Propulsion 14.1 4f812

Propulsion

EASA Part-66 Cat-B2 Module-14 Shahzad Khalil (Sr.Engg. Instructor)

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Contents 1 Basic mechanics and Engine operation 2 Engine construction 3 Fuel Control and FADEC

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Chapter-1: Basic mechanics and Engine Operation



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Classification of Engine Principles Of Jet Propulsion Theory Of Jet Propulsion Theory Of Gas Turbine Engines Working Cycle of Gas Turbine Engine Changes In Velocity And Pressure Factors Affecting Thrust Types of Gas Turbine Engine Performance Engine Station Designations Advantages/Disadvantages of Turbine Engine

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Engine (Heat) A device/machine which converts heat energy into mechanical energy.

1.1-Classification of Engine: Engine

Internal Combustion Engine (Combustion inside the Engine)

External Combustion Engine (Combustion outside the Engine ) Steam Engine

Air Dependent

Air Independent Rocket

Piston (Intermittent flow) Engine

Reaction (Continuous flow) engine

Ram Jet Pulse Jet

Gas Turbine Engine

Thrust Producing TurboJet TurboFan

Torque Producing TurboProp Turboshaft

1.2-Principles of jet propulsion The jet engine relies on the principle of taking in a mass of air and accelerating it rearwards. This means that according to Newton laws of motion a forward reaction will be produced. The important laws for jet propulsion are: Newton's 2nd Law of Motion. The second law states that an imbalance of forces on a body produces or tends to produce an acceleration in the direction of the greater force, and the acceleration is directly proportional to the force and inversely proportional to the mass of the body. F= m a Newton's 3rd Law of Motion. The third law states that for every action there is an equal and opposite reaction and the two is directed along the same straight line. Training Purposes Only

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Einstein's Law of Conservation of Energy. This law states that the amount of energy in the universe remains constant. It is not possible to create or destroy energy; however, it may be transformed. Boyle's Law: Boyle's Law states the volume of a definite quantity of dry gas is inversely proportional to the pressure, provided the temperature remains constant. Mathematically Boyle's law can be expressed as P1V1 = P2V2 Charles' Law: Charles's Law can be stated as the volume occupied by any sample of gas at a constant pressure is directly proportional to the absolute temperature. V / T =constant •

V is the volume

•

T is the absolute temperature (measured in Kelvin)

Charles's Law can be rearranged into two other useful equations. V1 / T1 = V2 / T2 Bernoulli's Principle: It states that in an incompressible, non-viscous fluid the sum of pressure and kinetic energy is always constant. This means that if the velocity of a gas or liquid is increased its pressure will decrease. The opposite is also true. If the velocity of a gas or liquid is decreased its pressure will increase. This fact relates directly to the law of conservation of energy.

Fig.1 Pressure and Velocity: Air is normally thought of in relation to its temperature, pressure, and volume. Within a gas turbine engine the air is put into motion so now another factor must be considered, velocity. Consider a constant airflow through a duct. As long as the duct crosssectional area remains unchanged, air will continue to flow at the same rate (disregard frictional loss). If the cross-sectional area of the duct should become smaller (convergent area), the airflow must increase velocity if it is to continue to flow the same number of pounds per second of airflow (Bernoulli's Principle). In order to obtain the necessary velocity energy to accomplish this, the air must give up some pressure and temperature energy (law of conservation of energy). The net result of flow through this restriction would be a decrease in pressure and temperature and an increase in velocity. The opposite would be true if air were to flow from a smaller into a larger duct (divergent area); velocity would then decrease, and pressure and temperature would increase. The throat of an automobile carburetor is a good example of the effect of airflow through a restriction (venturi); even on the hottest day the Training Purposes Only

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center portion of the carburetor feels cool. Convergent and divergent areas are used throughout a gas turbine engine to control pressure and velocity of the air-gas stream as it flows through the engine.

Fig. 2

1.3-Theory of Jet Propulsion The principle of jet propulsion can be illustrated by a toy balloon. When inflated and the stem is sealed, the pressure is exerted equally on all internal surfaces. Since the force of this internal pressure is balanced there will be no tendency for the balloon to move.

Fig. 3 If the stem is released the balloon will move in a direction away from the escaping jet of air. Although the flight of the balloon may appear erratic, it is at all times moving in a direction away from the open stem.

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Fig. 4 The balloon moves because of an unbalanced condition existing within it. The jet of air does not have to push against the outside atmosphere; it would function better in a vacuum. When the stem area of the balloon is released, a convergent nozzle is created. As the air flows through this area, velocity is increased accompanied by a decrease in air pressure. The flight of the balloon will be of short duration, though, because the working fluid-air in the balloon is soon gone. If a source of pressurized air were provided, it would be possible to sustain flight of the balloon.

Fig.6 A garden sprinkler rotated by the reaction of the water jets

Fig.5 Hero’s engine -the earliest form of jet reaction

The familiar whirling garden sprinkler (fig. 6) is a more practical example of the principle, for the mechanism rotates by virtue of the reaction to the water jets.

