m14 propulsion

Module 14 Propulsion

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EASA Part-66 Training Handbook Module 14 LINK & LEARN

Gas Turbine Engine

15.1 Fundamentals

Principles and Working Cycles of Gas Turbine Engines

15.1.2.1 Introduction

During the last 40 years, the development of gas turbine engines as propulsion systems for aircraft has been very fast. It is difficult to appreciate that before the 1950s very few people knew about this method of aircraft propulsion. Aircraft designers had been interested in the possibility of using a reaction turbine for a long time. But initially, the low speeds of early aircraft and the unsuitability of a piston engine for producing the large high-velocity airflow necessary for the 'jet' caused many problems.

Refer to Figure 1.

A French engineer, Rene Lorin, patented a jet propulsion engine in 1913. But this was an athodyd and, at that time, it could not be manufactured or used since suitable heat resisting materials had not been developed.

Note: An athodyd (or: pulse jet engine) is an open tube which is shaped to produce thrust when fuel is ignited inside it. Fuel is added to the incoming air as the athodyd moves through the air at high speed. This burning causes air expansion that speeds up the air and produces thrust (Figure 1, detail a)).

Secondly, jet propulsion would have been extremely inefficient at the low speeds of early aircraft. However, today's modern ram jet is very similar to Lorin's conception.

In 1930, Frank Whittle was granted his first patent for using a gas turbine to produce a propulsive jet (Figure 1, detail b)). But it took 11 years before his engine completed its first flight. The Whittle engine formed the basis of the modern gas turbine engine. The ROLLS-ROYCE, DERWENT, NENE or DART engines were derived directly from the Whittle engine.

The DERWENT and the NENE jet engines were mainly installed in military aircraft. The DART turboprop engine became well known as the power plant for the VICKERS Viscount aircraft. Although other aircraft may be fitted with later engines termed 'twin-spool', 'triple-spool', 'by-pass', 'ducted fan', 'unducted fan' or 'propfan', they are developments of Whittle's early engine.

Refer to Figure 2.

Although the jet engine appears to be very different from a piston engine with a propeller, it applies the same basic principle to produce propulsion. Both propel the aircraft solely by moving a large volume of air rearwards.

Although today's jet propulsion is popularly linked with the gas turbine engine, there are other types of jet-propelled engines, such as the ram jet, the pulse jet, the rocket, the turbo-ram jet and the turbo-rocket.

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15.1.2.2 Principles of Jet Propulsion

Jet propulsion is a practical application of Sir Isaac Newton's third law of motion which states that

'for every force acting on 2 body there is an opposite and equal reaction'.

For aircraft propulsion, the 'body' is atmospheric air that is caused to accelerate as it passes through the engine. The force required to cause this acceleration has an equal effect in the opposite direction, i.e. it acts on the components producing the acceleration.

A jet engine produces thrust in a way similar to the piston enginelpropeller combination. Both propel the aircraft by moving a large volume of air backwards: one in the form of a large air slipstream at comparatively low speed and the other in the form of a jet of gas at very high speed.

Refer to Figure 3

This same principle of reaction occurs in all forms of movement and has been usefully applied in many ways. The earliest known example of jet reaction is that of Hero's engine (Figure 3, detail a)) produced as a toy more than 2,000 years ago. This toy showed how the momentum of steam exiting a number of jets could impart an equal and opposite reaction to the jets themselves, causing the engine to rotate.

The whirling garden sprinkler (Figure 3, detail b)) is a more practical example of this principle: its mechanism rotates due to the reaction to the water jets. The high-pressure jets of modern fire-fighting equipment are another example of 'jet reaction': due to the reaction of the water jet, the hose cannot be held or controlled by one single fireman. Perhaps the simplest illustration of this principle is a toy balloon which, when the air or gas is released, rushes rapidly away in the direction opposite to the jet.

Jet reaction is definitely an internal phenomenon and does not result from the pressure of the jet on the atmosphere. In fact, the jet propulsion engine, whether rocket, athodyd or turbojet, is a piece of equipment designed to accelerate a stream of air or gas and to expel it at high velocity.

There are, of course, a number of ways of doing this. But in all instances, the resultant reaction (or: thrust) exerted on the engine is proportional to the mass or weight of air expelled by the engine and to the velocity change imparted to it.

Refer to Figure 2 again.

In other words, the same thrust can be provided

either by giving a large mass of air a little velocity increase

or by giving a small mass of air a large velocity increase.

In practice, the former is preferred, since by lowering the jet velocity relative to the atmosphere a higher propulsive efficiency is obtained.

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15.1.2.3 Types of Jet Engine according to Jet Propulsion Methods

The types of jet engine, whether ram jet, pulse jet, rocket, gas turbine, turbo-ram jet or turbo-rocket, differ only in the way in which the 'thrust provider' (or: engine) supplies and converts the energy into power for flight.

Ram Jet Engine

Refer to Figure 4.

The ram jet engine (Figure 4, detail a)) is an athodyd (or: aero-thermodynamic duct). It has no major rotating parts and consists of a duct with a divergent entry and a convergent or convergentidivergent exit.

When forward motion is imparted to it by an external source, air is forced into the air intake. Here, it loses velocity (or: kinet~c energy) and increases its pressure (or: potential energy) as it passes through the diverging duct. Then, the total energy is increased by the combustion of fuel. F~nally, the expanding gases are expelled to the atmosphere through the outlet duct.

A ram jet is often used as a power plant for missiles and target vehicles. But it is unsuitable as an aircraft power plant because it requires forward motion before any thrust can be produced.

Pulse Jet Engine

The pulse jet engine (Figure 4, detail b)) uses the principle of intermittent combustion. Unlike the ram jet, it can be run at a static condition. The engine is formed by an aerodynamic duct similar to the ram jet. But, due to the higher pressures involved, it is of more robust construction. The duct inlet has a series of inlet 'valves' that are spring-loaded in the 'open' position.

Air drawn in through the open valves passes into the combustion chamber and is heated by the burning of fuel injected into the chamber.

The resulting expansion causes a rise in pressure, forcing the valves to close and the expanding gases are then ejected rearwards. A depression created by the exhaust- ing gases allows the valves to open again and the cycle is repeated.

The pulse jet is unsuitable as an aircraft power plant because it has a high fuel consumption and is unable to reach the performance level of the modern gas turbine engine.

Rocket Engine

Although a rocket engine (Figure 4, detail c)) is a jet engine, there is one major difference: it does not use atmospheric air as the propulsive stream. Instead, it produces its own propelling medium by the combustion of liquid or chemically decomposed fuel with oxygen. It is able to operate outside the earth's atmosphere.

Gas Turbine Engine

The application of the gas turbine to jet propulsion has overcome the inherent weakness of the rocket and the athodyd: a means of producing thrust at low speeds was provided by the introduction of a turbine-driven compressor.


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The turbojet engine draws air from the atmosphere. After compressing and heating it (a process that occurs in all heat engines) the energy and momentum given to the air forces it out of the propelling nozzle at a velocity of up to 2,000 feet per second (approx. 610 m/s or 2,200 kmlh). On its way through the engine, the air gives up some of its energy and momentum to drive the turbine that powers the compressor.

The mechanical arrangement of the gas turbine engine is simple. It consists of only 2 main rotating parts (a compressor and a turbine) and one or a number of combustion chambers.

Note: The mechanical arrangements of various types of gas turbine engine are shown in Figures 5 to 7.

