Module 14 B2 Propulsion Final 2014 Notes

Faculty of Transport Engineering Technologies

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Module 14 Propulsion

These notes are intended for training guidance only and are not to be used as an authoritative

document for use in the civil aviation industry. In all cases, reference must always be made to

Annual

Review Completed by Date

2011 C. Gibson 05/08/2011 2012 C Gibson 31/08/12 2013 C. Gibson 14/05/13 2014 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031

Amendment and Annual Review Record

Amendment

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14.1 Turbine Engines ... 4

14.1.1 Constructional Arrangement and Operation of Turbojet, Turbofan, Turboshaft and Turbopro peller Engines ... 14

14.1.2 Electronic Engine Control and Fuel Metering Systems (FADEC) ... 22

14.2 Engine Indicating Systems ... 42

14.2.1 Exhaust Gas Temperature / Interstage Turbine Temperature Systems ... 42

14.2.2 Engine Speed ... 51

14.2.3 Engine Thrust Indication: Engine Pressure Ratio, Engine Turbine Discharge Pressure or Jet Pipe Pressure Systems ... 54

14.2.4 Oil Pressure & Temperature ... 57

14.2.5 Fuel Pressure, Temperature and Flow ... 59

14.2.6 Manifold Pressure ... 65

14.2.7 Engine Torque ... 66

14.2.8 Propeller Speed ... 68

14.3 Starting And Ignition Systems ... 70

14.3.1 Operation Of Engine Starting Systems And Components ... 70

14.3.2 Ignition Systems And Components ... 77

14.3.3 Maintenance Safety Requirements ... 82

Acronyms and Abbreviations ... 83

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14.1 Turbine Engines

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 Kitty Hawk, America with their historic flight in 1903. Since that early date, aircraft have developed steadily and, by 1939, aircraft speeds of 464 mph were being achieved by production aircraft. Aircraft with piston engines and propellers could climb to 56,000 feet and fly distances of up to 7,000 miles non-stop. 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 engine and propeller combinations suffered a loss in performance at high forward speeds and altitudes; clearly a new type of aircraft propulsion unit was needed if aircraft performance was to advance even more; thus the jet engine was born.

Principle Of Jet Propulsion

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". 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. The force which accelerates the steam reacts in the opposite direction on the engine, moving the engine away from the accelerating column of steam. A garden sprinkler uses a similar principle.

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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. Theoretically, all that is needed to produce useful thrust is a tube, with an inlet, some means of introducing and burning fuel and an exhaust. This is known as a ramjet, illustrated in fig.2.

The ramjet has no moving parts, all the reactive thrust being available to propel the aircraft to which it is fitted.

engine to burn the fuel, the aircraft has to be travelling at 300 knots or more.

This problem is overcome in the Turbojet (or Gas Turbine) engine by using exhaust gas to power a turbine, which in turn, drives a compressor fitted in the intake. 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 suitable for aircraft propulsion.

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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 t h a t has made modern high performance aircraft possible. The function of any propeller or gas turbine is to produce a propulsive thrust by accelerating a mass of air (or gas) rearwards. Let us now apply Newton’s Laws of Motion to see how Thrust is produced.

 In order to accelerate the air, a FORCE must be applied (Newton’s 1st Law).

 The acceleration is proportional to the applied force.

 There must be an equal and opposite REACTION (Newton’s 3rd Law) i.e. a forward acting force which is the Thrust.

The thrust obtained is proportional to the mass of air passing through the engine and to the velocity increase (acceleration) of the mass of air flow, i.e.:-

FORCE (Thrust) = MASS x ACCELERATION

The same amount of propulsive thrust can be obtained by either:

 Accelerating a large mass through a small increase in velocity.

 Accelerating a small mass through a large increase in velocity.

Thrust

A jet engine produces thrust in a manner similar to that of a piston engine / propeller combination but, whilst the propeller gives a small acceleration to a large mass of air, the turbine engine gives greater acceleration to a smaller mass of air flow. This point is illustrated in fig. 4.

Application of Principles

In addition to Newton's third law of motion, it is necessary to study mass flow of matter, Bernoulli's theorem and subsonic diffusion to understand how a gas turbine engine produces useful thrust.

Mass Flow of Matter

To understand how matter behaves when moving in a duct it is necessary to consider the mass flow of the matter. This is defined as the quantity of matter flowing in unit time, the mass flow may be expressed in lb/sec, kg/sec, or in any other convenient units.

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Mass Flow through a Ducted System

When a steady stream of air passes through a steady flow machine; such as a gas turbine engine, operating at fixed rev/min and air inlet density; the mass flow at any point in the system is of a constant value. If we consider the machine to be an open-ended duct, we find that the mass flow per second will depend on the density of the air and the volume flowing per sec.

Therefore: - Mass flow = density  area  velocity. This is known as the 'continuity equation' and it is true for any steady flow system regardless of changes in the cross-sectional area of the duct.

Bernoulli's Theorem

This theorem states that the sum of the pressure and kinetic energies in a fluid moving inside a duct is constant, even though pressure energy can be converted to kinetic energy and vice versa. This theorem can be applied to the relationship between pressure and velocity existing in the air flowing through a duct, such as a jet engine.

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Pressure Energy

In gas or fluid the pressure energy is more often called 'static pressure' and it can be defined as the pressure that would be felt by a body which was submerged in the medium (gas or fluid) and moving at the same velocity as the medium.

Kinetic Energy

This kind of energy is more often called 'dynamic pressure' and this term is used to define the extra pressure created by the movement of the medium. Dynamic pressure is proportional to ½ mass  velocity2 (i.e. ½mv2).

Continuity Equation and Bernoulli's Theorem Incompressible fluid

The combined effects of the continuity equation and Bernoulli's theorem are shown in the diagram below, when a steady flow of incompressible fluid flows through a duct of varying cross sectional area. This shows that:-

Mass flow remains constant as the cross-sectional area of the duct (and velocity) change.

Total pressure remains constant, but static pressure (PS) changes as area (and velocity) change.

Compressible Fluid (Atmosphere)

Compressible fluid flow refers to the air flow through a gas turbine engine and, because the air is compressible, flow at subsonic speeds causes a change in the density of the air as it progresses through the engine.

