Module 14 B2 Propulsion Final 2014 Notes

Faculty of Transport Engineering Technologies School of Aeronautical Engineering 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 the current documents for the most up to date information.

Amendment and Annual Review Record Amendment No

Incorporated by

Date

Annual Review 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031

Completed by C. Gibson C Gibson C. Gibson

Date 05/08/2011 31/08/12 14/05/13

School of Aeronautical Engineering 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 Bibliography and Recommended Further Reading ......................................................................................................... 83

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School of Aeronautical Engineering 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.

EASA Module 14 B2 Propulsion

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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School of Aeronautical Engineering Jet reaction is an internal phenomenon and it is not, as sometimes assumed, the result of the jet efflux impinging upon the atmosphere. The jet engine is designed to accelerate a stream of air to an exceptionally high velocity and to obtain useful thrust from the reaction. There are many ways of increasing the velocity of the air but, in all cases, the resultant reaction is the propulsive thrust exerted on the engine. 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.

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.

The ramjet has no moving parts, all the reactive thrust being available to propel the aircraft to which it is fitted. Unfortunately, in order to get sufficient airflow through the EASA Module 14 B2 Propulsion

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School of Aeronautical Engineering In 1941 the Whittle gas turbine engine powered the Gloster E28/39 aircraft and many of the present-day RollsRoyce 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.

EASA Module 14 B2 Propulsion

 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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School of Aeronautical Engineering

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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School of Aeronautical Engineering 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.

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

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 Compressible Fluid (Atmosphere) 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:-

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.

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

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School of Aeronautical Engineering 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.

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.

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.

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

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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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School of Aeronautical Engineering The divergent duct is used at various points in a gas 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.



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

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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.

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 conditions are encountered.

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School of Aeronautical Engineering Figure 9 shows a comparison of operation of a Gas Turbine 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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School of Aeronautical Engineering BRAYTON 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.

 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.

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.

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

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School of Aeronautical Engineering 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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School of Aeronautical Engineering 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.

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. Turbojet

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.

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School of Aeronautical Engineering This engine replaced the centrifugal type already 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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School of Aeronautical Engineering The rotor blades displace the air rearwards, producing a 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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School of Aeronautical Engineering The multi-stage axial flow compressor is most efficient 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.

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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 reoccurs 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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School of Aeronautical Engineering To ensure stable operation of the compressor over a wide 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.

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The advantages to be gained are: 

Higher compression ratio.

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Better airflow stability.

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Lower specific fuel consumption.

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Greater flexibility in operation.

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

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More rapid acceleration possible.

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Easier starting.

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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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School of Aeronautical Engineering   

Reduced fire risk and heat loss. Reduced noise level. Greater thrust yield from reheat.

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).

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School of Aeronautical Engineering

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 bypass 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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School of Aeronautical Engineering 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.

EASA Module 14 B2 Propulsion

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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School of Aeronautical Engineering 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 takeoff 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.

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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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School of Aeronautical Engineering 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.

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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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School of Aeronautical Engineering  Engine Driven High pressure Pump. The HP fuel 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.

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

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.

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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.

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

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

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School of Aeronautical Engineering Automatic Operation 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.

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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)  Air Fuel Ratio Control Unit (AFRCU)

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School of Aeronautical Engineering Maximum RPM is automatically controlled by a device 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.

The benefits of FADEC are:

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

2. Automation brings reduced pilot workload. 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.

3. Optimised engine control reduces maintenance

FADEC

4. Optimised airflow control allows the engine to

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.

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and optimises fuel consumption 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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School of Aeronautical Engineering A FADEC system has the following inputs: 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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School of Aeronautical Engineering Overview

Principles

Engine Control Unit (ECU)

FADEC Interface with Aircraft

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.

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).

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.

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School of Aeronautical Engineering The previous diagram shows the interface between the 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.

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.

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.

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School of Aeronautical Engineering

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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School of Aeronautical Engineering

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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School of Aeronautical Engineering 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.

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

Channel Selection The ECU will always select the "healthiest" channel as the Active channel based on a fault priority list. The fault

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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 engine). Page 33 of 71

School of Aeronautical Engineering

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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School of Aeronautical Engineering The ECU performs a self-test on power up, and selfmonitors 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.

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.

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.

Limits Protection

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.

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).

Fail Safe Control Thrust Regulation 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.

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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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School of Aeronautical Engineering As a backup there is a mechanical high pressure compressor (HP2 or HP3) governor located within the HMU. Thrust Control Modes

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.

Systems vary, therefore below are three typical systems: CFM 56 FADEC Control Modes 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 overboosting 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.