This simplicity, however, does not apply to all aspects of the engine: the thermodynamic and aerodynamic aspects are quite complex. They result from:

the high operating temperatures of the combustion chamber and the turbine

the effects of varying flows across the compressor and the turbine blades

the design of the exhaust system through which the gases are ejected to form the propulsive jet.

Refer to Figure 8.

At aircraft speeds below approx. 450 knots (knots = nautical miles (nm) per hour), the pure jet engine is less efficient than a propeller-type engine, since its propulsive efficiency largely depends on its forward speed. The pure turbojet engine is most suitable for high forward speeds. The propeller efficiency does, however, decrease rapidly above 350 knots due to the disturbance of the airflow caused by the high blade-tip speeds of the propeller.

The advantages of the turbinelpropeller combination have to some extent been offset by the introduction of the by-pass, ducted fan and propfan engines.

These engines provide larger airflows and lower jet velocities than the pure jet engine. They give a propulsive efficiency which is comparable to that of the turboprop engine and exceeds that of the pure jet engine.

Turboshaft Engine

A gas turbine engine 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 electric power generating plants and surface transportation systems, while in aviation, turboshaft engines are used to power many modern helicopters.

Refer to Figure 9.

The turboshaft power take-off may be coupled to and driven directly by the turbine that drives the compressor, but it is more likely to be driven by a turbine of its own. Engines using a separate turbine for the power take-off are called 'free turbine engines' or 'free-power turbine-type turboshaft engines'.

A free turbine turboshaft engine has two major sections, the gas generator and the free turbine section. The function of the gas generator is to produce the required energy to drive the free turbine system and it extracts about two third of the energy available from the combustion process leaving the other one third to drive the free-power turbine.



Turbo -ram Engine

Refer to Figure 10

The turbo-ram jet engine (Figure 10, detail a)) combines the turbojet engine (which can be used for speeds up to Mach 3) with the ram jet engine, which shows good performance at high Mach numbers.

The engine is surrounded by a duct that has a variable intake at the front and an afterburner jet pipe with a variable nozzle at the rear.

During take-off and acceleration, the engine works like a conventional turbojet with afterburner. At other flight conditions up to Mach 3, the afterburner is inoperative. As the aircraft accelerates beyond Mach 3, the turbojet is shut down and the intake air is diverted by guide vanes from the compressor. It is ducted straight into the afterburning jet pipe, which now works as a ram-jet combustion chamber.

This engine is suitable for an aircraft which requires high-speed and sustained- high-Mach-number cruise conditions.

Turbo-rocket Engine

The turbo-rocket engine (Figure 10, detail b)) is an alternative to the turbo-ram jet. However, there is one major difference: it carries its own oxygen to provide combustion.

The engine has a low-pressure compressor driven by a multi-stage turbine. The power required to drive the turbine is derived from combustion of kerosine and liquid oxygen in a rocket-type combustion chamber. Since the gas temperature is approx. 3,500 "C, additional fuel is sprayed into the combustion chamber for cooling purposes before the gas enters the turbine. This fuel-rich mixture (gas) is then diluted with air from the compressor. The surplus fuel is burnt in a conventional afterburning system.

Although the engine is smaller and lighter than the turbo-ram jet, it has a higher fuel consumption. This makes it more suitable for being used in an interceptor or space-launcher type of aircraft that requires high speed, high altitude performance and (normally) has a flight profile which is entirely accelerative and of short duration.

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EASA Part-66 Training Handbook Module 14

a) Lorin's jet engine

Combustion chamber

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Air intake Propelling nozzle

Fuel supply

b) Whittle-type turbojet engine

Combustion chamber Compressor / Turbine

u \ Fuel burner ~ e t pipe and

Air intake propelling nozzle

Figure 1 Principle of Jet Engines

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EASA Part-66 Training Handbook

Turbojet engine

mass

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Acceleration a 0

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Turboprop engine

ff = Acceleration mass

Figure 2 Comparison of Propulsion Systems



a) Hero's engine (probably the earliest form of jet reaction)

b) Rotation effect by the reaction of water jets

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Figure 3 Forms of Jet Reaction

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a) Ram jet engine

Fuel burners Combustion chamber

Air intake Propelling nozzle

b) Pulse jet engine Charging

shutter valve- Firina

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. .

Fuel \ Air intake supply Combustion chamber

c) Rocket engine Combustion Propelling Liauid fuel chamber nozzle

Oxygen Fuel injectors

Figure 4 Basic Methods of Jet Propulsion

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a) Double-entry single-stage centrifugal turbojet engine

b) Single-entry 2-stage centrifugal turboprop engine

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c) Twin-spool axial flow turboprop engine

Figure 5 Arrangement of Gas Turbine Engines (I)



a) Single-spool axial flow turbojet engine

b) Twin-spool turboshaft engine (with free-power turbine)

Figure 6 Arrangement of Gas Turbine Engines (11)


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a) Twin-spool by-pass turbojet engine (low by-pass ratio)

b) Triple-spool front fan turbojet engine (high by-pass ratio)

c) Propfan concept

d) Contra-rotating fan concept (high by-pass ratio)

Figure 7 Arrangement of Gas Turbine Engines (Ill)



Low by-pass ratio

High by-pass ratio, \

0 200 400 600 800 1,000 Airspeed (knots)

0 200 400 600 800 1,000 Airspeed (knots)

Figure 8 Comparative Efficiencies



a) Power conversion free turbine Fuel

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v Gas generator

b) An example of a free turbine engine that has been adapted for both turboprop and turboshaft applications

Exhaust outlet 7 Air 7 r- Compressor

l ~ r e e - ~ o w e r turbine ~ o m ~ r e s s o r turbine

Figure 9 Free Turbine Engine



a) Turbo-ram jet engine

Variable intake Intake guide vanes Variable nozzle (large area) (open) (large area)

Low Mach number

Variable intake (small area)

Intake guide vanes Variable nozzle (shut) (small area)

High Mach number

b) Turbo-rocket engine

Variable intake Afterburning Combustion chamber fuel burns

Oxygen and Variable nozzle fuel supply

Figure 10 Schematic Cross-section of a Turbo-ram and a Rocket Engine



15.2 Engine Performance

15.2.1 Working Cycle and Airflow of a Gas Turbine

General

The gas turbine engine is essentially a heat engine using air as a working fluid to provide thrust. To achieve this, the air passing through the engine has to be accelerated. This means, that the velocity (or: kinetic energy) of the air is increased. To obtain this increase, first of all the air pressure (potential energy) is increased, followed by the addition of heat energy. Finally, the potential energy is reconverted into kinetic energy in the form of a high-velocity jet efflux.

15.2.1.1 Working Cycle

The working cycle of the gas turbine engine is similar to that of the 4-stroke piston engine. However, in the gas turbine engine, combustion occurs at a constant pressure, whereas in the piston engine it occurs at a constant volume. Both engine cycles consist of induction, compression, combustion and exhaust.

These processes are intermittent in the case of the piston engine whilst they occur continuously in the gas turbine engine. In the piston engine, only one stroke is used in the production of power, the others being involved in the charging, compression and exhaust of the working fluid. In contrast, the gas turbine engine eliminates the 3 'idle' strokes, enabling more fuel to be burnt in a shorter time. Thus it produces a greater power output for a given size of engine.

Due to the continuous action of the gas turbine engine, and due to the fact that the combustion chamber is not an enclosed space, the pressure of the air does not rise during combustion (like that of the piston engine). But its volume increases. This process is known as 'heating at constant pressure'. Under these conditions there are no peak or fluctuating pressures to be withstood (as is the case with the piston engine with its peak pressures in excess of 1,000 psi).

