Leakage from the H.P. shut off valve and various fuel and air control units is directed to the drain tank which returns the fuel back to the L.P. pump.


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Focker 100 drains systemFigure 19.14




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15.20 FIRE PROTECTION SYSTEMS

DETECTION SYSTEMS

Because of the complexity of many modern aircraft, there has to be a reliable fire detection system, installed in the engine nacelles/cowlings. This will indicate to the flight crew both that there is a confirmed overheat/fire in the engine nacelle or bay and that the overheat/fire has been dealt with and the danger is passed.

N.B. The reason for using the terminology overheat/fire is that is the vast majority of ‘fire’ warnings are not actually fires, with combustion taking place, but are what are generally known as ‘hot gas leaks’. These warnings often follow the failure of bleed air ducts, jet pipe connections, etc.

There are three major different types of detection equipment, Thermal switch fire detection, Thermocouple fire detection and Continuous loop fire detection systems.

THERMAL SWITCH

The thermal switch system is sometimes known as the ‘spot detection’ system and consists of a number of separate detectors, which are located in the most likely positions within the nacelles where fire may occur. The detectors are all in parallel, allowing any one detector to give the warning to the flight deck.

The simple detector shown below will expand longitudinally, pulling the contacts together and setting off the alarm. As soon as the heat has been removed, (either by shutting down the engine or discharging the extinguisher bottle), the detector shortens again, allowing the spring wires to separate the contacts, cutting off the fire warning.

Thermal SwitchFigure 20.1


THERMOCOUPLES

The thermocouple system operates on the rate of temperature rise principle, rather than operating when a specific temperature is reached, such as with the spot detectors. The detection is actuated by a number of thermocouple ‘hot junctions’ in likely locations for fires. The thermocouples are in pairs, (a detector and a reference unit), but they will only detect fire, not a general bay overheat. This is because the detector has to have a temperature difference between it and the reference unit, which does not occur with a bay overheat situation but will with a direct fire source. The thermocouples themselves are manufactured exactly the same way as those units that measure the temperature of the jet pipe or turbine gasses. They consist of two metals, usually ‘Alumel’ and ‘Chromel’, which will generate a small current when heated, the current proportional to the temperature.

CONTINUOUS LOOP (Resistance/Capacitance)

The third, and most common, system for the detection of overheat/fire in engine bays and similar locations, is the Continuous Loop fire detection system. This system is often also known as the ‘firewire’ detection system. This system allows a more complete coverage of a fire hazard area than any type of spot detection systems. This system works on the same basic principle as the spot detectors, except that in place of individual detectors, a continuous-loop will detect an overheat or fire anywhere along the length of the tube.

The firewire consists of an Inconel outer tube with a central Nickel wire. Separating the inner from the outer is a eutectic salt, which changes resistance with the application of heat, an increase in temperature causing a decrease in resistance. This allows the measurement of both resistance and capacitance. On the latest installations loops are are normally run throughout the engine bay(s) in pairs. The result of this is that to get a 100% confirmed fire warning, there needs to be a correct change in both the resistance and capacitance in both firewire loops at the same time.

This system has the advantage that it will continue to operate, even if one of the firewires is completely broken. (Although the break will be detected during the firewire test carried out pre-flight).

Overleaf, is illustrated both the construction detail of the inside of a fire wire and a typical firewire installation as fitted to the Pratt & Whitney 124 engine on the ATR-72 aircraft.


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Firewire InstallationFigure 20.2

CONTINUOUS LOOP (Gas )

Another type of continuous loop fire detection system is the pressure-type. This tube contains a sealed gas filled tube containing an element that absorbs gas at a low temperature and releases it as the temperature rises. The tube is connected to a pressure switch that will close when the gas pressure in the tube reaches a pre-determined value. This pressure increase can result from either a localised sharp rise in temperature, (Fire), or a gentler rise in temperature over a longer length of tube, (Overheat).


CONTROLS AND INDICATIONS

The flight deck part of the fire detection and warning system can vary from aircraft to aircraft. As a general rule the display panel shown below could be taken as a typical example of a ‘fire’ panel on a modern aircraft applicable to the main engines. The push buttons labelled ‘LOOP’ are for warning and the isolation of faulty warning loops, whilst the ‘AGENT 1/2’ lights indicates whether the extinguisher bottles have been discharged.

Fire Detection and Extinguishing PanelFigure 20.3

NOTE: Many A.P.U. installations have a fully automatic fire detection, extinguishing and engine shutdown system.