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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 takeoff 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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School of Aeronautical Engineering Memo Mode 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.

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.

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.

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School of Aeronautical Engineering 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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School of Aeronautical Engineering It should be noted that if the ECU has a double channel 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.

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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 HPSOV position indication outputs to the EEC

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School of Aeronautical Engineering General

Fuel Metering Valve

The HMU has three hydraulic circuits:

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.

 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 valves control pressure to engine air system actuators as listed previously. EASA Module 14 B2 Propulsion

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 Shutoff valve which is controlled by the Fuel Control Switch on the Engine start / run lever.

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School of Aeronautical Engineering

INTENTIONALLY LEFT BLANK

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School of Aeronautical Engineering

14.2 Engine Indicating Systems Electrical Type 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:

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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 low or high. Page 42 of 71

School of Aeronautical Engineering Low temperature measuring devices utilise the Resistance 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.

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 Thermocouple, called the Seebeck effect after its’ discoverer.

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.

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

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School of Aeronautical Engineering A millivoltmeter (normally calibrated in degrees Celsius) is used to measure the thermo-e.m.f. Immersion Type Thermocouples

SAMPLING HOLES

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.

slower than pure jet

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.

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School of Aeronautical Engineering Rapid Response thermocouples are typically used to 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.

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

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 much smaller in diameter and, due to the mass of the sheath EASA Module 14 B2 Propulsion

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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School of Aeronautical Engineering Thermocouple Location 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.

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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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School of Aeronautical Engineering The parallel-connected thermocouple leads pass through 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.

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School of Aeronautical Engineering If the indicator is connected to the thermocouple system and its’ temperature increases this reduces the temperature difference which tends to make the indicator under-read but the operation of the bi-metal spring off-sets this and causes the indication to increase accordingly. The electrical compensation method will be explained later in these notes with the servo type indicator.

Practical Aircraft E.G.T. Measurement System Thermocouples are fitted in the area where the measurement is to be made They are connected in parallel to give the mean e.m.f. from all the thermocouples, this being an average of the temperature in this area. Extension Leads, made from the same material as the thermocouples, are used to connect the thermocouples in parallel and thence to the main engine junction box. A Ballast Resistor is selected on initial ground test, by the engine manufacturer, to compensate for engine build tolerances. It shunts away a portion of the signal (sometimes called `Temperature Scatter') to give a common temperature output signal at a reference engine speed. Its value is stamped on the engine data plate and recorded in the engine log book. If necessary, it must be replaced by one of the same value and specification.

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School of Aeronautical Engineering The extension and compensating leads also increase their resistance due to natural aging. To maintain the resistance of the circuit external of the indicator a Trimming Resistor is used to adjust this resistance to a pre-set, common value, as specified in the Aircraft Maintenance Manual. (Typically in the order of 15 ± 0.01Ω). This trimmer is usually of the bobbin type, the unwound portion being removed on adjustment. It may be fitted in either lead, but it must be compatible with that lead. (e.g. positive lead - Manganin wire; negative lead - Eureka wire). Compensating Leads, made from materials with the same thermo-electric properties as the thermocouples, are used to connect the signal from the thermocouples to the indicator. They have a very low resistance and form an integral part of the circuit resistance. They must never be repaired or shortened, any excess lead being coiled up and cleated in place: The level of signal from the thermocouples is in the order of only 20 to 40 millivolts, so differences in the circuit resistance would have a great effect on the indicator readings.

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Thermocouple components and leads are colour and size coded. Be warned that U.K. and U.S. colour codings are different (e.g. Chromel wires are coloured red under the U.K. system, white in the U.S.). Terminations are also differently sized and must be torqued to different values.

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School of Aeronautical Engineering Servo Operated Indicating System Many systems these days use servo operated instruments and these employ electrical temperature compensation. The thermocouple is connected to a Signal Processing Module whose main components consist of a Cold Junction reference circuit (C J Ref) and an error signal amplifier. The C J Ref Circuit is a bridge circuit and uses a thermistor to compensate for cold junction temperature errors as well as providing an error signal to the servo amplifier. The bridge circuit is fed with a low stabiliser d.c. voltage (typically 7V) which is also fed to a comparator and feedback potentiometer to effect computation of the thermocouple signal.