These peak pressures require the use of cylinders of heavy construction in the piston engine and the use of high-octane fuels, in contrast to the low-octane fuels and the light fabrication of combustion chambers used in gas turbine engines.

The working cycle of the gas turbine engine can be, in its simplest form, represented by the cycle shown on the pressurelvolume diagram in Figure 1:

point A represents air at atmospheric pressure that is compressed until reaching point 6 from B to C, heat is added to the air by injecting and burning fuel at a constant pressure; the volume of the air pressure is increased considerably

pressure losses in the combustion chambers are indicated by the drop between B and C from C to D, the gases resulting from combustion expand through the turbine and jet pipe back to the atmosphere; during this part of the cycle, some of the energy of the expanding gases is transferred into mechanical power by the turbine; the remainder provides the propulsive jet on its discharge to the atmosphere.


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The higher the temperature of combustion, the greater the expansion of the gases. The combustion temperature, however, must not exceed a certain value to provide a turbine gas entry temperature which is suitable for the design and materials of the turbine assembly.

The use of air-cooled blades in the turbine assembly permits a higher gas temperature and, consequently, a higher thermal efficiency.

Relation between Pressure, Volume and Temperature

During the working cycle of the gas turbine engine, the airflow receives and gives up heat. This results in changes of its pressure, volume and temperature. These changes are closely related to each other, because they follow a common principle that is a combination of the laws of Boyle and Charles.

This principle states, that the product of the pressure and the volume of the air at the various stages in the working cycle is proportional to the absolute temperature of the air at those stages. This relationship applies to any means used to change the state of the air. For example, whether energy is added by combustion or by compression (or is extracted by the turbine), the heat change is directly proportional to the work added to (or taken from) the gas.

There are 3 main conditions in the engine working cycle during which these changes occur:

R during compression, when work is done to increase the pressure and to decrease the volume of the air, there is a rise in the temperature

during combustion, when fuel is added to the air and burnt to increase the temperature, there is an increase in volume whilst the pressure remains al- most constant

during expansion, when energy is taken from the airflow by the turbine assembly, there is a decrease in temperature and in pressure with an increase in volume.

Changes in the temperature and the pressure of the air can be traced through an engine by using the airflow diagram in Figure 2. Because the airflow is continuous, volume changes are shown as changes in velocity.

The efficiency, with which these changes are made, will determine how far the desired relations between pressure, volume and temperature can be obtained. The more efficient the compressor is, the higher will be the pressure generated for a given work input, i.e. for a given temperature rise of the air. Conversely, the more efficiently the turbine uses the expanding gas, the greater will be the output of work for a given pressure drop of the gas.

The efficiency of the process of converting energy during expansion or compression is called 'adiabatic'. An 'adiabatic efficiency of 100 %' means, that no energy is lost during the process, neither by friction, conduction, nor by turbulence. Such a perfect process cannot be achieved in practice. However, an adiabatic efficiency of 90 % is still a good value for the compressor and the turbine.



Changes in Velocity and Pressure

During the passage of the air through the engine, aerodynamic and energy require- ments demand changes in its velocity and pressure.

For example, during compression, a rise in the pressure of the air is required and not an increase in its velocity.

After the air has been heated and its internal energy increased by combustion, an increase in the velocity of the gases is necessary to force the turbine to rotate.

At the propelling nozzle, a high exit velocity is required to provide the thrust on the aircraft. Local decelerations of the airflow are also required, e.g. in the combustion chambers to provide a low-velocity zone for the flame to burn.


These various changes are effected by means of the size and shape of the ducts through which the air passes on its way through the engine.

Where a conversion from velocity (kinetic energy) into pressure (potential energy) is required, the passages are divergent in shape. Conversely, where it is required to convert the energy stored in the combustion gases into velocity energy, a convergent passage (or: nozzle) is used. These shapes apply to gas turbine engines where the air velocity is subsonic or sonic, i.e. below or at the local speed of sound.

Refer to Figure 3.

Where supersonic speeds are encountered, such as in the propelling nozzle of a rocket, athodyd and some jet engines, a convergentldivergent nozzle (or: Venturi) is used to obtain the maximum conversion of the energy of the combustion gases into kinetic energy.

The efficiency, with which the energy changes are effected, depends on the proper design of the passages and nozzles. Any interference with the smooth airflow creates a loss in efficiency and may result in component failure due to vibration caused by eddies or turbulences of the airflow.

Airflow

Refer to Figures 4 and 5.

The path of the air through a gas turbine engine varies according to the design of the engine. A 'straight-through flow' system is the basic design, as it represents an engine with a relatively small frontal area. It is also suitable for application of the by-pass principle.

In contrast, the 'reverse flow' system represents an engine with greater frontal area, but with a reduced overall length.

The operation, however, of all engines is similar. The differences in the different designs are described in the following paragraphs.

Refer to Figure 6.

The by-pass principle effects a division of the airflow. All the air, which is taken in, is given an initial low compression. A percentage is then ducted to a by-pass, the remainder being delivered to the combustion system in the usual manner. This principle provides improved propulsive efficiency and lower fuel consumption.

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An important design feature of the by-pass engine is the by-pass ratio. This is the ratio of the amount of cool air, which is by-passed through the duct, to that of the air passed through the high-pressure system. With low by-pass ratios, i.e. in the order of 1 : I , the 2 streams are usually mixed before being exhausted from the engine.

The requirement for high by-pass ratios of up to 5 : 1 is largely met by using the front fan in a twin- or triple-spool configuration (on which the fan is, in fact, the low-pressure compressor). The fan engine may be regarded as an extension of the by-pass principle.

Very high by-pass ratios, in the order of 15 : 1 , are achieved by using propfans. These are a variation of the turboprop engine but with advanced-technology propellers capable of operating with high efficiency at high aircraft speeds.

On some front-fan engines, the by-pass airstream is ducted overboard either directly behind the fan through short ducts or at the rear of the engine through longer ducts (hence the term 'ducted fan'). Another variation, though seldom used, is that of the aft (rear) fan.



a) Working cycle

Compression Combustion Air intake (continuous) (continuous) Exhaust

b) Pressure/volume diagram

Volume

Legend: - Ambient air - Expansion (through turbine and nozzle) - Compression (pressure energy added) - Combustion (heat energy added)

Figure 1 Working Cycle of a Gas Turbine Engine



a) Divergent duct

Velocity: decreasing Pressure: increasing Temperature: increasing

Principle

b) Convergent duct

Velocity: increasing Pressure: decreasing Temperature: decreasing

Example: Typical axial flow compressor outlet casing

Principle Example: Flow through turbine nozzle guide vanes

Figure 2 Subsonic Airflow through Divergent and Convergent Ducts


Flow increases to -

Velocity increasing

Pressure decreasing

Velocity further increasing

Pressure further decreasing

Figure 3 Supersonic Airflow Through a Convergent/Divergent Nozzle (Venturi)

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Air intake

\

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Propelling nozzle

L

Compression Combustion Expansion Exhaust

"C Feetis psi

Figure 4 Typical Airflow in a Single-spool Turbojet Engine

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For training use only


a) 2-stage centrifugal flow (turboprop engine)

Low-pressure compressor High-pressure compressor

\

b) Twin-spool axial flow (turboprop engine)

Low-pressure High-pressure compressor compressor

. --

c) Twin-spool turboshaft engine with free-power turbine

Low-pressure compressor Reverse flow

< z z F

Figure 5 Airflow in Engines without By-pass


EASA Part-66 Training Handbook Module 14 LlN K & LEARN

a) Twin-spool axial flow (by-pass turbojet engine with low by-pass ratio)

Low-pressure By- pass By-pass air mixing compressor High-pressure flow with t b n x h a m t compressor

1 I gas stream

b) Triple-spool axial flow (front-fan turbojet engine with high by-pass ratio)

Low-pressure Intermediate pressure comDressor compressor

c) Axial flow (contra-rotating propfan with free-power turbine)

Contra-rotating propfan,

Compressor

d) Twin-spool axial flow (contra-rotating rear fan with free-power turbine)

Contra-rotating fan Low-pressure A

Figure 6 Airflow in Engines with By-pass



15.11.4 Full-authority Digital Engine Control

15.11.4.1 Introduction

History

In the years since World War II, jet engine controls have evolved significantly, providing ever-increasing functionality and exploiting successive advances in technology. Simple but bulky hydro-mechanical controls have given way to hydro- mechanical computers, analogue electronics and today's digital electronic controls, including full-authority digital engine controls commonly referred to as 'FADEC system'.