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1.4-Theory of Gas Turbine Engines If the balloon were converted into a length of pipe, and at the forward end an air compressor designed with blades somewhat like a fan were installed, this could provide a means to replenish the air supply within the balloon.

Fig. 7 A source of power is now required to turn the compressor. To extend the volume of air, fuel and ignition are introduced and combustion takes place. This greatly expands the volume of gas available.

Fig. 8 In the path of the now rapidly expanding gases, another fan or turbine can be placed. As the gases through the blades of the turbine, they cause it to rotate at high speed. By connecting the turbine to the compressor, we have a mechanical means to rotate the compressor to replenish the air supply. The gases still possessing energy are discharged to the atmosphere through a nozzle that accelerates the gas stream. The reaction is thrust or movement of the tube away from the escaping gas stream. We now have a simple turbojet engine.

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Fig. 9 The turbojet engine is a high-speed, high-altitude Powerplant. The simple turbojet engine has primarily one rotating unit, the compressor/turbine assembly. The turbine extracts from the gas stream the energy necessary to rotate the compressor. This furnishes the pressurized air to maintain the engine cycle. Burning the fuel-air mixture provides the stream of hot expanding gas from which approximately 60 percent of the energy is extracted to maintain the engine cycle. Of the total energy development, approximately 40 percent is available to develop useful thrust directly. The amount of energy required to rotate the compressor may at first seem too large; however, it should be ed that the compressor is accelerating a heavy mass (weight) of air towards the rear of the engine. In order to produce the gas stream, it was necessary to deliver compressed air by a mechanical means to a burner zone. The compressor, being the first rotating unit, is referred to as the N1 system. With a requirement for an engine that delivers rotational shaft power, the next step is to harness the remaining gas stream energy with another turbine (free turbine). By connecting the turbine to a shaft, rotational power can be delivered to drive an aircraft propeller, a helicopter rotor system, a generator, a tank, an air cushion vehicle (ACV), or whatever is needed. The power shaft can extend from the front, back, or from an external gearbox. All of these locations are in use on various types of engines at present. The following sketch shows a turboshaft engine with the power shaft extended out the front. The bottom sketch shows the same engine with the power shaft extending out the back. The basic portion of the turbine engine, the Gas producer (Generator), extracts approximately 60 percent of the gas stream energy (temperature/pressure) to sustain the engine cycle. To develop rotational shaft power, the remaining gas stream energy must drive another turbine. In engines today, a power turbine that is free and independent of the Gas Generator system accomplishes this task. The power turbine and shaft (N2 system) are not mechanically connected to the gas producer (N1 system). It is a free turbine. The gas stream ing across the turbines is the only link between these two systems. The free-turbine engine can operate over wide power ranges with a constant output-shaft speed.

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Fig. 10

With a turbojet engine, power (thrust) produced is roughly the difference between the velocity of the air entering the engine and the velocity of the air exiting from the engine. Efficiency of the engine (power producer versus fuel consumed) increases with speed until it is 100 percent efficient when the forward speed of the engine is equal to the rearward speed of the jet. Aircraft reciprocating engines operate on the four-stroke, five-event principle. Four strokes of the piston, two up and two down, are required to provide one power impulse to the crankshaft. Five events take place during these four strokes: the intake, compression, ignition, power, and exhaust events. These events must take place in the cylinder in the sequence given for the engine to operate.

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Fig.11 A comparison between the working cycle of a turbo-jet engine and a piston engine

Although the gas turbine engine differs radically in construction from the conventional fourstroke, five-event cycle reciprocating engine, both involve the same basic principle of operation. In the piston (reciprocating) engine, the functions of intake, compression, ignition, combustion, and exhaust all take place in the same cylinder and, therefore, each must completely occupy the chamber during its respective part of the combustion cycle. In the gas turbine engine, a separate section is devoted to each function, and all functions are performed at the same time without interruption.

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1.5- Working Cycle Otto Cycles (Reciprocating Engine): There is an element of similarity to both the reciprocating and jet engines, but the thermodynamic cycle of each is different from the other. The reciprocating engine operates on the Otto cycle, a constant volume cycle, consisting of four distinct operations. These operations are performed intermittently by a piston reciprocating in an enclosed cylinder. It is important to that the piston in a reciprocating engine delivers power only during one of its four strokes. Brayton Cycle Of Operation (Gas Turbine Engine) The turbine engine operates on the Brayton cycle, a constant pressure cycle containing the same four basic operations as the Otto cycle, but accomplishing them simultaneously and continuously so that an uninterrupted flow of power from the engine results.