Because of the vital importance of the warning, all cockpit fire alerts will have the highest priority within the cockpit alerting system, being a warning more than a caution. They will also have aural alerts and flashing lights indicating both the relevant fire extinguisher handle and the correct fuel shut-off lever.

As an example, a fire detected within an engine cowling would probably have the following effects:

A bell, ‘chime’ or other audible alert. A flashing ‘attention getting’ light usually coloured red. A ‘FIRE’ caption on the master warning panel. A light in the fire extinguisher operating handle/button. A light in the relevant fuel shut-off or H.P. cock lever.

Once the fire has been detected and the correct engine identified, from maybe two, three or four, the pilot will discharge the first, (of two), pressurised fire bottles containing an extinguishant into the relevant engine bay.

The chemical extinguishing fluid will be directed through pipes to spray nozzles in the bay, which will give the quickest spread of the chemical throughout the bay, smothering the fire. If the fire persists after the first bottle has been discharged, the pilot has the option of firing the second bottle into the same bay, which normally is sufficient to put the worst fire out.


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15.21 ENGINE MONITORING AND GROUND OPERATION

PROCEDURES FOR STARTING AND ENGINE RUN-UP

Firstly, it is essential that before any ground running is commenced, reference is made to the safety precautions and aircraft handling and storage notes that can be found within Module 7. These notes will cover flight line safety with specific mention of fuelling, towing, parking, noise, marshalling, etc. They will also cover such topics as chocking, taxiing, securing, de-icing and work during other inclement weather situations.

Before ground running an aircraft, reference must also be made to the current Ground Handling Procedures Manual. This will list not only all of the rules that must be followed, it will also give the references that are used to ensure the safety of all concerned both directly and indirectly with the ground run.

Other topics may need to be consulted in this manual. They are the aircraft noise limitations, the procedure to be followed in the event of a fire breaking out, where the ground power units are to be placed and the ground support equipment that will be required for the ground run.

The aircraft is parked with brakes applied and chocks in place; all personnel are in their correct places both inside and outside the aircraft and all equipment is in place, such as power unit and warning signs. It is now time to consider what is the object of the ground run, why is it taking place and what it is hoped to achieve.


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INTERPRETATION OF ENGINE POWER OUTPUT

The power output, as was mentioned earlier, is normally measured in E.P.R., (Turbo-jet) or Torque, (Turbo-propeller) aircraft. The ratios or units that these indicating systems use can be used to indicate the power of the aircraft engines ‘on-the-day’ and to indicate any loss of performance since the last engine run, (Provided the weather and all other circumstances are similar). These systems also show when any power boosting system, fitted to the aircraft, is operated and whether it is working to full capacity.

When the engine run is being carried out, it will normally be necessary for other parameters to be noted, at specific times, at certain engine powers and when called for by the ground running schedule itself. These figures could include the following parameters:

T.G.T. Oil Temperature Oil Pressure R.P.M. Fuel Flow E.P.R./Torque Bleed Air Temp. Bleed Air Press Vibration Level

Ground running is normally made much easier, especially when a complex, large and modern jet airliner is concerned, by the use of a ground running booklet. This has both the running procedure itself and also places for marking specific data or ‘tick boxes’ to confirm certain operations were carried out during the run.To give an idea of what might be found in a booklet for ground running, the ATR-72 engine run-up guide contains the following information:

General details of safety areas, control operations, adjustments & emergencies:

Engine control theory. Aircraft preparation. Propeller control theory. Engine starting procedure. Failure codes produced by the engine computers. Engine parameters and limits. Engine stopping procedure.

And, of course,

Operational tests during engine running.

The booklet also includes sets of Diagrams, (safe zones, etc.); Tables, (limits) and Graphs, (performance calculations).


TREND MONITORING/ANALYSIS

There is a much stronger reliance these days on trends, than there used to be. Years ago, if an engine started to use a little more fuel or oil, if it got progressively rougher when running or if it ran a little hotter than its companion engines, then little interest was shown until either the engine became too bad too use or, in the worst case, it failed completely. Today we monitor trends in a number of different ways and they can loosely be divided into three groups:

1. Those which monitor the condition of the oil-washed components such as gearboxes and bearings.

2. Techniques such as vibration analysis and non-destructive testing, (N.D.T.), that can be applied to both the air and oil washed components of the engine. (The N.D.T. inspection procedure is a complex and specialised technique that will be covered later).