Note: Some manfacturers name their FADEC system 'fully automated digital electronic control' system

Future

Refer to Figure 1

As shown in Figure 1, future FADECs are expected to have additional functions, such as performance seeking control, redundant control and condition monitoring, as well as advanced schedule control and multivariable robust control, achieving higher engine efficiency, safety, reliability, maintainability and longer life. It is further considered that the future engine control will be integrated with flight control to make an integrated aircraft control system.

Engines with their many control variables have certain degrees of freedom in generating the necessary thrust. This freedom enables the engines to prepare appropriate control modes for various missions. For example, selecting control variables that make fuel consumption minimum is appropriate from an economical point of view, while selecting control variables making the turbine inlet temperature lowest is appropriate from the viewpoint of extending engine life. By selecting control variables that make a certain parameter maximum or minimum, multivariable control achieves the most advantageous engine operation in terms of economy, safety, life and environment.

Optimizing Engine Performance

Refer to Figure 2.

The advent of FADEC established new standards in safety, functionality and engine handling across the flight envelope. The engine electronic control (EEC), the heart of the FADEC, is capable of total powerplant management from engine start to maximum power. It provides a powerful airframe interface for engine control, parameter display, health monitoring and maintenance functions.

The EEC has allowed a very significant reduction in aircraft wiring between the engine and airframe system. Gone are the heavy conventional harnesses and dedicated cockpit instrumentation. In their place, serial digital links (or: data busses) carry electronic signals. Simplified, weight-reduced interfaces replace complex hydro-mechanical and pneumatic controls and their heavy mechanical links to the cockpit.

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The workload of the flight crew has also been reduced. The crew no longer has to make constant adjustments to the engines during the start cycle or through take off, climb and descent. Instead they can take advantage of the EEC's 'set and forget' power management. This aid to carefree handling lets crew members concentrate on other in-flight priorities. Crews no longer need to respond in the traditional way to faults and failures. The fault-tolerant FADEC identifies problems with unprecedented precision for immediate in-flight correction or for post-flight diagnosis.

In addition to engine components, the EEC's diagnostics monitor the other engine accessories.

An engine- or accessory-related fault gets announced immediately to the aircraft central maintenance system by the EEC. Only those failures which require crew intervention are annunciated in the cockpit.

Maintenance crews can access and decode faults on the ground after the flight. The EEC also records faults and stores them in its memory, frequently providing more diagnostic detail than is available from the central aircraft system. Consequently, remembering specific problems no longer depends on the ability of personnel to keep them in mind or write them down. The EEC also contains self-diagnostics to monitor the regulating system's own performance. These features all contribute to rapid fault diagnosis and repair.

Furthermore, FADEC systems provide improved specific fuel consumption by regulating the engine with greater precision and flexibility. Close control of turbine temperature, tip clearance, over- and under-fuelling and optimized idling speeds all contribute to this achievement.

The powerful FADEC computing system has allowed the full potential of the jet engine to be exploited. It maximizes engine efficiency, enabling faster automatic starting sequences and optimizing engine thrust in the cruise mode. Engine life is maximized by maintaining the powerplant within red -ine and transient-operating limits at all times.

15.11.4.2 FADEC Application for Turbines

Airbus A 320 FADEC System

Refer to Figure 3.

The A320 FADEC system eliminates the problems experienced with hydromechani- cal engine control systems. It is fully integrated with the electronic flight control system (EFCS) and the automatic flight system (AFS). The duplicated FADEC systems control the engines, replacing the now redundant fuel control units (FCU).

The thrust levers, located on the centre pedestal, have 5 detents, informing the FADEC of the current phase of the flight. The FADEC accordingly controls engine thrust, providing automatically the optimum output and therefore, minimum fuel burn.

The FADEC performs the following functions: full engine fuel control Thrust setting compulation engine limit protection automatic start sequencing/monitoring thrust reverser controllfeed back



flight deck indication 17 parameters for engine health monitoring

capacity to accept additional sensors.

Refer to Figure 4.

Each engine is controlled by 2 FADEC systems which are called FADEC A and FADEC 9. All signals between the FADEC systems and the engine and between the FADEC systems and the aircraft are completely redundant. The FADEC systems are interconnected by cross-channel data link (CCDL). These buses are used to transmit engine data and FADEC status between the 2 FADEC systems.

Each FADEC receives command signals from the control pedestal and from the powerptant control panel and sends a command signal to the fuel pump metering unit (FPMU) torque motor which meters the fuel flow to the engine in order to reach the fan spool speed calculated by the FADEC thrust management section (N1 request),

The air data computer (ADC) system provide the ambient and airspeed data which are used by the FADEC system to calculate the maximum available thrust (N1 target) for each selected thrust rating mode. The thrust lever modulates linearly between 'IDLE' and 'THRUST SET' position.

In addition. the FADEC systems command the compressor variable geometry (CVG) actuators in order to optimize the compressor efficiency and compressor stall margins.

Both FADEC systems alternate in the powerplant control. While one FADEC controls the powerplant, the other remains in 'standby' mode. The standby FADEC monitors all inputs, performs all computations and performs built-in-test and fault detection, but the output drivers (fuel flow and CVG control), which command the engine, are powered off.

The active FADEC is alternated on each engine ground start in order to minimize the probability of latent failure within the powerplant control systemlaircraft interface. The selection logic resides within the FADEC systems.

Transfer from active FADEC to standby FADEC may be accomplished either automatically, in response to a detected fault, or manually through the FADEC selector knob located on the overhead panel. The manual selection overrides the automatic selection of the controlling FADEC unless the manually selected FADEC is not capable of controlling the engine.

Each FADEC is connected to one of the 2 FADEC systems on the opposite engine via an internacelle data bus. Through this bus, the FADEC systems communicate the information necessary to implement thrust reverse interlock and automatic take-off thrust control system (AlTCS) functions.

All measured powerplant control parameters, control system faults and status information are presented on the displays of the engine indicating and crew alerting system (EICAS).

15.1 1.4.3 Global Online Engine Diagnostics

In the past, if an operator was having problems with the engine or FADEC units, he had to either send them to the factory, or at times, a technical representative was dispatched to the customer in the field to diagnose and fix the problems.



Utilizing the latest online technology and the capabilities of the FADEC system, the engine and FADEC can now be diagnosed and adjusted directly from the factory to anywhere in the world.