Pressure

Fig. 12 The working cycle on a Pressure/Volume diagram The working cycle upon which the gas turbine engine functions is, in its simplest form, represented by the cycle shown on the pressure volume diagram in fig. Point A represents air at atmospheric pressure that is compressed along the line AB. From B to C heat is added to the air by introducing and burning fuel at constant pressure, thereby considerably increasing the volume of air. Pressure losses in the combustion chambers are indicated by the drop between B and C. From C to D the gases resulting from combustion expand through the turbine and jet pipe back to atmosphere. During this part of the cycle, some of the energy in the expanding gases is turned into mechanical power by the turbine; the remainder, on its discharge to atmosphere, provides a propulsive jet. Because the turbo-jet engine is a heat engine, the higher the temperature of combustion the greater is the expansion of the gases. The combustion temperature, however, must not exceed a value that gives a turbine gas entry temperature suitable for the design and materials of the turbine assembly. The use of aircooled blades in the turbine assembly permits a higher gas temperature and a consequently higher thermal efficiency. Training Purposes Only

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1.6-Changes in Velocity and Pressure During the age of the air through the engine, aerodynamic and energy requirements demand changes in its velocity and pressure. For instance: during compression, a rise in the pressure of the air is required and not an increase in its velocity. After the air has been heated and its internal energy increased by combustion, an increase in the velocity of the gases is necessary to force the turbine to rotate. At the propelling nozzle a high exit velocity is required, for it is the change in the momentum of the air that provides the thrust on the aircraft. Local decelerations of airflow are also required, as for instance, in the combustion chambers to provide a low velocity zone for the flame to burn. These various changes are effected by means of the size and shape of the ducts through which the air es on its way through the engine. Where a conversion from velocity (kinetic) energy to pressure is required, the ages are divergent in shape. Conversely, where it is required to convert the energy stored in the combustion gases to velocity energy, a convergent age or nozzle (Fig.13) is used. These shapes apply to the gas turbine engine where the airflow velocity is subsonic or sonic, i.e. at the local speed of sound. Where supersonic speeds are encountered, such as in the propelling nozzle of the rocket, athodyd and some jet engines, a convergent-divergent nozzle or venturi (Fig. 13) is used to obtain the maximum conversion of the energy in the combustion gases to kinetic energy. The design of the ages and nozzles is of great importance, for upon their good design will depend the efficiency with which the energy changes are affected. Any interference with the smooth airflow creates a loss in efficiency and could result in component failure due to vibration caused by eddies or turbulence of the airflow.

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Fig. 13 An air flow through divergent and convergent ducts.

Fig. 14 Supersonic airflow through a convergent-divergent nozzle or venturi.

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Free Turbine: The turbine engines are normally of the free-power turbine design, as shown in fig.15. In this engine, nearly two-thirds of the energy produced by combustion is extracted by the gas producer turbine to drive the compressor rotor. The power turbine extracts the remaining energy and converts it to shaft horsepower (shp), which is used to drive the output shaft of the engine. The gas then exits the engine through the exhaust section to the atmosphere.

Fig. 15 Typical Free-Power Turboshaft Engine.

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Fig. 16 Airflow Systems

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Fig.17 Airflow Systems 1.7-FACTORS AFFECTING THRUST The Jet engine is much more sensitive to operating variables. These are: Temperature of the air Pressure of air. Amount of humidity. Engine rpm Speed of aircraft (ram pressure rise). Jet Velocity Weight of fuel flow. Amount of air bled from the compressor. Turbine inlet temperature.

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1.8-Types of Gas Turbine Engine Turbojet: The turbojet is the engine in most common use today in high-speed, high-altitude aircraft. With this engine, air is drawn in by a compressor which raises internal pressures many times over atmospheric pressure. The compressed air then es into a combustion chamber where it is mixed with fuel to be ignited and burned. Burning the fuel-air mixture expands the gas, which is accelerated out the rear as a high-velocity jet-stream. In the turbine section of the engine, the hot expanded gas rotates a turbine wheel which furnishes power to keep the compressor going. The gas turbine engine operates on the principle of intake, compression, power, and exhaust, but unlike the reciprocating engine, these events are continuous. Approximately two-thirds of the total energy developed within the combustion chamber is absorbed by the turbine wheel to sustain operation of the compressor. The remaining energy is discharged from the rear of the engine as a high velocity jet, the reaction to which is thrust or forward movement of the engine. The turbojet is shown schematically in fig.19.

Fig. 18 Axial-Flow Turbojet Engine. TURBOPROP ENGINE AND TURBOSHAFT ENGINE The turboprop engine and turboshaft engines, shown in figures 19 and 20, are of the same basic type as the turbojet. Instead of ejecting high-velocity exhaust gases to obtain thrust, as in the turbojet, a turbine rotor converts the energy of the expanding gases to rotational shaft power. A propeller or helicopter transmission can be connected to the engine through reduction gearing. This energy may be extracted by the same turbine rotor that drives the compressor, or it may be a free-power turbine which is independent of the compressor turbine and only linked to it by the expanding gases.

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Fig. 19 Axial-Flow Turboprop Engine.

Fig. 20 Twin Spool Turbo shaft (with free power turbine) The free-power turbine is the type used in aircraft to harness the energy of the gases and convert this energy to rotational shaft power. This feature of having a free-power turbine enables the power output shaft to turn at a constant speed while the power producing capability of the engine can be varied to accommodate the increased loads applied to the power output shaft.