3. Those, which enable the air-washed components such as blades, guide vanes and combustion chambers to be inspected.

With oil washed components any mechanical wear from contact surfaces, such as gears and bearings, will produce debris which will be carried within the oil circulating around the engine. Analysis of debris in the oil system can provide a very useful method of assessing any trends in the wear from the internal engine components. This analysis can involve a number of different methods, including:

Magnetic detector plug debris analysis Oil filter debris analysis Spectrometric Oil Analysis Programme (SOAP).

MAGNETIC CHIP DETECTORS

The magnetic chip detectors, (M.C.D.), are small permanent magnets placed in the oil scavenge/return line to collect ferrous debris from the oil. At specific intervals, the plugs are removed and either visually inspected ‘on site’ or returned to a specialist department. The particles are examined, by a skilled and experienced technician using a microscope, and the debris particles identified as being from ball bearings, roller bearings, gearbox teeth, propeller mechanisms, etc.

From these inspections, the decision as to whether to reject the component or just to increase the frequency of sampling depends on the trend of the particle build-up. A slow and steady build-up can best be monitored by an increase in inspection frequency, whilst a sharp increase between one inspection and the next could require the immediate removal of the offending gearbox or engine.


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OIL FILTERS

The analysis of debris collected in the oil system filters, both pressure and scavenge, forms a useful technique to complement the M.C.D. method. The debris are carefully washed off the filter and collected for thorough examination. Particles of some metals will require further careful checking, whilst some metal debris might be considered quite normal.

The particles obtained from the filters will be catalogued for comparison with other, later, debris collections although, the frequency of filter inspection is less than for M.C.D.s. The information obtained provides a back up for establishing and confirming trends within the components.

S.O.A.P

The spectrometric oil analysis programme removes a sample of oil from the engine oil system. This will contain microscopic traces of all the metals that are washed by the oil. Chemical analysis of the sample particles enables the wear rate to be evaluated and quantified. In a typical oil system, traces of the following metals might be found:

Aluminium Iron Chromium Silver Copper Tin Magnesium Lead

VIBRATION

Vibration analysis is part of the on-condition maintenance policy that checks on the condition of rotating assemblies. The requirement is that defects can be detected sufficiently early to permit rectification before secondary and more serious damage occurs. Analysis of engine vibration signatures is an important tool for the detection of early failure in mechanical components.

A vibration monitoring system consists of a sensor, which converts the mechanical vibration of the machine to which it is attached, into electrical signals that are in proportion to the magnitude of the vibrations. In addition, there are amplifiers with their associated wiring, to produce some form of cockpit display. This display can be analogue, digital or even simply in the form of warning lights.

There will be a vibration limit set by the engine manufacturer, which will be the level to which the flight deck displays have their warning/caution limits set.


On some engines there is more than one limit, which might be both a caution and a warning. The caution would allow the pilot to close the engine down before little or no damage has occurred, whereas the warning would possibly indicate a serious failure within the engine itself.

Analogue and digital displays allow the crew to record readings for each engine, in flight, which can be later compared with previous flights to identify any increase in the levels of vibration.

Other forms of sensor can be found on some aircraft. One is the dual sensor that consists of two separate but integral units that permit the pilot to switch between them if one fails for any reason. Another useful variation is the ‘wide’ and ‘narrow’ band measurement, which means the reading can either be taken from over the whole range of vibrations emanating from the engine, or from one or two major rotating assemblies such as the N1 and N2 spools.

The illustrations below show two sensors that are mounted vertically and horizontally. Also shown is a typical linear (tape) display from a twinjet airliner which is showing both N1 and N2 vibration for both engines. The port engine has its vibration ‘in limits’ while the starboard engine has exceeded the caution limit on both spools.

Vibration sensors and displayFigure 21.1


`VISUAL/OPTICALINSPECTIONS

The examination of air washed components can be divided into two different inspections, unaided visual inspection and optical inspection. These will allow the engineer to inspect the internal parts of the engine, at various times, enabling any deterioration to be detected early, allowing the rectification to start before any damage has gone too far.

VISUAL

Inspection with the naked eye can detect obvious damage in the front stages of the compressor and the rear stages of the turbine. With twin and triple spool engines, the amount of the rotating parts that can be inspected visually is proportionally less.

Such inspections will detect cracks or other damage only when the operator is extremely vigilant. Jet pipe and turbine inspections cannot be carried out immediately after engine shutdown because of the heat. Visual inspection however is still extremely important for monitoring the condition of engines.