LINK & LEARN

(c) by Link & Learn Aviation Training GmbH 30178

Control signals

/ I I I I I I I

HMU / I

. --

Tfuel flow

t- Starter air Monitoring signals (optional)

Legend: VSV = variable stator vane HPT = HP turbine VBV = variable bleed valve CC = combustion chamber

Figure 2 FADEC System for Gas Turbine Engines

Figure 3 A 320 FADEC System


To other engine

I ECAS k Data bus I 1 ADC Data bus

Cockpit discretes

Trust lever

Thrust lever angle

Aircraft discretes

Max. take-off Max. continuous

Max. climb Max. cruise Engine start Engine stop

7

Take-off data store Take-off data increase Take-off data decrease

Anti-ice FADEC reset

Alternate FADEC select

4

3 0

2

I

FADEC ID

~ircraft Starter shut-off

-

jumpers Parity

relays

Data bus From other engine

ECS off

Cross-channel

C

metering unit

FADEC I3

- Engine sensors

A

( 30 (Main) I permanent I

Ignition relay

magnet 1-1 alternator I

Ignition exiter

A

Interface with Interface with the engine the aircraft and other FADEC systems

Figure 4 FADEC Block Diagram



15.1 4.1 Typical Engine Monitoring instruments

15.1 4.1.1 Pressure Measurement

Manifold Pressure Gauge

Refer to Figure 1.

The mechanism of a manifold pressure gauge contains a Bourdon-type measuring element.

This element is essentially a length of metal tube, specially shaped to give it an elliptical cross-section, and formed like the letter C. One end of the tube is sealed, while the other end is left open and fixed into a boss so that it may be connected to a source of pressure and form a closed system.

When pressure is applied to the interior of the tube, there is a tendency for the tube to change from an elliptical cross-section to a circular one, and also to straighten out.

Refer to Figure 2

The manifold pressure is indicated by a pointer on a scale at the front of the instrument.

Oil Pressure Measurement

Refer to Figure 3.

The engine oil supply system provides lubrication and cooling for various engine parts. Failure of the oil supply system will result in an engine failure.

To provide a warning of imminent engine failure, the oil supply to critical areas must be monitored by means of an oil pressure measuring system.

Oil pressure measurement may be done by the direct method (with or without a transmitter). Alternatively an electrical transmission method (via pressure switches) may be used in conjunction with an indicator.

Refer to Figure 4.

Oil pressure indicators are usually calibrated in pounds per square inch (psi), bar or kg/cm2.

A direct-reading oil pressure indicator of the Bourdon type consists of the following main components:

case with pressure inlet connection

Bourdon tube

gear and pinion devices

dial scale (calibrated in psi, bar or kg/cm2).



Functional Principle of a Direct-reading Oil Pressure Indicator

Refer to Figure 5.

Engine oil enters the transmitter casing and surrounds the capsule which is filled with a special fluid.

When the oil pressure increases the capsule is compressed. This increases the pressure of the special fluid which, in turn, is transmitted via a capillary tube to the Bourdon tube.

With increased pressure of the special fluid the Bourdon tube tends to straighten out. Thereby the pointer moves, via a coupling element, over the indicator scale.

Pressure Switches

Refer to Figures 6 and 7

A pressure switch unit consists of the following main parts: case with pressure and electrical connection open metal capsule electrical contact assembly actuating arm external warning light system.

The pressure switch consists of a housing, which is divided into 2 chambers by a diaphragm. One of the chambers is connected to the oil tube behind the filter, the second line is connected to the breather pressure.

If the pressure drops under an adjustable limit, the diaphragm actuates the switch. Then the warning light, usually located at the instrument panel, illuminates.

15.14.1.2 Temperature Measurement

Oil Temperature

The oil temperature indicating system is fitted to measure the temperature of the engine lubrication oil of all types of engine (piston and gas turbine engine).

The engine oil supply is important to the operation of the engine. A rise in oil temperature does not only cause a decrease of the cooling of engine parts but a decrease in the lubricating properties of the oil.

Refer to Figure 8.

The oil temperature measurement system in its most simple form, consists of a temperature sensing element (oil temperature bulb of the resistance type) and the oil temperature indicator (moving-coil indicator).

The temperature bulb is located at the delivery side of the oil pressure pump. The bulb is electrically connected to the indicator and to the aircraft's DC supply. A temperature switch may also be installed to operate a warning light.

Very often the oil temperature indication is integrated in a double- or triple-indication instrument:

oil pressureloil temperature oil pressure/oil temperature and cylinder head temperature.



Principle of Operation

Refer to Figure 9,

The resistance-bulb method makes use of the fact that the electrical resistance of most metals increases as their temperature rises.

A temperature resistance bulb contains a resistance element. The materials generally used for such elements are platinum, nickel or alloys such as nickel manganese (NiMg).

A wire made of one of these metals is wound around a coil former made of an insulating material. This assembly is covered by a metal sheath, called the bulb, that conducts heat quickly.

Refer to Figure 10.

The indicator is a moving coil or a ratiometer instrument suitably calibrated in degrees Celsius ("C).

Exhaust Gas Temperature

Refer to Figure 11.

The exhaust gas temperature (EGT) is a good indication of an engine's operating performance.

Power developed by a gas turbine engine is dependent on the air mass flowing through it as well as on the temperature drop. 'Temperature drop' is the difference between the temperature immediately in front of and behind the turbine. It is a measure of the energy extracted from the airstream by the turbine.

Maximum power can be obtained by establishing optimum temperatures. Therefore operating conditions need to be controlled carefully.

Various types of thermocouples are used to sense the exhaust gas temperature, depending on the engine type and systems which require these temperature measurements.

Refer to Figure 12.

The EGT measurement system of an aircraft monitors the exhaust gas temperature of each engine separately.

EGT indicators are normally calibrated in degrees Celsius although some indicators may still be found calibrated in degrees Fahrenheit.

Dual EGT Indication System of a Piston Engine

The instrument (Figure 11, detail a)) indicates the exhaust gas temperature of each engine separately. The 2 scales represent a relative EGT measurement, i.e. all mixture settings are determined in relation to peak temperatures (indicated by adjustable pointers or by asterisks).

The instrument contains 2 independent moving-coil millivoltmeters and indicates the thermocouple voltage (in mV) of the EGT elements. This is a function of the engine's intake fuellair mixture, as a change in mixture causes a change of the exhaust gas temperature.



Exhaust Gas Thermoelement

Refer to Figure 13.

A chromel alumel element is contained in the metal casing of the EGT thermocouple, the ends of which are brazed or welded together ('hot junction').

Refer to Figure 14.

Variation in exhaust gas temperature produces a millivolt current in the thermoelement, which is proportional to the temperature differential between the hot and the cold junction of the thermocouple.

EGT Indication System of a Gas Turbine Engine

The EGT indicator (Figure 11, detail b)) measures the temperature of the exhaust gas when it leaves the turbine exhaust section. The round panel-mounted instrument is a millivoltmeter which is supplied with voltage from the engine's thermocouple ring.

In simple exhaust gas temperature measuring systems, a group of probes are connected via extension leads to the moving-coil millivolt-meter in the indicator.

The moving-coil indicators are calibrated in degrees Celsius. They are usually calibrated with a specified external resistance value.

Exhaust Gas Thermoelement

Refer to Figure 15.

The EGT sensing element of a gas turbine engine is usually a thermocouple ring. The thermocouple ring is fitted to the rear flange of the outer casing of the turbine exhaust section and consists of an annular tube with (usually) 9 probes (thermocouples).

Each probe contains a chromel alumel thermocouple. The thermocouples are connected in parallel. When exhaust gases flow through openings in the probes, a voltage is generated and applied to the EGT-indicator(s).

lnterturbine Temperature

The interturbine temperature ( I T ) indication system (Figure 11, details c) and d)) provides the pilot with an indication of the engine temperature in the zone between the compressor turbine and the power turbine stator. The indicator is equipped with a pointer scale which indicates the IlT in "C.