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1.9-Performance ENGINE EFFICIENCIES The overall efficiency of an engine is the product of Thermal efficiency (Internal)

and

Propulsive Efficiency (External)

Thermal Efficiency: Thermal efficiency is a comparison between the heat released by the fuel and the kinetic energy ed into the gas stream as a result of the fuel burning. It is affected by the temperature drop across the turbine. The higher the turbine entry temperature the greater is the amount of thermal energy available to do work. It is a function of: Engine Pressure Ratio and mass flow Temperature at which air is heated Engine Pressure Ratio: The pressure found in the exhaust of a turbojet engine is an indication of the work done on the airflow through, and by, the engine. By measuring the pressure in the exhaust and comparing it to the pressure of the air found in the intake a ratio can be determined that will indicate how much work the engine is doing on the air:

EPR (Engine Pressure Ratio) =

Exhaust Pressure Intake Pressure

In the case of the high by- turbofan, there may be a requirement to measure the exhaust pressure from the cold stream (fan duct) as well as the hot stream exhaust. Because both the hot stream and the cold stream are producing thrust – although the hot stream does not produce much in comparison with the cold stream, it is useful to combine both for purposes of measurement. The pressures are added together and averaged before being compared with the intake pressure giving an ‘Integrated Engine Pressure Ratio’ (IEPR) that will indicate to the pilot and engineer how much work the engine is doing on the air in total. Propulsive Efficiency: it is related to an engine installed on an airframe. It can be defined as that proportion of engine work that can be converted into aircraft work, it is expressed as: Propulsive Efficiency=work done on Aircraft/work done on gas stream = twice aircraft speed/aircraft speed + engine speed

*100 * 100

= 2Vi / Vi+Vj *100 It indicates how efficient an engine is as a propelling unit. If aircraft is stationary, regardless of the amount of thrust produced, the fuel consumed is wasted as for as the aircraft propulsion is concerned, In-fact the propulsive efficiency is zero. However if the aircraft speed is equal to the jet speed the propulsive efficiency would be 100%. Training Purposes Only

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Propulsive efficiency increases as the difference between the aircraft and jet speed decreases. In other words the faster the aircraft flies, the closer the aircraft and jet velocities and the energy that is put into gas stream performs more useful work. At aircraft speeds below approximately 450 miles per hour, the pure jet engine is less efficient than a propeller-type engine, since its propulsive efficiency depends largely on its forward speed; the pure turbo-jet engine is, therefore, most suitable for high forward speeds. The propeller efficiency does, however, decrease rapidly above 350 miles per hour due to the disturbance of the airflow caused by the high blade-tip speeds of the propeller. These characteristics have led to some departure from the use of pure turbo-jet propulsion where aircraft operate at medium speeds by the introduction of a combination of propeller and gas turbine engine. The advantages of the propeller/turbine combination have to some extent been offset by the introduction of the by-, ducted fan engines. These engines deal with larger comparative airflows and lower jet velocities than the pure jet engine, thus giving a propulsive efficiency which is comparable to that of the turbo-prop and exceeds that of the pure jet engine (fig. 21).

Fig. 21 Specific Fuel Consumption: In order to compare different types of turbine engines it is necessary to use common units. Fuel flow/pound of thrust / hour----------------------- turbojet Fuel flow/shaft horsepower / hour--------------------- turboprop SFC is defined as the amount of fuel required to produce one pound of thrust in one hour ---- turbojet or the amount of fuel required to produce one shaft horsepower in one hour ---- turboprop Training Purposes Only

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Examples: CFM56-5B2

0·611 (Cruise)

V2500-A5

0·606 (Cruise)

PW2136

0·565 (Cruise)

RB211-535E4

0·598 (Cruise)

Thrust horsepower (THP) = thrust (lbs) * Aircraft Speed (ft/Sec) / 550 ------------ TurboJet Thrust horsepower (THP) = Shaft horsepower * Propeller Efficiency

------------Turboprop

1.10- ENGINE STATION DESIGNATIONS Station designations are assigned to the various sections of gas turbine engines to enable specific locations within the engine to be easily and accurately identified. The station numbers coincide with position from front to rear of the engine and are used as subscripts when designating different temperatures and pressures at the front, rear, or inside of the engine. For engine configurations other than the picture below should be made to manuals published by the engine manufacturer.

Fig. 22

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N = Speed ( rpm or percent ) N1 = Low Pressure Compressor Speed N2 = High pressure Compressor Speed N3 = Free Turbine Speed P = Pressure T = Temperature t = Total EGT = Exhaust Gas Temperature EPR = Engine Pressure Ratio EPR = Pt7 / Pt2 Ex.: Pt 2 = Total Pressure at Station 2 ( low pressure compressor inlet ) Pt 7 = Total Pressure at Station 7 ( turbine discharge total pressure )

Fig. 23

1.11. ADVANTAGES OF TURBINE ENGINES Keeping in mind the basic theory of turbine engines, compare the advantages and disadvantages of the turbine engine with the piston or reciprocating engine. •