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The air entering the duct at section A consists of air pressure (P1) and velocity (V1), then as the air enters the increased area of the duct at B it will spread out to fill the increased area and this will cause the air flow to slow down (continuity equation) and give a change in velocity to V2. The static pressure of the air will increase (Bernoulli's theorem) to become P2 in the wider section of the duct and, because air is compressible, the air density will also increase as it is compressed by the rise in pressure in section B of the duct.

Nozzles and Ducts

The energy changes throughout the gas turbine engine are effected by means of nozzles and ducts of various shapes and sizes

A duct which has a decreasing cross-sectional area is known as a CONVERGENT duct, the inlet area is greater than the area at the exit.

When air flows through such a duct, it increases in velocity and the static pressure is reduced. In other words, an increase in velocity is accompanied by a drop in pressure; there is also a drop in temperature. How the convergent duct as applied to gas turbines is shown in the diagram below.

If the duct has an increasing cross-sectional area it is said to be DIVERGENT and will convert kinetic energy into pressure energy.

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turbine where velocity is to be reduced and pressure increased, there is also an increase in temperature. A typical position for a divergent duct is shown in the diagram below. When air is compressed by this process it is called subsonic diffusion and it is a principle that is used extensively in jet engine design.

Later, we shall see the various changes that occur in velocity and pressure during the passage of an air stream through practical gas turbine engines. During these energy changes, the temperature will always follow the pressure. It is essential to note that energy changes through these ducts will NOT conform to the above if the following

 when additional energy is imparted to the air, e.g. by heat

 when energy is being extracted from the air, e.g. by doing work

 when the velocity of air passing through the duct reaches sonic speed. If this occurs the nozzle is said to be "choked".

NOTE The speed of sound is directly related to temperature.

When choking occurs, there can be no further increase in velocity until the temperature of the air is increased.

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and a Piston Engine.

Working Cycle of a Gas Turbine Engine

Heat engines convert the heat energy of the fuel into mechanical work. Piston engines and gas turbines are heat engines, both using air as the working fluid. In the Piston Engine the power output is intermittent, whereas in the Gas Turbine it is continuous.

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 is accelerated by heating. This means that the velocity of the air is increased before it is finally emitted in the form of a high velocity jet. The working cycle of a gas turbine is called the Brayton

Cycle. The working cycle on which the gas turbine engine

functions is, in its simplest form, represented by the P/V diagram (fig 10).

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BRAYTON CYCLE

 1-2 Compression: Work is done on the air in the Compressor resulting in a rise in its pressure and temperature and a decrease in its volume.

 2 to 3 Heat Energy (Combustion) increases the temperature and volume while the pressure remains virtually unchanged, hence the term:

CONSTANT PRESSURE CYCLE

 3 to 4 Expansion through the Turbine where energy is extracted (to drive the Compressor), resulting in a decrease in pressure, and temperature, whilst the volume of the gas increases.

 4 to 1 The air returns to ambient pressure ready for the cycle to start again.

The expansion process is completed through the Jet Pipe Nozzle, which produces a high velocity jet, the reaction to this providing the thrust, the gas finally reducing back to atmospheric pressure.

Note: The term ‘Constant Pressure’ only applies if the

engine is operating under a constant set of conditions. Even so, in practice there is a slight drop in the combustion system due to turbulence caused by the actual combustion itself.

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Changes in Temperature, Pressure & Velocity

The changes in temperature, pressure and velocity of the gases through a gas turbine engine are illustrated in the following diagram. The efficiency with which these changes are made will determine to what extent the desired relations between pressure, volume and temperature are obtained. The more efficient the compressor, the higher is the pressure generated for a given work input, i.e. for a given temperature rise of the gas. Conversely, the more efficiently the turbine uses the expanding gas, the greater is the output of work for a given temperature drop in gas.

During the passage of the air (gas) through the engine, aerodynamic and energy requirements demand changes in its velocity and pressure. For example, during

compression, a rise in the pressure of the air is required with no 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 cause the turbine to rotate.

Also, at the propelling nozzle, a high velocity is required, for it is the change in momentum of the air that provides the thrust on the aircraft. Local decelerations of gas flow are also required - for example, in the combustion chambers to provide a low velocity zone for the flame.

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14.1.1 Constructional Arrangement and Operation of Turbojet, Turbofan, Turboshaft and Turbopropeller Engines

There has been a great deal of development of gas turbine engines since the Whittle gas turbine engine first appeared. This engine was fitted with a single-sided centrifugal compressor, which had a low compression ratio (about 4 : 1 ). To increase this, it would have been necessary to increase the diameter of the compressor and, therefore, the frontal area. This would, in turn, have increased the weight considerably. A two-sided centrifugal compressor was an improvement, but similar penalties could not be avoided. The demand for greater power output, efficiency and flexibility led to further improvements in design, particularly by Rolls Royce with the axial flow, single and twin spool type compressors, the turbo-fan engine (including the RB series), up to the present Trent engine. Although we shall be discussing the components of the turbine engine later, it can be stated that all gas turbine engines have an intake assembly, a compressor assembly, a combustion assembly, a turbine assembly and an exhaust assembly.

The basic principles of each component remain the same, but the path of the air through the engine varies according to the design. A straight flow system is usual as it provides an engine with a small frontal area and is suitable for use of by-pass and ducted fan principles. We shall now introduce common types of gas turbine engines.

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mentioned, overcoming the disadvantage of its’ low compression ratio, high specific fuel consumption and large frontal area. The axial flow engine (air path parallel to the centre line of the engine) with its various stages of compression in the same casing (the Avon Mark 1 engine made by Rolls Royce had 12 stages of compression) gave higher compression ratio and a considerable improvement in performance and lower fuel consumption, as well as a smaller frontal area. Let us now look at this in more detail, together with any disadvantages of axial compressors.