The output of the bridge circuit is compared with the d.c. output from the wiper of the positional feedback potentiometer, and since the wiper is geared to the main pointer and digital counter of the indicator, then the difference and the potentiometer d.c. represents an error signal which is fed to the servo amplifier which feeds a d.c. voltage to the armature winding of a d.c. motor to adjust the indication accordingly. Since the motor also drives the potentiometer wiper the motor will continue to drive until a 'null' is reached i.e. the output bridge is balanced. The complete circuit is shown in the diagram at left. It will be noted from the diagram that the output from the first stage of the servo amplifier is also fed to a flag warning circuit. This acts as a servo loop monitor which detects any failure of the servo loop to back off the error signal voltage. If such a failure should occur then the flag circuit de-energises a solenoid controlled warning flag which appears across the digital counter display. The flag will also appear in the event of the 115V a.c. supply to the indicator falling below 100V. An overtemp ’tell-tale’ pointer is carried by the main pointer as the reading increases. When the temperature reduces, the main pointer moves, but the limit pointer is latched at the maximum temp reached. It can only be returned to its normal position by applying a separately switched 28V dc supply to a reset solenoid within the indicator.

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School of Aeronautical Engineering 14.2.2 Engine Speed

cage rotor which gives high torque for easy starting. The pointer is fed through a permanent magnet and drag cup.

Engine speed is an important parameter. It allows accurate control of the engine in gas turbines, this RPM will be measured as a percentage as opposed to direct revolutions per minute (RPM), and in most cases are referred to as 'N' gauges rather than RPM gauges or tachometers. Electrical Tachometers There are two types of tachometers. One is a tachogenerator which consists of an electrical generator, mechanically driven by an engine gearbox, electrically coupled to an indicator. The other is a tacho-probe, whose output is pulsed proportional to rotational speed, and which has several outputs and can, therefore, feed additional system, such as a Flight Data Recorder and an engine control system. Tacho-generators can be DC or AC; but DC types have commutators and brushes whereas the AC types have purely inductive coupled connections. Modern small aircraft therefore use AC tacho-generators. The AC generator gives a 3 phase output which h is fed to the indicator. The frequency of the output is proportional to engine speed, and is used to drive an AC synchronous motor in the indicator. This is combined with a squirrel

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School of Aeronautical Engineering Servo Operated Tachometer Systems More modern aircraft can be fitted with servo operated tachometers. The generator signals are initially fed to the signal processing module where they are converted to square waves by a squaring amplifier. The square waveform is differentiated by the signal shaping circuit to produce positive and negative triggering pulses, which are fed to a monostable via a buffer amplifier. The output of the buffer amplifier produces a train of pulses of constant width and amplitude at twice the frequency of the generator. This output is now fed to an integrator to produce a dc output called the 'Demand Signal' which is fed to the servo amplifier and monitor module. The demand signal is compared with the dc output from the wiper of a positional feedback potentiometer. Since the wiper is geared to the main pointer of the indicator, its output represents indicated speed. Any difference between the indicated speed and the demand speed results in an error signal, which is fed to the input of the servo amplifier whose output is connected to the armature winding of the servo motor. The indicator pointer and digital counter are then driven to the demanded speed position. At the same time, the feedback potentiometer wiper is also re-positioned to provide a feedback voltage to back-off the demand signal until the error is zero; at this point, the indicator will now display the demanded speed.

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The output voltage of the servo amplifier is also fed to a servo loop monitor, which detects any failure of the servo loop to back-up the error voltage. In the event of such failure, the monitor de-energises a solenoid-controlled warning flag, which appears across the digital counter display.

An overspeed ‘Tell-Tale’ pointer operates in a similar manner to that already described in the servo temperature indicator.

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School of Aeronautical Engineering Tacho-Probe & Indicator System

The pole pieces are in close proximity to the teeth of a gear wheel driven at the same speed as the compressor shaft or fan shaft as appropriate. To ensure correct orientation of the probe, a locating lug is provided in the mounting flange.

Commonly called a ‘phonic wheel’ probe; this system is also used in several types of large public transport aircraft, but has the advantage of being able to provide separate electrical outputs to other systems as well as tachometer indicators. In addition there is the advantage that a probe as shown above has no moving parts. The stainless steel hermetically sealed probe consists of a permanent magnet, a pole piece, and a number of cupronickel or nickel / chromium coils wound on a ferromagnetic core. Separate windings provide outputs to the indicator and other equipment requiring engine speed data.

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The diagram above illustrates how the permanent magnet produces a magnetic field around the sensing coils. As the gear teeth pass the pole, the intensity of the flux changes, due to the change in the air gap between the poles and the gear wheel teeth.