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Power-to-weight ratio. Turbine engines have a higher power-to-weight ratio than reciprocating engines. An example of this is the T55-L-l11. It weighs approximately 650 pounds and delivers 3,750 shaft horsepower. The power-to-weight ratio for this engine is 5.60 shp per pound, where the average reciprocating engine has a power-toweight ratio of approximately .67 shp per pound. Less maintenance. Turbine engines require less maintenance per flying hour than reciprocating engines generally do. As an aircraft maintenance Engineer, this advantage will appeal to you because of a greater aircraft availability and lower maintenance hour to flying hour ratio. The turbine engine also has fewer moving parts than a reciprocating engine; this is also an advantage over the reciprocating engine. Less drag. Because of the design, the turbine engine has a smaller frontal area than the reciprocating engine. A reciprocating engine requires a large frontal area which causes a great deal of drag on the aircraft. Turbine engines are more streamlined in design, causing less drag. Cold weather starting. The turbine engine does not require any oil dilution or preheating of the engine before starting. Also, once started, the reciprocating engine takes a long time to warm up to operating temperatures, whereas the turbine engine starts readily and is up to operating temperature immediately. Low oil consumption. The turbine engine, in general, has a lower rate of oil consumption than the reciprocating engine. The turbine engine does not require the oil reservoir capacity to be as large as the reciprocating engine's; because of this, a weight and economy factor is an additional advantage.

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1.12. DISADVANTAGES OF TURBINE ENGINES The disadvantages of the turbine engine are discussed in the following subparagraphs. •

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Foreign object damage. One of the major problems faced by the turbine engine is foreign object damage (FOD). A turbine engine requires tremendous quantities of air. This air is sucked into the engine at extremely high velocities, and it will draw up anything that comes near the inlet area. High temperatures. In the combustion chamber, the temperature is raised to about 3, 500° F. in the hottest part of the flame. Because this temperature is above the melting point of most metals, proper cooling and flame dilution must be employed at all times to insure that the engine is not damaged. Slow acceleration. The acceleration rate of a turbine engine is very slow in comparison with that of a reciprocating engine. The pilot must be aware of the time lag in the turbine engine acceleration between the instant when power is requested and when power is available. High fuel consumption. Turbine engines are very uneconomical when it comes to the amount of fuel they consume. The Lycoming T53 turbine engine, for instance, uses approximately 1.5 gallons per minute of fuel. Compare it to a reciprocating engine of approximately the same horsepower which has a fuel consumption rate of 1 gallon per minute. Cost. The initial cost of a turbine engine is very high when compared to the cost of a reciprocating engine. For example the T53-L-13B engine costs about $63,000, and the cost of a reciprocating engine of approximately the same horsepower is $20,000.

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Chapter 2-Engine Construction



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Air inlet Section Compressor Combustion Chamber Turbine Exhaust Afterburners

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2.1-Air Inlet Section The air inlet duct must provide clean and unrestricted airflow to the engine. Clean and undisturbed inlet airflow extends engine life by preventing erosion, corrosion, and foreign object damage (FOD). The amount of air required by a gas turbine engine is approximately ten times that of a reciprocating engine. The air inlet is generally a large, smooth aluminum duct which must be designed to conduct the air into the compressor with minimum turbulence and restriction. The air inlet section may have a variety of names according to the desire of the manufacturer. It may be called the front frame and accessory section, the air inlet assembly, the front bearing and shroud assembly, or any other term descriptive of its function. Usually, the outer shell of the front frame is ed to the center portion by braces that are often called struts.

Fig. 1 Air Inlet

The anti-icing system directs compressor discharge air into these struts. The temperature of this air prevents the formation of ice that might prove damaging to the engine. 2.2-Compressor Section The compressor is the section of the engine that produces an increase in air pressure. It is made up of rotating and stationary vane assemblies. The first stage compressor rotor blades accelerate the air rearward into the first stage vane assemblies. The first stage vane assemblies slow the air down and direct it into the second stage compressor rotor blades. The second stage compressor rotor blades accelerate the air rearward into the second stage vane assemblies, and so on through the compressor rotor blades and vanes until air enters the diff section. The highest total air velocity is at the inlet of the diff. As the air es rearward through the diff, the velocity of the air decreases and the static pressure increases. The highest static pressure is at the diff outlet. The compressor rotor may be thought of as an air pump. The volume of air pumped by the compressor rotor is basically proportional to the rotor rpm. However, air density, the weight of a given volume of air, also varies this proportional relationship. The weight per unit volume of air is affected by temperature, compressor air inlet pressure, humidity, and ram air pressure (free stream air pressure provided by the forward motion of the engine). If compressor air inlet temperature is increased, air density is reduced. If compressor air inlet pressure is increased, air density is increased. If humidity increases, air density is decreased. Humidity, by comparison with temperature, and pressure changes, has a very small effect on density. With increased forward speed, ram air pressure increases and air temperature and Training Purposes Only

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pressure increase. The weight of air pumped by the compressor rotor is determined by rpm and air density. Compressor efficiency determines the power necessary to create the pressure rise of a given airflow, and it affects the temperature change which takes place in the combustion chamber. There are three basic compressors used in gas turbine engines: the centrifugal-flow, the axialflow, and axial-centrifugal-flow compressors. The axial-centrifugal-flow compressor is a combination of the other two and operates with characteristics of both. •

Centrifugal-flow compressor. Figure 2 shows the basic components of a centrifugalflow compressor: rotor, stator, and compressor manifold.