The Axial Flow Compressor

The compressor consists of a series of discs which carry blades of an aerofoil section (the Rotor). The rotor is surrounded by a casing which houses fixed blades, also of an aerofoil section (the Stators). A row of stator blades is located behind each row of rotor blades to form a compressor stage. Several stages go together to make up the compressor. An additional set of stators is located prior to the first set of rotor blades. These are the Intake Guide Vanes. On some compressors the angular setting of these vanes is automatically controlled to suit varying airflow conditions. This ensures air enters the first stage compressor rotors smoothly and at the optimum angle.

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rise in pressure and simultaneously imparting a high rotational or ‘swirl’ velocity. The air next enters the first stage stators where a proportion of the high kinetic energy is converted into a further rise in pressure by the divergent passages between the stators, with a consequent fall in velocity. These passages also correct the ‘swirl’ imparted to the air by the rotor blades and present it at the correct angle for entry into the rotor of the next compressor stage, where the process is repeated.

Each stage produces an increase in pressure. This is relatively small (1.1 to 1.2 times the inlet pressure) equally produced by the rotor and the stator. The final row of stators act as straighteners to remove any ‘swirl’ from the air before it enters the combustion system.

As the air density increases through the compressor from inlet to outlet, the cross-sectional area of the air annulus is progressively reduced. This maintains a constant net axial velocity and also maintains the pressure rise from the low to the high pressure end of the compressor.

Although the swirl velocity increases and decreases through rotor and stator vanes respectively, the axial velocity through all the stages remains approximately constant.

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when the airflow meets the rotor blades at the optimum angle of attack. This angle is determined when the engine is designed dependant on required mass flow, pressure ratio and the compressor r.p.m. range. Compressor stages are matched to give optimum efficiency at the high r.p.m. range of operation (take-off and climb).

Surge in Axial Flow Compressors

When operating at lower speeds the air meets the first stage rotors at too great an angle of attack. The airflow pattern across the blades will break down and the blades will stall, in the same way as any other aerofoil. When this occurs the stall may spread downstream to the subsequent compressor stages until the whole airflow pattern breaks down.

At low engine speeds another factor can affect the rear stages of the compressor. Due to the reduced pressure ratio, the air attempts to occupy a greater volume. As the space available is controlled by the volume of the annular space, the result is a ‘choking’ of the later compressor stages. When choking occurs, the velocity of the inflow through the compressor will decrease until the first stage stalls. This would be followed by subsequent stages until all stages have stalled.

High pressure at the compressor outlet will now cause a reversal of airflow towards the compressor inlet. This causes the compressor to SURGE. Choking is now relieved and normal airflow is restored until choking re- occurs and the pattern repeats.

If surging continues it may cause severe turbine a n d c o m p r e s s o r damage so the engine must be shut down immediately.

Surging can also be caused by over fuelling. If the engine is at low r.p.m. and the throttle is opened gradually, there will be a gradual increase in gas temperature and velocity, resulting in increased power at the turbine and the engine will accelerate.

If, however, the throttle is opened too rapidly there will be a rapid high fuel delivery rate. Acceleration response will lag, due to the large inertia of the rotating assembly. There will be a fast increase in gas velocity through the NGVs and turbine, causing choking at the turbine. Air velocity through the compressor will reduce until the first and successive stages stall causing the engine to surge. To overcome this some turbo-jet fuel systems are fitted with an over fuelling control to regulate the fuelling rate to match the lag in acceleration of the compressor.

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speed range anti-surge devices are fitted to the engine. These include Bleed Valves which open or close automatically in response to an r.p.m. signal. At low r.p.m. they will open and release the excess volume of air (due to the low pressure ratio) to atmosphere. This prevents choking of the later stages, maintaining air velocity and thus eliminating compressor stall.

Inlet Guide Vanes can be automatically adjusted such that the airflow into the compressor continues to meet the first stage rotors at the correct angle of attack, dependent on engine r.p.m. and air intake temperature. Some engines also incorporate one or more stages of variable incidence stators to alleviate surge problems.

Twin Spool Axial Flow Compressor Engine

This engine has a compounded compressor assembly in which the compressors are driven by separate turbines, through co-axial shafts; the only connection between the two rotating assemblies is the gas stream. This allows each half of the compressor to be run at its most efficient speed. The low pressure assembly rotates at a lower rev/min and accepts air from the intake and passes it to the high pressure drum, resulting in higher pressures and increased stability.

The advantages to be gained are:

 Higher compression ratio.

 Better airflow stability.

 Lower specific fuel consumption.

 Greater flexibility in operation.

 Reduction in the possibility of 'stall' and 'surge'.

 More rapid acceleration possible.

 Easier starting.

 Greater power at altitude.

The twin spool axial flow compressor engine is illustrated below. This example shows a turbo-propeller engine. The rear (low speed) turbine drives the front (low speed) compressor and also, through a reduction gearbox, a propeller

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By-Pass Twin Spool Axial Flow Compressor Engine

The by-pass engine was developed to permit the use of higher turbine temperatures to obtain higher thrust. About half of the low pressure air is passed through the annular by-pass duct surrounding the high pressure compressor assembly and combustion system to re-join the hot gas stream after the turbine. This results in higher combined flow of cooler, slower gases to atmosphere. The advantages to be gained in addition to those mentioned are:

 Higher propulsive efficiency.

 High thermal efficiency.

 Better power / weight ratio. (smaller, lighter high pressure compressor, combustion system and turbine).

 Reduced fire risk and heat loss.

 Reduced noise level.

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Turbo-Fan Engine

This is a high-ratio by-pass engine with a large diameter front fan driven by the low pressure turbine, and operating within a cowl to provide a separate low velocity, high-mass air flow; the air is ducted to flow concentrically with the hot jet and does not mix in an exhaust unit as in the medium by-pass engine. The front fan may have more than one stage and the by-pass ratio is 3:1 or more. As the high pressure compressor is required to pass only a proportion of the total mass flow, both the compressor and combustion system are of smaller and lighter construction than those engines already mentioned. An illustration of a turbo-fan engine is shown here.

Some turbo-fans have three concentric shafts with an intermediate compressor (as fitted to the Rolls Royce RB178-51, the Rolls Royce / Turbomeca RB172 Adour, the Rolls Royce RB203-01 (Trent) and the Rolls Royce RB211). The advantages to be gained are as for the by- pass engine mentioned previously, but with greater propulsive efficiency and much lower specific fuel consumption (SFC) due to the large mass flow and lower jet velocities.