OPTICAL

Optical inspection of the inside of a gas turbine engine, including the hot end, is best carried out using a remotely illuminated optical probe, sometimes known as a borescope. These probes can be obtained in a variety of lengths, diameters and, for difficult places, in a flexible form.

On earlier engines, the inspection probes had to be inserted into already existing holes in the engine casing. These might be igniter or thermocouple installation holes, which would have the item removed to allow use of the holes by the borescope. The limitation of this system was that the parts of the engine that were visible to the inserted probes were very limited.

Most modern engines however, have purpose-built inspection holes fitted in the casing at the correct places, thus allowing complete inspection of all of the air washed parts of the engine from the first stage of the compressor to the final stage of the turbine.

However, with rigid borescopes it can still sometimes be difficult to orientate the probe to give the desired view. With modern probe sets, there will be a fibre-optic probe, which is flexible. This can be either fed directly in to the engine or via a purpose built guide tube, which will turn the flexible probe to the exact position for viewing.


Several borescope sets have, in addition to the range of probes mentioned earlier, a low light CCTV camera mounted on to the eyepiece of the probe. This means that more than one person can watch the inspection at the same time. Also, the viewing can be recorded on video both for a permanent record and to allow trend monitoring by comparison of wear, damage and discoloration with earlier recordings.

The illustration below of the Rolls Royce Tay 650 engine shows the borescope and visual inspection access to the inner parts of the engine. It will be seen that almost all of the internal parts of the engine can be inspected by the mixed use of both rigid probes, (solid lines), and flexible probes, (dotted lines).

Engine Inspection LocationsFigure 21.2


`The illustration below shows a flexible borescope being used to inspect the first stage of a turbine assembly via ports on the fan duct and in the combustion chamber.

Boroscope InspectionFigure 21.3


INSPECTION CRITERIA, TOLERANCES AND DATA

Modern engines have exceptionally high serviceability ratings and it is not unusual to find some engines with 20,000 or 30,000 running hours in one position, on the same aircraft. Because of this, there is little to do with respect to ‘inspecting’ the engine on a routine, day-to-day basis.

As a general rule, aircraft in service have a daily inspection, often overnight, between cease flying on one day and the early departure on the next. During the day, if the aircraft is not an inter-continental flyer, there will be pre-departure inspections, prior to the aircraft departing on its next flight leg. Finally, at specified intervals, such as every so many days, an ‘A’ check would be carried out.

All of these checks require an inspection of the engine. This is usually little more than a look into the intake for Foreign Object Damage(F.O.D.); a check of the jet pipe for signs of turbine damage; a check of the engine, gearbox and I.D.G. oil levels and, finally, a look around the engine cowlings, for any signs of damage and oil or fuel leaks.

It is more likely that any ‘inspection’ of the engine will be involved after the aircraft captain has reported a fault/ defect. Trouble shooting an engine will often involve checking some part of the engine’s operation against figures produced by the manufacturer and published in the maintenance manuals.

As an example, the following checks, etc. are typical of those which might be encountered whilst carrying out trouble-shooting on a reportedly defective engine:

POWER – The power of a gas turbine engine in service is never measured in pounds thrust, as it is on the test bed at the manufacturers. As mentioned earlier, the measurement of the power of a gas turbine engine in service is usually as engine pressure ratio, (E.P.R.).

This E.P.R. might be checked against reference figures to prove whether the engine is producing full power whist at maximum revolutions. (If the engine is being run in any conditions other than at sea level and at a temperature of 150C. then adjustments to the figures will have to be made).

TORQUE – Torque is the measure of power being produced by a turbopropeller or turboshaft engine. The torque pressure produced by the engine at maximum power will be noted on a certificate from the manufacturers.

Ground runs will confirm, again with adjustments for airfield height and outside air temperature, O.A.T., whether the engine is producing the torque, (shaft horsepower), that it should be, at full throttle.

T.G.T. – The reading of the T.G.T. on the cockpit gauge can give an indication of the health of that engine. If the engine has some problem internally, it is possible that more fuel may have to be burnt in the combustors to make-up the shortfall in power. This increased fuel flow will show as a higher than normal T.G.T. reading.


`The T.G.T. figures during start-up, maximum power and other engine situations, are published in the aircraft’s flight manual. The pilot (and any engineer ground running the engines) must strictly follow these limitations, as well as reporting when the figures are NOT in accordance with the limits.