Fig. 2 Typical Single-stage Centrifugal Compressor As the impeller (rotor) revolves at high speed, air is drawn into the blades near the center. Centrifugal force accelerates this air and causes it to move outward from the axis of rotation toward the rim of the rotor where it is forced through the diff section at high velocity and high kinetic energy. The pressure rise is produced by reducing the velocity of the air in the diff, thereby converting velocity energy to pressure energy. The centrifugal compressor is capable of a relatively high compression ratio per stage. This compressor is not used on larger engines because of size and weight. Because of the high tip speed problem in this design, the centrifugal compressor finds its greatest use on the smaller engines where simplicity, flexibility of operation, and ruggedness are the principal requirements rather than small frontal area and ability to handle high airflows and pressures with low loss of efficiency. Training Purposes Only

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Axial-flow compressor. The air is compressed, as the name implies, in a direction parallel to the axis of the engine. The compressor is made of a series of rotating airfoils called rotor blades, and a stationary set of airfoils called stator vanes. A stage consists of two rows of blades, one rotating and one stationary. The compression ratio of each stage is 1:1.2. The entire compressor is made up of a series of alternating rotor and stator vane stages as shown in fig.3.

Fig. 3 Axial-flow Compressor.

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Axial flow compressors have the advantage of being capable of very high compression ratios with relatively high efficiencies; see figure 4. Because of the small frontal area created by this type of compressor, it is ideal for installation on high-speed aircraft. Unfortunately, the delicate blading and close tolerances, especially toward the rear of the compressor where the blades are smaller and more numerous per stage, make this compressor highly susceptible to foreign-object damage. Because of the close fits required for efficient air-pumping and higher compression ratios, this type of compressor is very complex and very expensive to manufacture. For these reasons the axial-flow design finds its greatest application where required efficiency and output override the considerations of cost, simplicity, and flexibility of operation. However, due to modern technology, the cost of the small axial-flow compressor, used in aircraft, is coming down.

Fig. 4 Compressor Efficiencies and Pressure Ratios. Compressor Construction Centrifugal - flow compressors: are usually made of titanium. The diff is generally manufactured of a stainless steel alloy. Axial-flow compressors: The rotor blades are generally cast of stainless-steel alloy. Some manufacturers use molybdenum coated titanium blades to dampen vibrations on some stages of rotor blades. 2.3-Combustion Chamber: Today, three basic combustion chambers are in use. They are the annular combustion chamber, the can type, and the combination of the two called the can-annular. The combustion section contains the combustion chambers, igniter plugs, and fuel nozzles or vaporizing tubes. It is designed to burn a fuel-air mixture and deliver the combusted gases to the turbine at a temperature which will not exceed the allowable limit at the turbine inlet. Fuel is introduced at the front end of the burner in a highly atomized spray from the fuel nozzles. Combustion air flows in around the fuel nozzle and mixes with the fuel to form a correct fuel-air mixture. This is called primary air and represents approximately 25 percent of total air taken into the engine. The fuel-air mixture which is to be burned is a ratio of 15 parts of air to 1 part of fuel by weight. The remaining 75 percent of the air is used to form an air Training Purposes Only

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blanket around the burning gases and to lower the temperature. This temperature may reach as high as 3500° F. By using 75 percent of the air for cooling, the temperature operating range can be brought down to about half, so the turbine section will not be destroyed by excessive heat. The air used for burning is called primary air- and that for cooling is secondary air. Igniter plugs function only during starting, being cut out of the circuit as soon as combustion is self-ing. On engine shutdown, or, if the engine fails to start, the combustion chamber drain valve, a pressure-actuated valve, automatically drains any remaining unburned fuel from the combustion chamber. All combustion chambers contain the same basic elements: a casing or outer shell, a perforated inner liner or flame tube, fuel nozzles, and some means of initial ignition. The most severe operating periods in combustion chambers are encountered in the engine idling and maximum rpm ranges. Sustained operation under these conditions must be avoided to prevent combustion chamber liner failure. The annular-type combustion chamber shown in fig. 5 . The annular combustion chamber permits building an engine of a small and compact design.

Fig. 5 Annular-type Combustion Chamber. Training Purposes Only

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The can-type combustion chamber is one made up of individual combustion chambers. This type of combustion chamber is so arranged that air from the compressor enters each individual chamber through the adapter.