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Turboshaft

A gas turbine engine that delivers power through a shaft to operate something other than a propeller is referred to as a turboshaft engine. Turboshaft engines are similar to turboprop engines. The power take-off may be coupled directly to the engine turbine, or the shaft may be driven by a turbine. The free turbine located downstream of the engine turbine. The free turbine rotated independently being connected to the main engine only by the hot stream of gases. This principle is used in the majority of turboshaft engines currently produced, and is being used extensively in helicopters, ships, electric generators etc.

Turbopropeller

The turboprop (turbo-propeller) engine is a combination of a gas turbine and a propeller. They are basically similar to turbojet engines in that both have a compressor, combustion chamber(s), turbine and a jet nozzle, all of which operate in the same manner on both engines. However, the difference is that the turbine in the turboprop engine usually has more stages than that in the turbojet engine. In addition to operating the compressor and accessories, the turboprop turbine transmits increased power forward, through a shaft and a reduction gear train, to drive the propeller. The increased power is generated by the exhaust gases passing additional stages of the turbine. The exhaust gases also contribute to engine power output through jet reaction, although the amount of energy available for jet thrust if considerably reduced.

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14.1.2 Electronic Engine Control and Fuel Metering Systems (FADEC)

An understanding of mechanical fuel control will help you understand what the Full Authority Digital Engine Control System (FADEC) does. The thrust of a turbo jet is controlled by varying the amount of fuel burnt in the combustion system, and in order to operate to safe temperature limits, the amount of fuel that is burnt must be governed by the amount of air that is available at the time. The air supply is dependent upon the RPM of the compressor and the density of the air at its inlet, so under a constant set of atmospheric conditions the RPM of the compressor is an indication of the engine thrust. The pilot has control of the fuel flow to the combustion system and is able to select any compressor RPM, between ground idling and maximum permissible which is required for take- off conditions, by the operation of a cockpit lever.

Atmospheric conditions can vary resulting in changes of air density at the compressor inlet. A reduction in air density will cause a reduction in the amount of air delivered to the combustion system at a selected RPM, with a consequent increase in the combustion chamber temperature. If the fuel flow is not reduced, a rise in compressor RPM will occur, accompanied with overheating of the combustion and turbine assemblies.

An increase in air density will result in an increase in the amount of air delivered to the combustion system at a selected RPM, and unless the fuel flow is increased a reduction in RPM will occur.

Changes in air density at the compressor inlet are caused by:

 Effects of Altitude. The density of the air gets progressively less as the altitude is increased, therefore less fuel will be required in order to maintain the selected RPM.

 Effects of Forward Speed. The faster the aircraft flies then the faster the air is forced into the aircraft intake. A well designed aircraft intake will slow down this airflow, converting its’ kinetic energy into pressure energy, so that it arrives at the compressor inlet at an optimum velocity with an increase in pressure and hence density. This is known as Ram Effect, and plays an important part in the performance of a turbo-jet. Within certain limits the greater the ram effect, the greater the air mass flow and more fuel can be burnt at the selected RPM, producing more thrust.

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Purpose of the Engine Fuel System

The purpose of the engine fuel system is to deliver to the combustion system, in a readily combustible form, the correct amount of fuel over the whole operating range of the engine, under the control of the pilot.

Layout of Typical System Components

The diagram opposite illustrates the layout of components of a representative fuel system. Some of the components in the system are fitted to the aircraft and other are fitted to the engine. The aircraft mounted components are:

 Fuel Tanks. These store sufficient fuel for the aircraft's designed flight duration.

 Booster Pumps. These ensure a constant supply of fuel at low pressure to the inlet of the engine driven HP Fuel Pump.

 Low Pressure Cock. This isolates the engine fuel system from the aircraft fuel system for servicing requirements.

Note: These aircraft mounted components will be dealt

with in greater detail during the Aircraft System Phase.

The engine mounted components are:

 Low Pressure Filter. Fuel enters the engine fuel system at the LP filter. A low pressure switch is often fitted to the filter case and this operates a warning light in the cockpit if the fuel pressure on the outlet side of the filter falls below a certain value.

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pump receives filtered low pressure fuel at its inlet and raises the pressure sufficiently to cause the fuel to flow through the burners into the combustion chambers at the correct rate determined by the throttle position and atmospheric conditions.

 Throttle. The throttle is set manually by the pilot and its position determines the amount of fuel delivered to the burners and hence the engine speed and thrust. Movement of the throttle schedules the HP pump to deliver fuel at the appropriate rate. Note: The throttle levers are aircraft mounted components but the throttle is mounted on the engine.

 Barometric Pressure Control. The BPC is sensitive to throttle movements and engine air intake conditions. Its purpose is to relay fuel flow requirements to the HP fuel pump in response to changes in throttle position, and to modify that fuel flow in response to varying engine air intake pressures, thus maintaining automatically the selected RPM.

 Acceleration Control Unit. When the throttle is opened the BPC will schedule the HP fuel pump to increase fuel flow. The HP fuel pump is able to respond to the demand very quickly, but because of the inertia of the compressor, fuel flow tends to rise faster than the airflow. To prevent compressor surge due to over fuelling, the ACU is sensitive to air/fuel ratio and limits the rate of over fuelling during the early stages of a rapid engine acceleration.

 High Pressure Cock. The fuel flow to the burners passes through the HP cock which is manually operated from the cockpit. The cock has two positions, fully open to permit engine running, or fully closed to stop the engine by shutting off the fuel supply to the burners.

 Pressurising Valve. The pressurising valve ensures that the fuel pressure in the burner manifolds is high enough for efficient burner operation.

 Burners. The purpose of the burners is to present the fuel into the combustion chamber in a readily combustible form.