Refer to Figure 16.

A typical sensing system consists of a wiring harness (twin leads) incorporating a terminal block, 2 busbars and 9 (sometimes 8) individual chromel alumel thermocouple probes connected in parallel by the busbar arrangement.

The probes stick out from individual, threaded bosses in the power turbine discharge airstream. The hot exhaust gases of the engine generate a voltage in the thermocouples which is proportional to the exhaust gas temperature. This voltage is applied to the IlT indicators via the wiring harness and the aircraft wiring.

The IlT indicator is a millivoltmeter which is supplied with voltage from the thermocouple wiring harness via a compensating resistor.



15.14.1.3 Speed Measurement

Engine Speed Indication

Refer to Figure 17.

The measurement of the engine speed (in revolutions per minute (rprn)) enables the pilot to control the accurate performance of the engine (in conjunction with other engine parameters, e.g. inlet manifold pressure, EGT).

A typical engine indicating system consists of the following main components: rprn indicator

rprn transmitter

rpm wiring (independent of aircraft power supply).

Engine speed indicating systems are used in all types of powered aircraft.

Refer to Figure 18.

The engine speed indicating system of a piston engine measures and displays the speed of the engine's crankshaft.

Refer to Figure 19.

The engine speed indicating system of a gas turbine engine (and of turboprop engines) measures and displays the speed at which the power turbine shaft rotates.

The engine speed indicating system of a helicopter measures and displays the speed of the engine's crankshaft or of the power turbine shaft, and also the speed of the main rotor.

The indicating instruments are normally called 'tachometers'. The method most commonly used for measuring the engine speed and transmitting the value to the indicator is by electrical means. However, in several types of general aviation aircraft, mechanically operated tachometers (with a flexible drive) are used.

There are 2 main types of electrically operating speed measurement system:

one using a generator as the transmitter

the other using a tacho probe.

The engine speed may either be shown in actual revolutions per minute or as a percentage (whereby 100 % may only be exceeded during take-off and landing).

Speed Measurement System with a Generator

Refer to Figure 20.

A speed measuring system incorporating a generator consists of a 3-phase AC generator acting as the rpm transmitter

a 3-phase AC synchronous motor inside the rpm indicator

the necessary wiring.

The generator contains a permanent magnet rotor which rotates inside the stator. It is driven by the engine gearbox.




As the rotor rotates past the stator windings, 3 phases of alternating electromotive force (EMF) are induced into the windings. These 3 phases are 120" apart from each other.

The frequency of the induced EMF is directly proportional to the rotor's speed. The magnitude of the EMF depends on the strength of the magnet and the number of turns on the phase coils.

Since the rotor speed is determined by the engine speed at a fixed ratio, the frequency of the induced EMF is a measure of engine speed.

The output voltage of the generator is supplied to the 3-phase coils of the indicator to produce a resultant magnetic field. This resultant field rotates at a speed corresponding to the generator output frequency.

As the stator field in the indicator rotates it causes a torque to rotate the rotor into the same direction and at the same speed.

Refer to Figure 21.

When the engine is running, the 3-phase AC voltage produced by the rpm transmitter is applied to the rotor of the synchro in the indicator. The rotor shaft rotating inside the bell of the metering system carries a magnetic disc. This disc induces eddy currents which, in turn, produce torque. The torque turns the pointer of the indicator, which indicates the speed of the corresponding engine.

Speed Measurement System with a Tacho Probe

Refer to Figure 22.

A speed measuring system incorporating a tacho probe consists of the following main components:

tacho probe (electromagnetic pick-up) acting as the rpm transmitter

torque synchro inside the rpm indicator

the necessary wiring.

This type of system is used to measure and indicate compressor speeds of an engine. In some turbo fan engines, the speed of the fan can also be measured.

One advantage of a probe is its ability to provide separate electrical output signals to other systems (if required). Another advantage is that a probe has no moving parts.

A tacho probe is made of stainless steel and hermetically sealed (airtight seal) to prevent any foreign matter (dirt, dust) from entering. It consists of a permanent magnet, a pole piece and a number of coils wound around a central ferromagnetic core. A flange on the probe allows it to be mounted at that position inside the engine where the speed measurement is to be made, i.e. where the probe's pole pieces are close to the teeth of a certain gear wheel. This gear wheel is known as a 'phonic wheel'. It is driven at the same speed as the compressor shaft or fan shaft.



The indicator contains a signal-processing module (servo-amplifier), a motor (torque synchro) and a feedback circuit consisting of a potentiometer and a buffer amplifier.

Normally, there are 2 spring-loaded pointers which indicate the rpm. A power supply module provides the necessary AC and DC supplies for the indicator.



The probe and gear wheel act as a magnetic flux switch that induces electromotive forces into the sensing coils. These forces are directly proportional to the compressor's (or: fan's) speed.

The probe's permanent magnet produces a magnetic field around the sensing coils. As the teeth of the gear wheel pass the pole pieces of the probe, the flux intensity through each pole piece varies, according to the varying width of the air gap between the pole pieces and the gear wheel teeth.

As long as the intensity of the flux changes, an EMF is induced into the sensing coils. Its amplitude depends on the rate of change of flux intensity.

The top of each gear tooth and the bottom of each tooth space are flat. At these points, the intensity of the flux does not change and, as a result, the induced EMF will be zero. However, because the sides of each gear tooth are angled, there is a change in flux intensity as long as the angled sides of the teeth pass the pole pieces.

The induced EMF reaches its maximum when the maximum rate of change of flux occurs.

The frequency of the changes of the EMF depends on the number of teeth in the gear wheel (or fan blades).

The signals from the probe, which are to be used for speed indication, are supplied to the indicator's signal processing module. In the module, the signals are added to the outputs of the servo-potentiometer and the buffer amplifier. After summation, the signals are amplified by the servo-amplifier in order to cause the torque synchro to rotate the pointers.

At the same time, the wiper of the servo-potentiometer is moved in order to control the summation of signals to the servo-amplifier.

In the case of a power supply or signal failure the main pointer is returned to an off-scale position by a preloaded spring.



15.14.1.4 Quantity Measurement

Fuel Flow Indicating Systems

Refer to Figure 23.

A fuel flow indicating system measures the rate of the fuel flow from the fuel tanks to the engine. Fuel flowmeter systems are used in large piston-engined aircraft and in all gas-turbine-engined aircraft. An accurate knowledge of the fuel mass left in the fuel tanks and of the rate at which the fuel flows to the engine enables the pilots to calculate the remaining travelling range of the aircraft during flight.

Basically all fuel flow measuring systems consist of a transmitter and an indicator although some systems also use an amplifier. The transmitter measures and transmits an electrical signal proportional to the rate of flow of the fuel.

Refer to Figure 24.

A fuel flow indicator is electrically operated and may be a synchro, a moving coil or may be servo-operated.

Indicators may be used in groups (one for each engine), or a single indicator may be used with a selector switch. It is also possible that several transmitters can be connected to one indicator in order to display the total flow for all engines.

Fuel flowmeters are usually calibrated in literslmin, pounds per hour or kilograms per hour. They may have totalisers showing the total weight of the remaining fuel or the fuel used since the flight started.

There are many fuel flow measurement systems in use on different aircraft, however all of them fall in one out of the 2 following categories:

independent fuel flowmeter

integrated flowmeter.

Independent Fuel Flowmeter

Refer to Figure 25.

The independent fuel flowmeter system has a rotating vane transmitter, an indicator and requires 28 V DC for operation.