REVERSER OPERATION – It is possible that the crew may report that one reverser is slower in operation than its companion(s). Testing the operation of these units will involve timing their operation, both deploying and stowing, either during an engine run or using some external power source to drive the reversers. If not within the published time limits, rectification will be necessary.

R.P.M./% – A number of different values of R.P.M. are published in the engine manuals, including the idle and maximum values. On many engines however, this will be represented not as R.P.M. but as a percentage of maximum engine speed, (100%).

It may be necessary, at times, to convert from percentage to R.P.M., so the maximum 100% value will be published in the manuals to facilitate this. Some other settings that may require checking are flight idle, reverse thrust (less than 100%), and cruise power.

ACCELERATION TIMES – It is vital that the engines accelerate both as quickly as possible, in case the pilot needs to ‘go around’ after being unable to land, and equally, so that there is no asymmetric thrust, which might cause an unwanted turning moment.

To achieve these requirements, it is normal for the engines to be subjected to what are known as ‘slam checks’ at intervals, or whenever the crew reports any problems. The slam is the rapid movement of the throttle from idle to maximum, whilst timing how long the engine takes to ‘catch-up’ to maximum power. Adjustments will be made if the engine is too slow, too fast or if it does not match the timing of the other one(s).

ACCESSORIES – Many of the components fitted on to the engine, such as pumps, generators, measuring devices, etc., have to meet ‘installed performance figures’, meaning for instance, that a fuel pump must have a specific output at a certain engine speed. These figures are again published in the engine maintenance manuals.


COMPRESSOR WASHING/CLEANING

The atmosphere near the ground is filled with small particles of dirt, oil, soot, salt (near the sea) and other foreign matter. A very large volume of air is drawn into the compressor, with centrifugal force throwing the particles outwards so that they build up a coating on the engine casing, the guide vanes and the compressor blades.

An accumulation of dirt on the compressor blades reduces the aerodynamic efficiency of the compressor itself, which results in a loss of engine performance. This deterioration in performance due to the build-up of dirt on the blades is the same as an aircraft wing under icing conditions. Unsatisfactory acceleration and high exhaust gas temperatures can be the result of this build-up.

Gas path erosion occurs from the ingestion of sand, dirt, dust and other fine airborne particles. This erosion can occur in both the compressor and the turbine sections of the engine. The accumulative effect of continuous ingestion of these abrasive particles can result in the erosion of the surface coatings, and even further into the base metal of the fan, the compressor blades and vanes.

Two common methods for removing dirt, salt and corrosive deposits are fluid (liquid) wash and an abrasive grit blast.

FLUID CLEANING

The fluid cleaning procedure is easily accomplished by first spraying an emulsion surface cleaner and then applying a rinse solution to the compressor. This is carried out whilst the engine is either being motored over by the starter or during low speed operation.

Note: It cannot be overstressed that the wash procedure must be performed in strict accordance with the instructions laid down in the manufacturer’s manual.

Usually, when the water wash is performed solely to remove salt deposits, the compressor wash is known as ‘desalination’. If the solution wash is performed solely to remove baked on deposits to improve engine performance, the wash may be known as a ‘performance recovery wash’.

Motoring washing is carried out at whatever speed the starter will rotate the engine up to, typically between 15% to 25%, with the cleaning mixture injected at high pressure.Running washing is carried out around ground idle, again typically 60% with a lower water pressure.

The use of washing, on a regular basis, can vastly extend the life of an engine, especially when the engine is given regular performance ground runs, measured against standard figures.


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ABRASIVE GRIT PROCESS

A second, more vigorous method of compressor cleaning is to inject abrasive grit into the engine at selected power settings. The grit used may be ground walnut shells or apricot pits. The type and amount of material and the procedure to be used is prescribed by the engine manufacturer.

Whilst the intervals between operations is much longer, the abrasive effect on some parts of the engine and the fact the grit is burned up in the turbine, (giving no cleaning), means that this process is not carried out as often as washing.

Sometimes there is a ‘rule-of-thumb’, which says that if the deposit on the compressors is still wet, then a liquid wash is all that is required. However, if the deposit has hardened then abrasive grit is required.

Whether the material used for cleaning the compressor is liquid or solid, there will be a correct procedure and most likely apparatus for delivering the cleaning medium to the front of the compressor safely.


FOREIGN OBJECT DAMAGE (F.O.D.)

F.O.D. is the term used whenever such items as stones, grit, nuts, bolts; in fact anything which can be found around the aircraft movements area, are sucked into the intakes of a running jet engine. The damage that occurs is always expensive; it just depends on whether it is a case of minor damage to a single compressor blade or the total destruction of the engine. Any signs of damage in the intake area must always be assumed to indicate further, worst, damage to the engine.