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A device sensitive to fuel flow to the burners and air pressure at the engine intake, schedules a change in fuel pump output in response to signals of varying air intake pressure. Common names for such components are:

 Barometric Pressure Control (BPC)

 Barometric Fuel Control Unit (BFCU)

 Altitude Sensing Unit (ASU)

Manual Operation

The pilot selects the required RPM by movement of a cockpit lever which is mechanically connected to a throttle valve in the engine fuel system. The result of opening the throttle causes the fuel pump to schedule a greater fuel flow to the burners. The gas temperature in the combustion chamber rises and the acceleration of the gases through the turbine increases. This results in a higher compressor RPM and a greater airflow, thus providing an increase in thrust.

During engine acceleration, a device sensitive to fuel flow to the burners and air delivery from the compressor, limits fuel pump output in response to signals of excessive fuel to air ratio during the early stages of engine acceleration. As the engine accelerates the same device schedule an increase in fuel pump output in response to signals of increasing compressor air delivery. Common names for such components are:

 Acceleration Control Unit (ACU)

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driven by the engine via a train of gears. This limits the output to the fuel pump in response to a signal of maximum engine RPM. This component, known as the Max RPM Governor, is adjustable, and is often incorporated in the fuel pump.

Note: Components controlling the fuel flow may be

mounted on the engine, whether individually, or grouped together in one main unit known as the Fuel Control Unit.

FADEC

FADEC (Full Authority Digital Engine Control System) is the name given to the system that controls the engine on modern Gas Turbine Engines. This part of these notes discusses the common features of FADEC and also the different applications used by the large commercial passenger aircraft engine manufacturers, Rolls Royce (RR) and General Electric(GE) and their derivatives IAE and CFM.

FADEC replaces the hydro-mechanical fuel control systems as exemplified by the Rolls Royce Spey or JT8D. It can also be utilized to increase the engines’ efficiency, by incorporating such devices as Variable Stator Vanes and Automatic Turbine Rotor Clearance Control.

The benefits of FADEC are:

1. Substitution of Hydro-mechanical control system reduces weight and hence fuel consumption.

2. Automation brings reduced pilot workload. 3. Optimised engine control reduces maintenance

and optimises fuel consumption

4. Optimised airflow control allows the engine to work nearer the surge line, thus increasing thrust, whilst reducing the chance of surge or flameout.

A FADEC system consists of a Central Processor Unit called an Electronic Engine Control (EEC) or an Engine Control Unit (ECU), a Hydro Mechanical Unit (HMU) and sensors.

The Central Processor Unit, for the purposes of these notes will be referred to as the ECU.

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1. Analogue signals from electrical sensors.

2. Digital signals, usually on an ARINC 429 Data Bus, from aircraft computers such as the Air Data Computer (ADC), Thrust Management Computer (TMC) and Flight Management Computer (FMC). 3. Thrust lever signals are transmitted by Rotational

Variable Differential Transformers, mechanically connected to a conventional thrust drum, which is moved by the Manual Thrust Lever and the Auto Thrust Servo Motor.

4. Pressure inputs - apart from those received from the ADC. Po and PS3 (Intake and Compressor

Delivery Pressure) signals are tapped directly into pressure transducers located within the ECU.

5. Feedback signals from any moving mechanical device, such as Thrust Reverser and Variable Bypass Valves, utilise Linear or Rotary Variable Differential Transducers (LVDTs or RVDTs).

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Overview

Engine Control Unit (ECU)

The ECU is a dual channel processor that computes all functions of the FADEC system, based on its inputs and stored data, and then commands the HMU to take appropriate actions. Every second a typical ECU can monitor 200 measurements from more than 40 sensors to ensure the engine runs safely and efficiently. The ECU also provides ARINC 429 data to the Flight Management Computer (FMC), Thrust Management Computer (TMC) and EICAS (Boeing) or ECAM (Airbus) cockpit display computers.

Hydro Mechanical Unit (HMU)

The HMU provides an interface between the electrical analogue output from the ECU and the fuel. It is achieved by an Electrical Hydraulic Servo Valve (EHSV) actuating a Fuel Metering Valve (FMV), thus controlling fuel supply to the burners. In addition the HMU will have EHSVs controlling fuel muscle pressure to Variable Stator Vanes (VSVs) and Variable Bleed Valves (VBVs), if fitted.

Principles

FADEC Interface with Aircraft

Inputs to FADEC

 Thrust Lever Resolver- Two analogue signals come from the thrust lever resolvers. They represent the Thrust Lever Angle (TLA), this angle is, however, most often called the Throttle or Thrust Resolver Angle (TRA).

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A310 and its’ PW 4000 series engine. Thrust required is input to the Thrust Control Computer (TCC); either by the pilot via the Thrust Rating Panel, or from the Flight Management System when engaged in Performance (Vertical Profile) Mode.

The Auto-throttle Actuator drives the Throttle Control Levers to the appropriate position, for the thrust required, via the Coupling Units and Dynamometric Rods. It also drives the Resolver Unit, positioning the Thrust Resolver Angular position (TRA) to the thrust required as shown below.

TRA signals are sent to the TCC for positional feedback and to the FADEC as demand signals. The FADEC monitors actual thrust and compares this with thrust demanded (TRA). If these differ, the FADEC controls the FMU to bring them into line.

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Figure below shows the flow paths for a CFM 56-5 Engine, which is a typical FADEC engine. Please note the Following: 1 FADEC is a very useful tool for gathering information for a condition monitoring system. Customers can choose whether to have Condition Monitoring for their system, therefore the sensors required are customer options and are marked *.

2 TLA stands for Thrust Lever Angle. This signal is received from the RVDT fitted to the thrust lever drum. However this angle is sometimes quoted as the TRA Throttle or Thrust Resolver Angle)

3 The ECU is powered by its’ own alternator or by aircraft 28v DC Aircraft Bus for Starting, Testing and Maintenance. 115 VAC aircraft power is required for the AC igniter circuit.

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The Engine-Control Unit (ECU)

The ECU is a dual channel processor housed within a single container, however all hardware within the container is partitioned into the two channels. Normally mounted on the fan casing cooling is either by natural Fan Case Cooling Air or directly by a dedicated Fan Air Ducting.

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ECU Architecture

Dual Channels

The FADEC System is fully redundant, built around two independent control channels. Dual inputs, dual outputs, and automatic switching from one channel to the other, eliminate any dormant failure.