Fig. 6 Can-type Combustion Chamber Can-annular combustion chamber. This combustion chamber uses characteristics of both annular and can-type combustion chambers. The can-annular combustion chamber consists of an outer shell, with a number of individual cylindrical liners mounted about the engine axis as

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shown in figure Fig. 7 Can-Annular Combustion Chamber. Training Purposes Only

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2.4- Turbine Section: A portion of the kinetic energy of the expanding gases is extracted by the turbine section, and this energy is transformed into shaft horsepower which is used to drive the compressor and accessories. In turboprop and turboshaft engines, additional turbine rotors are designed to extract all of the energy possible from the remaining gases to drive a powershaft. Types of Turbines. Gas turbine manufacturers have concentrated on the axial-flow turbine shown in Fig.9. This turbine is used in all gas-turbine-powered aircraft today. However, some manufacturers are building engines with a radial inflow turbine, illustrated in fig.10. The radial inflow turbine has the advantage of ruggedness and simplicity, and it is relatively inexpensive and easy to manufacture when compared to the axial-flow turbine. The radial flow turbine is similar in design and construction to the centrifugal-flow compressor described earlier. Radial turbine wheels used for small engines are well suited for a higher range of specific speeds and work at relatively high efficiency.

Fig. 8 Axial-flow Turbine Rotor.

Fig.9 Radial Inflow Turbine. The axial-flow turbine consists of two main elements, a set of stationary vanes followed by a turbine rotor. Axial-flow turbines may be of the single-rotor or multiple-rotor type. A stage consists of two main components: a turbine nozzle and a turbine rotor or wheel.

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Fig. 10 Multiple-rotor, Multiple-stage Turbine. Turbine Construction: The turbine rotor is one of the most highly stressed parts in the engine. It operates at a temperature of approximately 1,700° F. Because of the high rotational speeds, over 40,000 rpm for the smaller engines, the turbine rotor is under severe centrifugal loads. Consequently, the turbine disk is made of specially alloyed steel. 2.5- Exhaust Section: The hot gases are exhausted overboard through the exhaust diff section. Internally, this section s the power turbine and aft portion of the Powershaft. An exhaust system that es the turbine discharge gases to atmosphere at a velocity, and in the required direction, to provide the resultant thrust. The velocity and pressure of the exhaust gases create the thrust in the turbo-jet engine but in the turbo-propeller engine only a small amount of thrust is contributed by the exhaust gases, because most of the energy has been absorbed by the turbine for driving the propeller.

Fig. 11 Basic Exhaust System Training Purposes Only

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The temperature of the gas entering the exhaust system is between 550°C and 850°C according to the type of engine. Turbo-propeller and by engines have the coolest exhaust gas flow. The use of a thrust reverser, noise suppressor and a variable-area propelling nozzle entails a more complicated system. The by engine may also include a mixer unit to encourage a thorough mixing of the hot and cold gas streams. Turboshaft engines used in helicopters do not develop thrust by use of the exhaust duct. If thrust were developed by the engine exhaust gas, it would be impossible to maintain a stationary hover; therefore, helicopters use divergent ducts. These ducts reduce gas velocity and dissipate any thrust remaining in the exhaust gases. On fixed wing aircraft, the exhaust duct may be the convergent type, which accelerates the remaining gases to produce thrust which adds additional shaft horsepower to the engine rating.

Fig. 12 Divergent Exhaust Duct.

2.6-Afterburners In addition to the basic components of a gas turbine engine, one other process is occasionally employed to increase the thrust of a given engine. Afterburning (or reheat) is a method of augmenting the basic thrust of an engine to improve the aircraft takeoff and climb performance.

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Afterburning consists of the introduction and burning of raw fuel between the engine turbine and the jet pipe propelling nozzle, utilizing the unburned oxygen in the exhaust gas to combustion. The resultant increase in the temperature of the exhaust gas increases the velocity of the jet leaving the propelling nozzle and therefore increases the engine thrust. This increased thrust could be obtained by the use of a larger engine, but this would increase the weight, frontal area and overall fuel consumption. Afterburning provides the best method of thrust augmentation for short periods. Afterburners are very inefficient as they require a disproportionate increase in fuel consumption for the extra thrust they produce. Afterburning is used in cases where fuel efficiency is not critical, such as when aircraft take off from short runways, and in combat, where a rapid increase in speed may occasionally be required.

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Fig. 13 Afterburner Engine Terminology: Directional references are shown in Fig. 15.