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School of Aeronautical Engineering The change in flux density generates an emf in the sensing coils which is fed to the indicator. The probe and gear teeth may therefore be considered as a magnetic flux switch that induces emf's directly proportional to the gear wheel and compressor or fan shaft speed. 14.2.3 Engine Thrust Indication: Engine Pressure Ratio, Engine Turbine Discharge Pressure or Jet Pipe Pressure Systems Thrust Meters The take-off thrust from a turbo-jet engine is normally shown on the Thrust Meter, of which there are two basic types. The first type measures turbine discharge or jet pipe pressure. The most common type is the Percentage Thrust Meter. The second type which is called an Engine Pressure Ratio (EPR) Gauge, measures the ratio of jet pipe pressure to compressor inlet pressure. In both types of indicator an indication of thrust output is given, although when the turbine discharge pressure only is measured, correction is necessary for any variation in inlet pressure. In addition both types may require corrections for variations of ambient air temperature. EASA Module 14 B2 Propulsion

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School of Aeronautical Engineering Power monitoring is done by means of a direct reading type of pressure ratio indication that indicates power in terms of percentage over a range of 50 to 100%. In addition it incorporates a manually controlled device permitting the thrust indications to be compensated for variation in ambient atmosphere conditions. The compensation is accomplished by rotating a setting knob which adjusts a counter of a three digit display each number at which refers to an appropriate atmospheric condition obtained from performance curves. At the same time it rotates the mechanism and positions the 'bug' to a new datum value on the instrument face. Engine Pressure Ratio (EPR) This method of indication used intake pressure and jet pipe manifold pressure, the pressure ratio of which varies directly with thrust (see diagram following). In general, an EPR system consists of an engine inlet pressure sensing probe, a number of pressure sensing probes projected into the exhaust unit of an engine, a pressure ratio transmitter and an indicator. The interconnection of these components, based on a system in current use, is schematically shown in the diagram on the following page.

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The inlet pressure-sensing probe is similar to a pitot pressure probe, and is mounted so that it faces into the airstream in the engine intake or, as in some power-plant installations, on the pylon and in the vicinity of the air intake. The probe is protected against icing by a supply of warm air from the engine anti-ice system. The exhaust pressure-sensing probes are interconnected by pipelines which terminate at a manifold, thus averaging the pressures. A pipeline from the manifold and another from the inlet pressures probe, are each connected to the pressure ratio transmitter which comprises a bellows type of pressure-sensing transducer, a linear voltage differential transformer (LVDT) a two-phase servomotor, amplifier and a potentiometer. The transducer bellows are arranged in two pairs at right angles and supported in a frame which, in turn, is supported in a gimbal and yoke assembly. The gimbal is mechanically coupled to the servomotor via a gear train, while the yoke is coupled to the core of the LVDT.

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Engine Pressure Ratio system schematic Diagram

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School of Aeronautical Engineering The servomotor also drives the wiper of the potentiometer which adjusts the output voltage signals to the EPR indicator in terms of changes in pressure ratio, alters the potentiometer output signal to the indicator the pointer and digital counter of which are servo-driven to indicate the new pressure ratio. Simultaneously, the motor drives the transducer gimbal and LVDT coils in the same direction as the initial yoke movement so that the relative movement now produced between the LVDT coils and core starts reducing the signal to the servomotor, until it is finally cancelled and the system stabilised at the new pressure ratio.

Engine Oil Temperature Measurement The construction of a typical sensing element commonly used for sensing liquid temperatures is shown below.

14.2.4 Oil Pressure & Temperature Oil within an engine plays the vital role of lubricating bearings, some of which are highly stressed and are required to operate at high temperatures. As the lubricating efficiency of oil deteriorates at high temperatures, it is essential that the oil is cooled and that the pilot receives an indication that the system is operating satisfactorily. The pilot also needs to know that the pressure at which the oil is being delivered to the bearings is adequate.

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In practice a Wheatstone bridge is used to measure the change in resistance in oil and fuel temperature measurement systems as shown opposite. Resistors A, B and C are known values; that of the heat sensitive bulb D changes with temperature. This causes imbalance in the bridge and current flow from Y to X. The meter is calibrated in degrees C.

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Engine Oil Pressure Measurement Oil pressure is sensed by an electrical transducer/transmitter also located upstream of the bearings. A change in oil pressure causes a change in current from the transmitter and hence the indicator receives a change in current which is proportional to the change in oil pressure. Oil pressure is sensed by an electrical transducer/transmitter also located upstream of the bearings.

Oil temperature is taken by a temperature sensitive element fitted in the oil system upstream of the bearings. Changes in temperature of the oil cause changes in electrical resistance and hence alterations in the current to the indicator. See previous notes on Temperature Measurement.