The transmitter is constructed of a cast metal body with inlet and outlet connections, joined together via a spiral-shaped metering chamber. Inside the metering chamber there is a metering vane pivoted in such a way that it can be rotated by the fuel passing through the chamber.

There is a small gap between the edge of the vane and the chamber wall. Because of the shape of the chamber the area of this gap increases as the vane is moved from its 'zero' position.

The vane is mounted on a shaft, one end of which is attached to a calibrated spring. Therefore, the greater the fuel flow, the further the vane is rotated. The other end of the metering vane shaft sticks out through its bearing. It carries a ring-type magnetic coupling between the vane and the electrical transmitting unit.


EASA Part-66 Training Handbook Module 14 LlN K & LEARN

The transmitting unit may be either an AC synchro or a precision potentiometer. The shaft of the potentiometer carries a 2-pole bar magnet which is located inside the ring magnet. Interaction of the magnetic fields provide a magnetic lock (or coupling) so that the wiper arm (or: synchro rotor) can follow the movement of the metering vane without friction.

A spring-loaded by-pass valve located on the top of the transmitter, provides an alternative path for the fuel if the fuel pressure rises above a preset limit. This may happen if the vane becomes jammed in a 'low-flow' position.

A damping chamber (not shown in Figure 25) is mounted to one side of the transmitter and connected to the metering chamber by a small bleed hole. Inside the damping chamber there is a counterweight and a circular vane secured to the same end of the metering vane shaft as the control spring.

Fuel is allowed to fill the damping chamber through the bleed hole and completely immerse the counteweight assembly. Any oscillation of the metering assembly and transmission element is therefore compensated for by liquid damping (by the fuel) and is unaffected by the fuel flow.

The indicator is a moving-coil milli-ammeter which carries one or more pointer(s). Additionally, there is an amplifier in the indicator case which amplifies the signals from the fuel flow transmitter.

Where an indicator makes use of synchronous transmission, the pointer is operated by the rotor of a receiver synchro.

Integrated Flowmeter System

Refer to Figure 26.

The integrated flowmeter system of the true-mass-flow type consists of 3 principal units:

flow transmitter

electronic relay (or computer, not shown in the Figure)

indicator.

Its operation depends on the following principle:

'The torque required to accelerate a fluid to a given angular velocity, is a measure of the flow rate of the fluid's mass. '

When fuel is flowing the construction of the system causes relative angular displacements between a constant-speed impeller and a sensing rotor (drum) located behind the impeller. Inductive-type pick-offs sense the angular displacements in terms of signal pulses proportional to the flow rate. These pulses are supplied to the indicator via an amplifier or a computer.



Fuel Flow Transmitter

The transmitter consists of an alloy body which contains an AC electric motor driving an impeller assembly at a constant speed. Channels through the impeller allow the fuel to flow through at an angular velocity which is equal to that of the impeller.

A 'drum' called reaction turbine (or: torque sensing rotor) is mounted behind the impeller, separated by a stationary disc. The drum is rotated by the fuel flowing through the turbine channels. The drum is restrained by a calibrated spring which is attached to the end of the turbine shaft. On the same end of the turbine shaft a synchro transmitter is mounted, which measures the amount of the turbine deflection against the spring tension.

Fuel Flow Indicator

The indicator contains a transformer, an amplifier, a servo-motor, a tachogenerator and a pointer mechanism, which operates via a magnetic drum cap assembly. A digital counter mechanism is mechanically connected to the servo-motor.

Fuel Quantity Indicating Systems

Fuel quantity indicating systems measure the amount of fuel left in the aircraft's fuel tanks. The scales of fuel quantity indicators may be calibrated in gallons, pounds, liters or kilograms. Some may even have 'full' and 'empty' indications only.

Refer to Figure 27.

Fuel quantity indicating systems vary in construction and operating principles depending on the type of aircraft and the type of fuel system fitted. The 2 main methods use the principle of electrically transmitting the signals from the tank to an indicator. They are the:

float-type system

capacitance-type system.

Float-type Fuel Quantity Indicating System

Refer to Figure 28.

This system is mainly used in light aircraft and helicopters. The system comprises 2 units:

a transmitter (called 'tank unit' or 'sender unit'), which is located in the fuel tank an indicator located in the cockpit.

Refer to Figure 29.

A tank unit consists of a mechanical float assembly which controls an electrical transmitting device. The float may be made of cork (specially treated to prevent fuel absorption), or it may be in the form of a lightweight metal cylinder suitably sealed.

The float is attached to an arm pivoted so as to permit angular movement. This movement is transmitted to an electrical element consisting of either a wiper arm and a potentiometer or a Desynn type of transmitter (synchro).




As the fuel level changes, the arm with the float moves through certain angles and re-positions the wiper arm (or the brushes) in order to vary the resistance. Thereby the flow of current to the indicator is varied as well. These variations in current flow cause the deflection of a moving coil (or: rotor) within the indicator, which, in turn, positions a pointer over a suitably calibrated scale.

Indicators depend on the type of transmitter used. Where the transmitter is of the Desynn type, the indicator must be a Desynn synchro as well. If the transmitter is a potentiometer, the indicator may be a ratiometer or an ammeter.

Capacitance-type Fuel Quantity Indicating System

Refer to Figure 30.

A system indicating the fuel quantity by measuring capacitance is usually installed in high-performance aircraft. In its basic form a capacitance-type fuel-gauge system consists of a variable capacitor located in the fuel tank, an amplifier and an indicator.

The complete circuit forms an electrical bridge, which is continuously being re-balanced as a result of the differences between the capacitances of the tank capacitor and a reference capacitor. The signal produced is amplified and operates a motor, which positions a pointer to indicate the capacitance change of the tank capacitor (and so the change in fuel quantity).

Refer to Figure 31.

A basic capacitance fuel quantity indicating system consists of 3 units:

tank unit

amplifier

E indicator.

The tank unit is a variable capacitor and consists of a pair of closely spaced metal cylinders mounted one inside the other. The unit extends from the top to the bottom of the tank; its actual construction depends on the shape of the fuel tank.


The 2 parallel cylinders, partly immersed in the fuel, form a capacitor. The capacitance depends on how much fuel is in the tank. A mixture of fuel and air, which varies according to the fuel consumption, acts as the di-electric constant. This causes the capacity of the capacitor to vary as well.

The amplifier amplifies the signal changes produced by the tank unit, in order to operate a motor in the indicator.

The indicator contains a motor which drives a pointer (or a digital mechanism). At the same time the motor moves a wiper arm over a potentiometer used for balancing the bridge (as feedback).

The individual components and units are interconnected by means of coaxial cables. These cables are normally supplied in ready-made lengths. They are of a known, constant capacitance value because capacitance-type systems are sensitive to any changes in the capacitance of the tank unit circuits.



15.14.1.5 Engine Vibration Moni tor ing and Indicating System

Engine vibrations are unwanted but unfortunately they cannot be eliminated entirely (even in turbine engines, which do not have any reciprocating parts). Vibrations can only be kept at the lowest possible level.

During operation, vibrations above acceptable levels may always occur as a result of mechanical faults. E.g. a turbine blade may crack or creep, or an uneven temperature distribution around the turbine blades and rotor discs may occur. Either of these failures will cause an unbalanced condition of the main rotating assembly.

Refer to Figure 32.

An engine vibration monitoring and indicating system consists of the following main components:

vibration pick-up (linear-velocity detector) auxiliary monitor (amplifier) indicator (moving-coil micro-ammeter).