Channel Selection

The ECU will always select the "healthiest" channel as the

priority list contains critical faults such as processor, memory or power failures, as well as other failures that involve a channels’ capability to control the FMV, VSV, or VBV torque motor(s).

During engine run status, each channel within the ECU will determine whether to be in the active state or standby state every 30 milliseconds based on a comparison of its own health and the health of the cross-channel. Either channel can become active if its health is better than the cross-channels health. Likewise it will become standby if its health is not as good as the cross-channels health. If the two channels have equal health status, the channels will alternate on each engine shutdown and the standby channel will become the active channel on the next start.

Channel Transfer

Assuming the opposite channel is of equal or greater health, channel Active/Standby transfer will occur after the engine has been run above 76% N2 and subsequently shutdown (N2 less than 35%).

Electrical Inputs

All command inputs to the FADEC system are duplicated. Only some secondary parameters used for monitoring and indicating are single (e.g. the EGT input on the CF6

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To increase the fault tolerant design, the parameters are exchanged between the two control channels via the cross channel data link.

Pressure Inputs

Pressure tappings from the engine are plumbed directly into the ECU, either discretely to each channel or a single tapping that is split within the ECU and then sent to discrete channel transducers.

Hardwired Inputs

Information exchanged between aircraft computers and the ECU is transmitted over digital data buses. In addition signals are hardwired directly from the aircraft where a computer is not used. (Thrust Reverser feedback via RVDT's or TLA via an RVDT)

Outputs

All the ECU outputs are double, but only the channel in control supplies the engine control signals to the various receptors such as torque motors, actuators or solenoids. Further information on output signal receivers can be found in the HMU section.

The ECU is equipped with BITE, which provides maintenance information, and test capabilities via an aircraft mounted component called Multifunction Control Display Unit (MCDU, Airbus) or Propulsion Interface Monitoring Unit (PIMU, Boeing).

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monitors during operation. In addition operation of a ground test switch powers up the ECU which carries out a real time ground test. For Boeing airframes the ECU stores faults in the ECU volatile memory until the aircraft lands. On landing the faults are streamed to the PIMU which holds the fault until a BITE test is carried out. There is a PIMU for each engine. An EICAS message will advise maintenance staff to carry out this procedure even if the pilot has not noticed the problem.

AIRBUS faults will be stored in the MCDU in real time. BITE interrogation is airframe specific and cannot be covered in a generic FADEC publication.

Using the BITE system, the ECU can detect and isolate failures in real time and hence allows switching of engine control from the faulty channel to the healthy one.

Fail Safe Control

If a standby channel is faulty and the channel in control is unable to ensure one engine function, this control is moved to a fail-safe position. For example, if the standby channel is faulty and the channel in control is unable to control VBV position, the valves are operated to the open position.

Main Interfaces

To perform all its tasks the ECU interfaces with, aircraft computers, either directly or via the Engine Interface Monitoring Unit (EIMU). Principle among these, are the aircraft Left and Right Air Data Computers which supply data, notably Ambient Temperature (Tamb); Total Air Temperature (TAT); Static Pressure (PSO) and Total

Pressure (PT). All of these are required to determine that

the thrust commanded remains constant for the ambient conditions and that thrust and EGT limits are not exceeded.

Limits Protection

The ECU has a dual channel limit protection section comprising max limits for N1, N2 and N3 (RR only) In addition various max limits are protected depending on the system, most commonly Compressor Delivery (Burner) Pressure. (Ps3).

Thrust Regulation

Thrust regulation on high bypass engine is calculated using ADC inputs to calculate the required fuel to provide the commanded thrust. The thrust is measured in terms of N1 speed or EPR (RR Trent). For the EPR engine in the event of EPR signal failure it reverts to control by N1.

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compressor (HP2 or HP3) governor located within the HMU.

Thrust Control Modes

Systems vary, therefore below are three typical systems:

CF6 FADEC Control Modes

In the event that an ADC signal is lost then the ECU will use the opposite channel signal. In the event that the channels inputs do not agree as to which signal is accurate then the ECU will revert to an alternate mode using the last known ambient pressure signal. This is also known as the soft reversionary mode.

The soft reversionary mode can cause throttle stagger as the other engine is still operating in the normal mode. To prevent this, the ECU mode switches can be pushed for both engines, to select hard reversionary mode, which means they are using the fixed corner point ambient temperature for that engine. Because Tamb may be higher than corner point there is now a danger of over-boosting the engine. The pilot will always throttle back before selecting hard reversionary and be aware of max N1 indication to prevent over-boosting or over-temping the engine.

R.R. Trent FADEC Control Modes

The primary thrust control loop uses EPR. In the event that EPR computation is impossible then the ECU reverts to the N1 mode where N1 is used to control thrust. In the N1 mode Auto Throttle is no longer available.

CFM 56 FADEC Control Modes

The engine operates in one of three thrust modes, AUTO - MEMO – MANUAL. Entering/exiting these three modes is controlled by inputs to the Engine Interface Unit (EIU).

Auto Thrust Mode

The auto thrust mode is only available between idle and Max Climb Thrust when the aircraft is in flight. After take-off the throttle is pulled back to the max climb position, the auto thrust system will be active and the Automatic Flight system will provide an N1 target to provide either:

 Max Climb Thrust.

 An Optimum Thrust.

 A Minimum Thrust.

An Aircraft Speed (Mach Number) in association with the auto pilot.

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The Memo Mode is entered automatically from Auto mode if the N1 target is invalid. One of the instinctive disconnect buttons on the throttle is activated. Auto thrust is disconnected by the EIU. In the memo mode, the thrust is frozen to the last actual N1 value and will remain frozen until the throttle lever is moved manually, or auto thrust is reset.

Manual Thrust Mode

This mode is entered any time the conditions for Auto or Memo are not present in this mode. Thrust is a function of throttle lever position.

Date Entry Plug & Wiring Harnesses

The Data Entry Plug is mounted on the FADEC channel A housing on the upper left side. It provides engine trim data for thrust rating, optional equipment configuration and EPR/thrust relationships to the ECU for the specific engine only.