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A change in oil pressure causes a change in current from the transmitter and hence the indicator receives a change in current which is proportional to the change in oil pressure.

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School of Aeronautical Engineering 14.2.5 Fuel Pressure, Temperature and Flow

Fuel Flow Indicating Systems

Fuel Pressure and Temperature

As stated fuel flow indicating systems have two main units. These are the fuel flow sender or transmitter and the indicator itself. The systems are included in aircraft to measure the rate of fuel flow to the engines. In order to operate successfully the following criteria must be met:

Fuel Temperature and Pressure of the low pressure fuel supply are electrically transmitted to their respective indicators. They are similar in operation to those for oil temperature and pressure, so no further description will be given in these notes.

1. They must be able to indicate the rate of fuel flow accurately.

Fuel Flow 2. The transmitter must not impede the flow of fuel. Although the amount of fuel consumed during a given flight may vary slightly between engines of the same type, fuel flow does provide a useful indication of the satisfactory operation of the engine and of the amount of fuel being consumed during the flight. A typical system consists of a fuel flow transmitter, which is fitted into the low pressure fuel system, and an indicator, which shows the rate of fuel flow and the total fuel used in gallons, pounds or kilogrammes per hour. The transmitter measures the fuel flow electrically and an associated electronic unit gives a signal to the indicator proportional to the fuel flow (see diagram following).

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3. If a mechanical breakdown occurs, then the maximum rate of fuel flow to the engine should be provided. 4. They must include compensation for changes in fuel temperature.

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School of Aeronautical Engineering Types of Fuel Flow Systems The following notes describe three types of fuel flow meters. Rotating Vane Fuel Flow System Transmitter The transmitter has 4 sections:

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1.

The rotating vane, whose shaft determines the electrical output.

2.

The damping section, with a fuel filled compartment containing a damping vane to remove oscillations of the moving vane, the damping vane also acts as a counter-balance to the moving vane, this section also houses the calibration spring.

3.

Information transmission section, - a ring magnet is attached to the moving vane shaft and transmits shaft movement to a bar magnet in the electrical section. The use of a ring and bar magnet eliminates the risk of fuel in the electrical section.

4.

The electrical section houses a bar magnet, whose movement varies the output from a potentiometer and therefore the output to the indicator.

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School of Aeronautical Engineering Indicator The indicator is of the integrated type that indicates the fuel flow rate and the total fuel consumed. It contains an inverter circuit, which provides various signals to the circuit: 1. 2. 3. 4. 5.

Motor reference phase. Tacho - generator feed-back reference phase. Amplifier input. Fuel flow transmitter potentiometer supply. An anti-phase signal for the resetting of the total fuel consumed counters.

The indicator consists mainly of a low inertia two-phase induction motor, which provides 2 integrated outputs (the instantaneous flow rate and the total fuel consumed). The motor also drives a feedback tacho-generator, which provides damping through a negative feedback to the amplifier proportional to the motor speed. The fuel flow pointer is driven via a magnetic drag cup assembly. The total fuel consumed counters are operated by a mechanical drive via a gearbox.

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Fuel entering the metering chamber is straightened before it impinges on the vane, which rotates against the tension of the calibration spring. The chamber is non-linear in shape (involute) to produce a linear vane shaft movement, which is conveyed to a potentiometer via ring and bar magnets. The potentiometer output is fed to the amplifier to drive an induction motor coupled to a gearbox producing 3 outputs: 1. A drive to a drag/disc assembly to operate the flow rate pointer. 2. A mechanical drive to operate the total fuel consumed counters. 3. A drive to a tacho-generator, which produces a negative feedback signal proportional to the rate of fuel flow.

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School of Aeronautical Engineering Impeller Fuel Flow System This type of flow meter is designed to continually display the rate of fuel flow, and the total fuel quantity consumed, in terms of mass units. The function of the flow meter transmitter depends upon the volumetric rate of fuel flow and therefore if the system is measuring fuel mass, a correction will have to be made for the density and temperature of the fuel.

The output of the transmitter is fed to an integrator within the indicator, to be amplified and shaped for the operation of the fuel flow rate and total fuel consumed indicators.

The transmitter consists of a light alloy casting with guide vanes and an electrical ‘pick-off’ coil. Inside the casting there is a helical vane rotor which has a magnet embedded in it. When the impeller rotates (due to the fuel flow) a sinusoidal signal, at a frequency proportional to the speed of the rotor and hence the rate of fuel flow, will be induced in the pick-off coil. Temperature Correction For a decrease in temperature, the fuel becomes denser, giving a lower flow rate. This results in a decrease in the signal frequency and therefore the indicator would underread. A temperature sensor output is fed to the indicator and applied to the computed outputs from the transmitter to give a temperature corrected indication.