On most aircraft the front stage(s) of the engine can be seen with little effort, however, certain aircraft, such as the DC-10/MD-11series and the Boeing 727 both have ‘buried’ centre engines. These will require great care when inspecting the engines for signs of damage, due to the fact that an engineer would be invisible to others who might start the engine. There must be others involved as safety-men and the flight deck must be ‘placarded’ with DO NOT START signs.

The damage that the compressor blades have received must be classified in accordance with the manual, which will dictate whether it is a case of a single blade replacement, a fan stage replacement or a total engine change.

If the damage received by the blade is slight, (as defined by the repair manual), then the damage can be ‘dressed out’ using a selection of fine abrasive tools. These change any sharp edged damage into a smooth curved hollow, which will not cause later failures due to cracking.

The examples of blade damage, illustrated below, show some extreme cases of how a blade can be damaged and, below that, how the damage can be blended or scalloped, if the damage is classified as slight.


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Blade Damage Examples

Figure 30.1

Foreign object damage can also be caused by poor workmanship and husbandry by maintenance engineers. When working either in the engine intakes or in the vicinity of them, great care must be taken not to drop any ‘hardware’ such as nuts and bolts, or even larger items such as tools.

Whilst it is the responsibility of “someone else” to sweep the ramp, taxiways and other aircraft movement areas, everyone who works around aircraft should acquire the habit of:

IF YOU SEE SOMETHING ON THE GROUND THAT SHOULD NOT BE THERE…

…PICK IT UP!




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15.22 ENGINE STORAGE AND PRESERVATION (Level B1 only) (ref; Civil Aircraft Airworthiness Information and procedures. CAAIP Leaflet 7-4)

INTRODUCTION

Under normal operating conditions the interior parts of an engine are protected against corrosion by the continuous application of lubricating oil and oil mist, the operating temperatures of which, are sufficient to dispel any moisture which may tend to form, particularly during the wide temperature change which take place during engine shut down and start up. After shut down the residual oil film gives protection for a short period of time which varies, depending on the local environmental conditions. When the engine is not in regular service, parts which have been exposed to the bi-products of combustion and internal parts in contact with acidic oil, are prone to corrosion. If engines are expected to be out of use for an extended period of time, they should be ground run periodically or, some form of anti-corrosive protection or treatment must be applied internally and externally to prevent deterioration.

The type of engine protection applied depends on how long it is expected to be out of service, if it is installed in the aircraft and if it can be turned.

The procedures adopted and the level of protection applied will very from one manufacturer to another. In all cases, the approved Maintenance Manual procedure should be complied with.

INSTALLED TURBINE ENGINES:- which are not going to be run for periods up to seven days normally do not require protection apart from the installation of external blanking covers to air intakes, exhausts and any other openings. This provides protection against the ingress of dust, rain, sand, snow etc.

STORAGE PROCEDURES

Leaflet 7-4 of CAAIPPREPARING FOR STORAGE AND DISPATCH

The preparation of the engine/module for storage and/or despatch is of major importance, since storage and transportation calls for special treatment to preserve and protect the engine from deterioration and damage. To resist corrosion during storage, the fuel system is inhibited by special oil and all apertures are sealed off. The internal of the engine and engine components such as fuel pumps, hydraulic pumps, air starter motors, etc. are also protected by inhibiting oils or powders. External surfaces are usually protected by paper impregnated with inhibiting powder or oil and the engine is enclosed in a re-usable bag or plastic sheeting into which a specific amount of desiccant silica-gel is inserted. If transportation by tail or sea is involved, the inhibited or bagged engine is usually packed in a wooden crate or metal case.


PRESERVATION AND DEPRESERVATION

Preservation: is the procedure recommended as the minimum necessary to protect a power plant against: - liquid or debris (FOD) entering the power plant Corrosion Atmospheric conditions during periods of storage and inactivity.

Depreservation: Is the procedure recommended to restore a preserved engine back to an operational state.

NOTE: For the purpose of this procedure, “operational state” is defined as: “A power plant that can be started”. (On or Off the aircraft.)

PERIODS OF STORAGE (Installed engines) can be divided into two catagories:

1. Short Term Storage – from 7 days up to 1 month.2. Long Term Storage – from 1 month up to 6 months.



Documents

Mod 15 (b1) Gas Turbines