The Data Entry Plug is a contact connector. Jumper wire connections in the plug provide the ECU specific information about the engine. This information is used for fuel scheduling and engine rating calculations. The plug is configured for the specific engine characteristics. It is attached to the engine by a lanyard and remains with the engine if the ECU is changed.

Note: Some engine types have separate Rating and Identification plugs (e.g. the GE CF6-80).

Wiring Harness

The Wiring Harness is routed as required around the engine and to the strut connections. It provides input and output signal paths for the FADEC.

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Power Supplies

Permanent Magnet Alternator (PMA)

A dual coil Permanent Magnet Alternator driven from the External or Accessory Gearbox powers the ECU. The dual output is fed independently to the two Channels. The PMA can provide all power requirements once the engine is running above 15% N2 (N3 for RR Engine).

For engine starting an aircraft 28V DC supply is used. In addition a 28V DC Bus supplies power for ground testing the system and for back up in the case of the primary 28V DC Bus failing.

Aircraft 28 V DC is also always available in the event of PMA supply failing to both channels.

28V DC is applied to the ECU when: The start switch is activated

The Fuel switch is placed to on (for an in-flight windmilling start)

When ground test power is applied

The aircraft supplies a 115V AC 400HZ power source to each channel f o r ignition exciter # 1 and ignition exciter # 2. The inputs are routed to the exciters or terminated within the ECU by switching relays.

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failure then the engine will not start as the exciters can only be powered via the ECU.

Hydro Mechanical Unit (HMU)

Primary outputs from the ECU are directed to the torque motors of the EHSVs located on the HMU and to the torque motor controlling the primary fuel metering valve. The fuel metering subsystem is completely contained in the HMU. The HMU is mounted on the front, right side of the accessory gearbox. It is driven by a mechanical

connection to the gearbox. The HMU responds to electrical signals from the ECU to meter fuel flow for combustion and to modulate servo fuel flow to operate the engine air systems. The HMU also receives signals from the aircraft fuel control system to control an internal high pressure fuel shutoff valve (HPSOV).

There are four external electrical connectors for electrical interfaces with the aircraft and ECU. Four fuel ports connect the HMU with the fuel pump and fuel nozzles. There are five hydraulic connections for control interfaces with the engine fuel and air systems. Each hydraulic interface is controlled by an electro-hydraulic servo valve (EHSV) that varies servo fuel pressure in response to ECU signals.

The fuel connections to the HMU are:

 Fuel inlet from the fuel pump

 Fuel discharge to the fuel nozzles

 Fuel bypass discharge to the fuel pump

 Servo fuel inlet from the servo fuel heater. The hydraulic connections from the HMU are:

 Servo fuel pressure to the low pressure turbine case cooling (LPTCC) valve

 Servo fuel pressure to the high pressure turbine case cooling (HPTCC) valve

 Servo fuel reference pressure to the LPTCC and HPTCC valves

 Servo fuel pressure to the variable bypass valves (VBVs)

 Servo fuel pressure to the variable stator vanes (VSVs).

The electrical connections to the HMU are:

 Fuel control signals from EEC channel A

 Fuel control signals from EEC channel B

 HPSOV solenoid inputs from the fuel control valves

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The HMU has three hydraulic circuits:

 A fuel metering circuit

 A bypass circuit

 A servo control circuit.

The fuel metering circuit controls fuel flow to the fuel nozzles in the engine combustor. It has a fuel metering valve and a high pressure fuel shutoff valve (HPSOV). Unmetered fuel from the fuel pump goes to the FMV. Metered fuel from the FMV goes to the HPSOV. If the HPSOV is open, metered fuel is routed to the fuel nozzles. The bypass circuit is composed of a bypass valve, a differential pressure (delta P) regulator, and an over-speed governor. The fuel pump supplies more fuel than needed for the metered fuel flow. The bypass circuit returns excess fuel to the fuel pump.

The servo control circuit divides the fuel supply from the servo fuel heater into regulated and unregulated servo flows. These flows operate actuators located both inside and outside of the HMU. The circuit has a servo regulating and distribution section and five electro-magnetic servo valves. One of these servo valves supplies servo pressure for FMV control and is discussed below. The other servo

Fuel Metering Valve

A fuel metering valve (FMV) inside the HMU controls fuel flow to the nozzles. The hydraulically driven metering valve is controlled by the fuel metering valve EHSV. The EHSV has two coils, one for each ECU channel. The controlling ECU channel increases current through its EHSV coil to hydraulically open the FMV. If neither coil has power, the FMV closes. The FMV has two position indicating resolvers. One resolver is excited by, and provides a position feedback signal to, ECU channel A. The other resolver goes to ECU channel B.

The differences between an HMU and a Mechanical System are:

 The LP cock is replaced by an Isolation Valve which is controlled by the fire handle in the cockpit.

 The HP cock is replaced by a Pressurising and Shut- off valve which is controlled by the Fuel Control Switch on the Engine start / run lever.

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14.2 Engine Indicating Systems

Introduction

The following notes provide the student with generic information on Engine Indicating Instrumentation as found

on most General Aviation and pre-Electronic

Instrumentation type of aircraft. Information on the Airbus Electronic Centralised Aircraft Monitoring (ECAM) and the Boeing Engine Indicating and Crew Alerting Systems (EICAS) are to be found in Module 5 B1 and B2 notes.

14.2.1 Exhaust Gas Temperature / Interstage Turbine Temperature Systems

Temperature measurement falls into two distinct categories, High Temperature measurement and Low Temperature measurement. High temperature measuring devices measure such things as Exhaust Gas Temperature (E.G.T.) and Cylinder Head temperature. Low temperature measuring devices measure such things as Fuel and Oil temperatures.

There are a variety of ways in which temperature can be measured as follows:

Electrical Type

A change in the temperature of an electrical conductor can cause a change in the resistance of that conductor. Thus measuring the resistance of an electrical conductor can indicate the temperature of that conductor. This is known as the “Resistance” type.

Dissimilar metals when joined together at one end can produce an electrical potential called a thermo E.M.F. This e.m.f. is dependent upon the temperature difference between the junctions, temperature measuring devices using this principle are known as “Thermo-Electric” measuring devices.