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School of Aeronautical Engineering As alternatives to the basic counter and milliammeter the signals can be fully digitised and then fed to electronic digital display indicators. Further modifications to the signal are carried out by temperature and density compensators. Impeller/Turbine Fuel Flow Indicating System This type of fuel flow transmitter has 4 sections: 1.

A static frequency controller, which will maintain its output to close tolerances - i.e. ± 0.3% frequency.

2.

An impeller, which is driven through reduction gearing by a motor using the frequency controller output - this keeps the impeller at a constant speed.

3.

A turbine that contains fuel-straightening vanes. The turbine is mechanically independent of the impeller, but is restrained by a spring.

4.

An electrical transmitter, whose output is controlled by the position of the turbine.

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The picture at on the following page shows an integrating flow meter system as used in the Boeing 737 aircraft. The constant speed impeller imparts an angular momentum to the fuel, proportional to the rate of fuel flow. This angular momentum of the fuel is applied to the straightening vanes in the turbine, causing it to rotate until the calibrated restraining springs balance the force due to the momentum of the fuel. The deflection of the turbine shaft positions the LVDT to a position corresponding to the fuel flow in the line. A signal voltage (up to 5V at maximum fuel flow) is induced in the secondary of the LVDT and is supplied to the indicator servomotor via the closed contact of the reset switch and amplifier. The servomotor rotates at a rate proportional to the flow rate, driving the Flow Rate Pointer via a magnetic drag cup and the Fuel Used counters via a mechanical gearbox. The Reset Switch is located on a panel in the cockpit and, when pressed, energises the reset relay, whose contacts supply 115 V a.c. to the servo amplifier and motor, causing it to drive the fuel used counters rapidly to zero. The decoupling disc prevents any hydraulic coupling between the impeller and turbine at low flow rates.

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School of Aeronautical Engineering combustion chamber. As you increase the throttle setting, more fuel and air is flowing to the engine; therefore, MAP increases. When the engine is not running, the manifold pressure gauge indicates ambient air pressure (i.e., 29.92 in. Hg). When the engine is started, the manifold pressure indication will decrease to a value less than ambient pressure (i.e., idle at 12 in. Hg). Correspondingly, engine failure or power loss is indicated on the manifold gauge as an increase in manifold pressure to a value corresponding to the ambient air pressure at the altitude where the failure occurred.

14.2.6 Manifold Pressure On airplanes that are equipped with a constant-speed propeller and piston engine, power output is controlled by the throttle and indicated by a manifold pressure gauge. The gauge measures the absolute pressure of the fuel/air mixture inside the intake manifold and is more correctly a measure of manifold absolute pressure (MAP). At a constant r.p.m. and altitude, the amount of power produced is directly related to the fuel/air flow being delivered to the

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Engine power output is indicated on the manifold pressure gauge.

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School of Aeronautical Engineering The manifold pressure gauge is color-coded to indicate the engine’s operating range. The face of the manifold pressure gauge contains a green arc to show the normal operating range, and a red radial line to indicate the upper limit of manifold pressure. For any given r.p.m., there is a manifold pressure that should not be exceeded. If manifold pressure is excessive for a given r.p.m., the pressure within the cylinders could be exceeded, thus placing undue stress on the cylinders. If repeated too frequently, this stress could weaken the cylinder components, and eventually cause engine failure. 14.2.7 Engine Torque In large supercharged piston engines most of the propulsive force is produced at the propeller. This is also true for turbo-prop engines, as only a small part of the force is derived from the jet thrust. Measurement of this propulsive force is carried out using a Torque meter Indicating System measuring the torque developed at the propeller. (Diagram on following page). Pulse Probe Torque meter System The torque meter assembly provides the means of transmitting and measuring torque produces by the power section.

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A correct indication of torque is essential due to the fact that torque will vary over a considerable range, as it is affected greatly by ambient air conditions at the engine inlet, and by turbine inlet temperature. Since this is a constant speed engine from take-off to landing, engine rpm is not a factor to be considered in the variation of torque. The torque meter shaft assembly is the rotating portion of the torque meter system. It includes two concentric shafts, two sleeve bearings, and the engine to torque meter coupling. The torque meter inner shaft (torque shaft) is a solid steel shaft which carries the torsional load. It is bolted to the safety coupling at the forward end and is splined to the compressor extension shaft at the aft end. The torque meter outer shaft (reference shaft) is connected to the inner shaft by locating key at the aft end. Concentricity between the two shafts is maintained by the centre and front sleeve bearings. Both shafts rotate as a unit. The forward end of each shaft is flanged and has equally spaced rectangular exciter teeth machined on these flanges. As the amount of torque transmitted increases, the torsional deflection of the inner shaft causes a displacement between the exciter teeth of the two flanges. The displacement is detected by electromagnetic pickups mounted in the torque meter housing.