The vibration pick-up unit is mounted at a right angle to the engine's axis. Some systems may use 2 pick-up units: one monitoring the vibration level around the turbine section and the other around the diffuser section.

Refer to Figure 33.

The vibration indicator is an instrument that continuously monitors and displays the amplitude of turbine engine vibration. The instrument is calibrated in such a way that, if the amplitude reaches the critical level of vibration, the pointer will register half-scale deflection. This 'critical level of vibration amplitude' depends on the type of turbine engine that is being monitored.


Refer to Figure 34.

The pick-up unit is a linear-velocity detector that converts mechanical energy of vibration into an electrical signal. The magnitude of this signal is proportional to the energy of vibration. The conversion is done by means of a spring-supported permanent magnet, which is suspended in a coil attached to the interior of the case.

As the engine vibrates, the pick-up unit and the coil move with it. The magnet, however, tends to remain fixed in space because of the influence of inertia. The motion of the coil causes its windings to cut the field of the magnet inducing a voltage in the coil. This voltage (or: the signal) is amplified and transmitted via an electrical filter network and a rectifier to the indicator.

Auxiliary Monitor

The filter and amplifier networks produce a highly stable highpass response for frequencies from 14 Hz to 1,000 Hz. They provide a signal whose amplitude is proportional to the vibration amplitude. This signal is then rectified by a diode pump circuit and applied as follows:

for display in the indicator as a measure of low vibration amplitude to supply the vibration level information to a warning circuit.

The warning circuit is designed to trip at a preset level (normally at the critical level of vibration amplitude) on the indicator scale. When reaching this preset level the warning circuit causes the indicator and other remote warning lamps to light up.



lndica tor

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The indicator is a micro-ammeter, which is vertically mounted. It continuously displays the vibration level as a relative reading (typically from '0' to '5'). The indicator is calibrated so that the critical level of vibration amplitude corresponds to a pointer's position at approx. half-scale of the indicator.

Furthermore, the indicator incorporates a reference datum pointer that is actuated by a lever arm mechanism. This pointer can manually be set by means of a slotted adjusting control to indicate the normal level of vibration amplitude. This provides a visual reference to the pilot for checking any changes of the vibration amplitude during flight.

Power Supply

The AC power enters the auxiliary monitor at the plug connector. Power is supplied directly to the indicator through the auxiliary monitor socket connector and the indicator connector.

Power supply boards inside the units convert the 11 5 V ACl400 Hz power supply to a stabilised 24 V DC for amplifier supply and to 0.5 V ACl400 Hz for test purposes.



connections

h

B m

Springs

Figure 1 Functional Principle of a Bourdon Tube



Figure 2 Manifold Pressure Indicator (Example)



I LOW OIL PRESSURE (OR FILTER BYPASS) j Oil

Fuel

Figure 3 Oil System Monitoring



Capsule with \ transmitting fluid \

System fluid pressure inlet

Figure 5 Oil Pressure Indicating System (Example)



a) Warning system Bus +28 V DC

- - -

Control box warninglcaution

- - - - - - - - - Engine

Switch low Power press oil supply g

ru

Pressure connection

b) Presssure switch

Oil tube behind filter

Breather pressure \ Diaphragm

Figure 6 Oil Pressure Warning System and Pressure Switch



Electrical connector

Pre a d j ~

Figure 7 Oil Pressure Switch



Temperature bulb (resistance type)

Figure 8 Oil Temperature Indicating System



a) Oil temperature indicator (moving-coil type)

n , Pointer Hairsorina 8 - \ 11 Bridqe piece - .

Coil and former

Adjusting device

Balance Core arm

Shunt

b) Oil temperature transmitter (resistance-bulb type)

Plug Union nut receptacle I

1 Former Bulb

/ / Calibrating Resistance 2-pin socket Contacting (or: balancing) coil element

strips

Figure 9 Oil Temperature Transmitter and Indicator



Temperature bulb

Figure 10 Electrical Oil Temperature Measuring Circuit



a) Dual EGT indicator (piston engine)

b) Servo-operated EGT indicator (gas turbine engine)

c) ITT indicator (turboprop engine)

d) ITT indicator (gas turbine engine)

Figure 11 EGT and ITT Indicators (Examples)



EGT probe 1

Figure 12 Tapping Points of EGT Probes



Front adju screws

sting

\

Chromel l e d +, LH engine Alumel lead -, LH engine

I

Connection

Connection

Probe

Clamp

Figure 13 EGT Indication System (Piston Engine)


m D cn D w nl ?'

p-Thermoelement *- Balance wire -+- Lead wire -4 Q) Q) (probe)

1 I I I -I 7 ru.

I 2. I s ca

Hot junction connector

Indicator

Reference line Cold junction

Figure 14 EGT Indication System (Functional Diagram)


Chramel alurnel / wiring harness

Figure 15 Thermocouple Carrier of a Gas Turbine Engine


block

Chromel terniinal

Threaded coupling

u Probe assembly Jasher

Figure 16 ITT Indication (Schematic Diagram)


rpm indicator

Figure 17 Typical rpm Indicating System (Schematic Diagram)



m

RPM generator

v

Figure 18 Tapping Point for rpm (Piston Engine)



NI (LP compressor) 1 T a -, Q

N2 (HP compressor)

Figure 19 Tapping Points for RPM (Gas Turbine Engine)



Laminated squirrel-cage rotor

I

Figure 20 Functional Principle of an rpm Indicating System



Electrical

Spot-welded connection Core

connector permanent magnet Sensing coils

I I

ar wheel

------//

Reference voltage

Buffer

Tacho probe signal

Figure 22 Tacho Probe and Torque Tachometer (Schematic Diagram)

Signal processing

module


cu

2 115VAC 400 Hz

w

(c) by Link & Learn Aviation Training GmbH 66178

module

- Power supply

m

1 4 V DC

-


Power setting

REVERSE FEATHER

0 0 0

transmitter I++- r - I t - i

From fuel tank

Figure 23 Fuel Flow Indicating System (Schematic Diagram)



Figure 24 Fuel Flowmeters (Examples)



By- pass valve \ m

Fuel outlet

0

Calibrated ~e te r ing vane spring

Fuel inlet

I Metering unit

LINK & LEARN

Remote indicator

Figure 25 Independent Fuel Flowmeter



Restraining spring Annular space

Constant-speed impeller

E A B C D l 115 vac mi - - / E A B C D

To 3-phase power supply

Indicator

Figure 26 Integrated Fuel Flowmeter System



Figure 27 Fuel Quantity Indicators (Examples)



Potentiometer I winding

\ I Electrical connector

Figure 28 Fuel Quantity Indicating System (Float Type)



Transmitter Indicator

Wiper arm Toroidal resistor

Fuel tank Indicator

Figure 29 Float-type Fuel Quantity Indicating System



Indicator

Reference capacity (varies with fuel

+ - - - - - - - - permittivity)

Fixed trim capacitor

I \

Isolating transformer

Legend:

------ Mechanical connections

Electrical connections

Representing capacity of tank units

Figure 30 Fuel Quantity Indicating System (Capacitance Type)



Bus + 2 8 V D C

Tank control circuit breaker

I Indicator

LINK & LEARN

Figure 31 Capacitance-type Fuel Quantity Indicating System



Indicator Amplifier

Figure 32 Engine Vibration Monitoring and Indicating System



Reference pointer

19-way connector

I

Reference pointer

Figure 33 Engine Vibration Indicator (Example)


m14 propulsion

Documents

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pulse jet engine

todays jet propulsion

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whittle engine

piston engine

nene jet engines

modern gas turbine engine