Radiation Type

The radiation emitted by a body at any wavelength is dependent upon the temperature of that body. This is known as a body’s “Emissivity”. Thus the temperature of a body can be determined by that body’s “Emissivity”.

The majority of aircraft temperature measuring devices utilise only the Electrical Type of measuring device, which can be divided into two sub-groups dependent upon whether the temperature range to be measured is

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type, whilst high temperature measuring devices utilize the Thermo-Electric Type. Temperature measurement using the Resistance type is known as Resistance Thermometry and the Thermo-Electric measuring types are known as Pyrometry.

Temperature Sensing Elements

The sensing element consists of a resistance coil wound on an insulated former, the ends of the coil being connected to a Two-Pin socket via contact strips. The resistance coil may be made from various materials, e.g. Nickel or Platinum, which possess positive linear temperature coefficients of resistance.

It is most important that the correct bulb is used with a specific indicator to avoid indication errors. Gauges are calibrated for a specific type of resistance wire and are marked accordingly, i.e. Plat. Law or Nickel Law.

Temperature bulbs may be filled with hydrogen to improve their response time. The cable interconnecting the bulb and indicator forms part of the temperature bulbs resistance and should therefore not vary from a specific stated ohm value.

Thermocouples Seebeck Effect

If two dissimilar metal wires are fused together at both ends to form a continuous loop; and the temperature of one junction is raised above the temperature of the other junction; a thermo-e.m.f. is produced, whose value will be directly proportional to the difference in temperature between the two loop ends. This is known as a Thermo- couple, called the Seebeck effect after its’ discoverer.

The Seebeck Effect is utilised when measuring the high temperatures of aircraft engine cylinder heads and jet engine exhaust pipes.

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used to measure the thermo-e.m.f.

Immersion Type Thermocouples

The Immersion type thermocouple is used to measure the temperature of gases. It is typically used as the sensing element of turbine engine gas temperature indicating systems. The Chromel/Alumel hot junction and wires are usually encased in ceramic insulation within a metal protection sheath (typically Inconel), the complete assembly forming a probe that can be immersed in the gas stream at specific points where measurement is required. There are two classifications of Immersion type thermocouples known as Stagnation and Rapid Response types. The classification depends upon whether the probe is to be used with high velocity or low velocity gases.

In pure jet engines the gas velocities are high, so in these engines Stagnation thermocouples are employed.

It will be noted from diagram (a) opposite, that the entry and exit holes (known as sampling holes) are staggered and unequal in size. This allows the gases to slow down and stagnate at the hot junction, allowing the thermocouple time to respond to the change of gas temperature. A typical response time would be 1 - 2 seconds.

SAMPLING HOLES

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measure exhaust gas temperatures of turboprop engines. Since the gas velocities of a turboprop engine are lower than that of a pure jet engine a different type of probe is used as shown in (b). It will be noted that the sampling holes are directly opposite each other and are of the same size. The gases flow directly over the thermocouple allowing the couple to react more quickly. A typical response time would be between 0.5 seconds and 1 second.

Some temperature probes are used to supply more than one system in which case more than one element is required as shown in (c). Insulation of the thermocouple elements from each other is provided by compacted magnesium oxide (MgO), which also serves to maintain the elements in position.

Nozzle Guide Vane Thermocouple

A third type of thermocouple is designed to measure gas temperatures between turbine stages. The hot junction is housed inside a sheath, which is specially shaped to form the leading edge of a stator guide vane and is therefore usually referred to as a Nozzle-Guide-Vane thermocouple. Gases flow over the hot junction, which is positioned between sampling holes of equal diameter as in the rapid response thermocouple. However, since the holes are

and its proximity to the guide vane, the couple response is much slower than the rapid response type.

It is required, in some types of turbine engine to measure the temperature of the engine cooling air. This requires a different design of thermocouple from those discussed previously. The temperature sensor in this case is also a Chromel / Alumel thermocouple element, designed to be positioned over a vent hole and between a mounting boss on the engine and an overheat detector switch.

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It is important to position probes correctly as the temperatures measured relate to engine performance. The ideal position for temperature measurement is either at the turbine blades themselves or at the turbine entry but this presents certain practical difficulties, consequently thermocouple probes are located at the exhaust or jet pipe unit, and between the turbine stages at one end of the stator positions.

For accurate measurement it is necessary to sample temperatures from a number of points evenly distributed over a cross section of the gas flow. This compensates for the fact that differences in temperature can exist between various layers of airflow through the turbine and exhaust unit.

The measuring system therefore consists of a group of at least 5 thermocouples distributed evenly in the gas flow and connected in parallel in order to measure the average temperature condition. This arrangement is known as a 'Harness Assembly' as in the diagram opposite.

Thermocouple Harness Assemblies

A typical example of a thermocouple harness is shown in the following diagram. The 5 probes in this case each contain 2 thermocouple elements; one for temperature indication and one for temperature control. In some engines probes and thermocouple lead junction boxes may be designed as separate units but in the illustration given the probes are welded to stainless steel junction boxes thus forming single items.

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Inconel conduits which are also welded to ferrules at the junction boxes. The leads terminate at a main junction or 'Take off' box, to which the leads of the remainder of the circuits are connected. This allows for easy replacement / removal of the harness.

Cold Junction Temperature Compensation

Since the indicator of any thermocouple system forms the cold junction part of the thermocouple then any change in ambient temperature at the indicator will cause an indication error. For example, if the hot junction temperature remained constant and the ambient temperature of the indicator increased then the temperature difference would decrease resulting in the indicator Under-Reading. Conversely if the ambient temperature of the indicator were to decrease the temperature difference between hot and cold junctions would increase so the indicator would over-read.

There are two methods of compensating for cold junction errors, mechanical and electrical. A typical example of a mechanical method is the bi-metallic strip as shown in the diagram below.

With the indicator disconnected from the thermocouple system the bi-metal spring response to ambient temperature changes at the indicator, an increase in temperature causing the spring to unwind resulting in the hairspring element assembly moving round to indicate an increase in temperature. Conversely, temperature decrease will result in the element indicating a lower temperature. The indicator therefore acts like a direct reading bi-metal type of thermometer.