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The torque meter magnetic pickup assembly contains two identical pickups. One pickup is located over the torque meter inner shaft flange, and the other is located over the torque meter outer shaft flange. When a load is applied, the torque shaft twists, displacing the teeth on the reference flange. This displacement causes a change in the phase displacement between impulses produces at the magnetic pickup mounted over the flange. The air gap between the pickup assembly and the exciter teeth affects pulse voltage.

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School of Aeronautical Engineering 14.2.8 Propeller Speed The following is a description of a typical Propeller Speed System: NP (Propeller RPM) A separate speed indicating system is provided for each propeller. Each system consists of a propeller speed (NP ) sensor connected electrically to the rpm indicator in the cockpit. The propeller magnetic pickup speed sensor is supplied as part of the basic engine. It is mounted on the reduction gearbox and provides a signal (0 to 5 volts peak to peak), equivalent to 0 to 1,200 propeller rpm. The signal is sent to the cockpit indicator. The No. 1 and No.2 N P indicators are powered by 28V DC. Lighting for the instruments is powered from the 5V DC lighting system. Circuits in each indicator compute the AC signal from its associated sensor and provide an equivalent indication of NP . A 0 to 5 VDC signal, proportional to propeller rpm, is relayed to the flight data recorder.

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INTENTIONALLY BLANK

A press-to-test button on the indicator, when pressed with power on the indicator, causes the pointer to align with a blue dot at 1,050 rpm on the dial and 1,050 to be displayed in the digital display. In the event of indicator failure, the needle moves off scale below zero, and the digital display is blanked.

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School of Aeronautical Engineering 14.3 Starting And Ignition Systems INTRODUCTION Two main systems are required to ensure that a gas turbine engine will start satisfactorily. Firstly, provision must be made for the compressor and turbine to be rotated up to a speed at which adequate air passes into the combustion system to mix with the fuel from the fuel spray nozzles. Secondly, provision must be made for ignition of the air/fuel mixture in the combustion system. 14.3.1 Operation Of Engine Starting Systems And Components During normal engine starting, the two systems must operate simultaneously. It must also be possible to motor the engine over without ignition for maintenance checks and to blow out residual fuel after a failed start. In addition, it must be possible to operate the ignition system for relighting the engine during flight.

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

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

Methods of Starting The starting procedure for all jet engines is basically the same, but can be achieved by various methods.

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School of Aeronautical Engineering The type and power source for the starter varies with engine and aircraft requirements. The power sources can be electrical, gas, air or hydraulic and each method has its merits. The requirements for a military aircraft, for example, are totally different to those for a commercial airliner. The starter motor must, however, always produce a high torque and then transmit this torque to the engine in a smooth manner to accelerate it to self- sustaining speed.

The electrical supply voltage can be progressively increased by the removal of resistances in the circuit as the engine increases in speed. The ignition system is also actuated and supplied at the same time as the start is initiated. Once the engine is running, the starter supply is cancelled by the drop in supply current or by the action of a timer mechanism. Either way, the starter slows down and the clutch or ratchet mechanism ensures that the engine can accelerate free from the starter drive shaft.

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

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The diagram in figure 13.1 shows a simplified electric starter circuit. It contains most of the components found in many starter circuits such as master switch, start button and main starter relay. Overspeed relays usually disconnect the starter motor electrically, once the amount of current being drawn falls below a value which can only be reached if the engine is self - sustaining.

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Figure 13.1. Electric Starting Circuit

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School of Aeronautical Engineering AIR STARTING Air starting is used on most commercial and some military jet engines. It has many advantages over other starting systems and is comparatively light, simple and economical to operate. An air starter motor transmits power through a reduction gear and clutch to the starter output shaft, which is connected to the engine. A typical air starter is a basic air turbine that rotates at high RPM when HP air is passed through it from the on board Auxiliary Power Unit, (APU), a cross-feed from a running engine, or an external air supply. The air supply, from whichever source, is controlled by an electrically operated control and pressure-reducing valve that is opened when an engine start is selected. It is automatically closed at a pre-determined starter speed. The clutch automatically disengages as the engine accelerates up to idling RPM and the rotation of the starter ceases.

Fig. 13.2. Air Start System

Figure 13.2 shows a typical air starting system and Figure 13.3 shows a cut-away of the actual starter motor showing its rotor.

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Fig 13.3 Air Starter System and Cut Away Starter

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

Fig 13.4 Impingement Air Starter System

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

Fig 13.5 Self – contained Gas Turbine Stater

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School of Aeronautical Engineering HYDRAULIC STARTING

14.3.2 Ignition Systems And Components

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

High Energy Ignition Systems

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

Each HE ignition unit receives a low voltage supply, controlled by the starting system circuit, from the aircraft’s electrical system. The electrical energy is stored in the unit until, at a pre-determined value, the energy is dissipated as a high voltage, high current discharge across the plug. A choke is fitted to extend the duration of the discharge and safety resistors are fitted to ensure dissipation of energy in the capacitors.

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

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

A typical, simple ignition system is illustrated in Figure 13.6 and shows how the inputs are modified, through several stages, to give a high voltage, direct current to the HT terminal of the Igniter.

To be able to operate at both levels, combined systems, giving high and low level outputs are most popular. Such a system would consist of one unit emitting a high output to one Igniter plug and a second unit giving a low output to a second Igniter plug. Some Igniter units have been manufactured which contain both high and low outputs, which means that two igniters can be operating at either level depending on the conditions and the relevant cockpit selection. Fig 13.6 Typical Ignition System

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

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

The Igniter plugs operate in the same way as sparking plugs, except that they are only required to start the engines, they are then switched off until the next start. There are two basic types of Igniter plug, the air gap type and the surface discharge type. The air gap type require a potential difference in the region of 20,000 volts, whist the surface discharge type only requires a voltage in the region of 2,000 volts. As igniters are used for both low-tension D.C. systems and high-tension A.C. systems and are NOT interchangeable, care must be taken to use the correct item, as recommended by the manufacturer, in their overhaul/maintenance manuals.

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School of Aeronautical Engineering The two illustrations one unit of a dual ignition system (left) and a surface discharge igniter (right).

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School of Aeronautical Engineering RELIGHTING The jet engine requires the facility for relighting should the flame in the combustion chamber become extinguished during flight. This ‘relighting’ can only be safely accomplished if the aircraft is at the correct speed and below a certain altitude. Figure 13.9 illustrates the relighting envelope for a specific aircraft. If the aircraft is too slow or too fast, or if it is above about 25,000 feet, there is little chance for the engine to relight. Within this envelope the airflow will rotate the compressor at a speed satisfactory for relighting.

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School of Aeronautical Engineering 14.3.3 Maintenance Safety Requirements It has already been mentioned that the HEIUs can deliver a fatal shock if they are handled whilst still live. It is also possible to be shocked if either the Igniter leads or the igniters themselves are handled before 1 minute has elapsed after removing all power from the system. DO NOT depend on just the starter master switch being placed into the ‘OFF’ position as it is possible someone may switch it to ‘ON’ whilst you are working some way from the cockpit, on aft mounted engines for example. At least pull AND LABEL AS ‘INOP’, any circuit breakers applicable to the HEIUs. Also, disconnect the Low-Tension connectors on the Igniter box itself to be doubly sure.

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

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

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School of Aeronautical Engineering Acronyms and Abbreviations ADC Air Data Computer ARINC Air Radio Inc. BCD Binary Coded Decimal BITE Built In Test equipment CDU Control Display Unit CU Control Unit ECAM Electronic Centralised Aircraft Monitoring ECU Engine Control Unit EEC Electronic Engine Control EGT Exhaust Gas Temperature EHSV Elector-Hydraulic Servo Valve EICAS Engine Indicating Crew Alerting System EIMU Engine Interface Monitoring Unit EPR Engine Pressure Ratio FADEC Full Authority Digital Engine Control FMS Flight Management System FMV Fuel Metering Valve HMU Hydro-Mechanical Unit LCD Liquid Crystal Display LED Light Emitting Diode LRU Line Replaceable Unit MSU Mode Select Unit PIMU Propulsion Interface Management Unit TLA Throttle Lever Angle TMC Thrust Management Computer TRA Throttle Resolver Angle VBV Variable Bleed Vanes VSV Variable Stator Vanes EASA Module 14 B2 Propulsion

Bibliography and Recommended Further Reading FADEC for EASA Pt 66 by Total Training Support Rolls Royce The JET ENGINE ISBN 0 902 121 2 35 COBC Instrument Notes (Module 13)

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END OF MODULE 14

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