Fig.14 Directional References

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Foreign Object Damage (FOD) Due to the design and function of a turbine engine’s air inlet, the possibility of ingestion of debris always exists. This causes significant damage, particularly to the compressor and turbine sections. When ingestion of debris occurs, it is called foreign object damage (FOD). Typical FOD consists of small nicks and dents caused by ingestion of small objects from the ramp, taxiway, or runway, but FOD damage caused by bird strikes or ice ingestion also occur. Sometimes FOD results in total destruction of an engine. Prevention of FOD is a high priority. Some engine inlets have a tendency to form a vortex between the ground and the inlet during ground operations. A vortex dissipater may be installed on these engines. Other devices, such as screens and/or deflectors, may also be utilized. Preflight procedures include a visual inspection for any sign of FOD. Turbine Engine Hot/Hung Start When the EGT exceeds the safe limit of an aircraft, it experiences a “hot start.” It is caused by too much fuel entering the combustion chamber, or insufficient turbine rpm. Any time an engine has a hot start, refer to the appropriate maintenance manual for inspection requirements. If the engine fails to accelerate to the proper speed after ignition or does not accelerate to idle rpm, a hung or false start has occurred. A hung start may be caused by an insufficient starting power source or fuel control malfunction. Compressor Stalls Compressor blades are small airfoils and are subject to the same aerodynamic principles that apply to any airfoil. A compressor blade has an angle of attack which is a result of inlet air velocity and the compressor’s rotational velocity. These two forces combine to form a vector, which defines the airfoil’s actual angle of attack to the approaching inlet air. A compressor stall is an imbalance between the two vector quantities, inlet velocity and compressor rotational speed. Compressor stalls occur when the compressor blades’ angle of attack exceeds the critical angle of attack. At this point, smooth airflow is interrupted and turbulence is created with pressure fluctuations. Compressor stalls cause air flowing in the compressor to slow down and stagnate, sometimes reversing direction. Compressor stalls can be transient and intermittent or steady and severe. Indications of a transient/intermittent stall are usually an intermittent “bang” as backfire and flow reversal take place. If the stall develops and becomes steady, strong vibration and a loud roar may develop from the continuous flow reversal. Often, the flight deck gauges do not show a mild or transient stall, but they do indicate a developed stall. Typical instrument indications include fluctuations in rpm and an increase in exhaust gas temperature. Most transient stalls are not harmful to the engine and often correct themselves after one or two pulsations. The possibility of severe engine damage from a steady state stall is immediate. Recovery must be accomplished by quickly reducing power, decreasing the aircraft’s angle of attack, and increasing airspeed. Although all gas turbine engines are subject to compressor stalls, most models have systems that inhibit them. One system uses a variable inlet guide vane (VIGV) and variable stator vanes, which direct the incoming air into the rotor blades at an appropriate angle. Flameout A flameout occurs in the operation of a gas turbine engine in which the fire in the engine unintentionally goes out. If the rich limit of the fuel/air ratio is exceeded in the combustion chamber, the flame will blow out. This condition is often referred to as a rich flameout. It generally results from very fast engine acceleration, in which an overly rich mixture causes the fuel temperature to drop below the combustion. Training Purposes Only

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This may be due to prolonged unusual attitudes, a malfunctioning fuel control system, turbulence, icing or running out of fuel. Symptoms of a flameout normally are the same as those following an engine failure. If the flameout is due to a transitory condition, such as an imbalance between fuel flow and engine speed, an airstart may be attempted once the condition is corrected. AIRCRAFT ENGINE FUELS KEROSENE The fuel generally used in civil turbo-jet and turbo-propeller engines is known as Jet A1 (Specification D Eng. RD 2494). It has a relative density of approximately 0·8, a high flash point, and does not give off easily ignitable vapours at normal ground temperatures. It is a ‘narrow-cut’ kerosene. TYPICAL JET FUELS JET A

A kerosene type fuel with a freezing point around -40°C. It is available only in the U.S.A. SG Range = 0·775 to 0·83

JET A1

See below

JET B

This is a wide range distillate known as a wide cut gasoline. Not in common use. SG Range = around 0·76

JP 4

This is a wide range distillate known as a wide cut gasoline. When certain additives are present it may be known as AVTAG. For military use. SG Range = around 0·76

JP 5

High flash point kerosene mainly for aircraft carrier use. May be known as AVCAT. SG Range = around 0·83

PROPERTIES OF JET A1 FLASH POINT

38°C Minimum

(The flash point is the lowest temperature of a flammable liquid at which an ignition in the gas phase initiated by a test flame will propagate) SPECIFIC GRAVITY

0·81 at 15°C

CALORIFIC GRAVITY

18,560 BThU/lb OR 150,400 BThU/gallon

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VISCOSITY-

from 22 Centistokes at -60°C to 1·2 Centistokes at +43°C

FREEZING TEMPERATURE

-40°C maximum.

Note: It must be ed that kerosene is hygroscopic. This means that it will attract and absorb moisture from the atmosphere. The amount of water saturation content in kerosene at 65°C can be as much as 0·045% by weight. When kerosene freezes it does not become a solid block (like pure water) but becomes crystalline which is pumpable but clogs filters. Individual crystals of ice in the fuel can measure up to 2µ across but, under very slow freezing conditions, can measure up to 5µ. Volatility: means the degree of ease at which a liquid can be changed into a vapour. The greater the volatility, the lower the ambient temperature required to convert it to a gaseous state. Furthermore, the lower the ambient pressure on the liquid, the lower the temperature at which it will boil (turn into vapour. Warning: Skin with kerosene, and other aviation fuels, can give rise to a skin reaction, or dermatitis, which may, in certain instances, be serious. This reaction of the skin will vary from person to person and may depend upon the duration of . There is a greater danger of a serious reaction if the clothing, or anything in with the skin, is soaked with the fuel and remains in . Immediate first aid treatment is, therefore, important and should consist of removing contaminated clothing, etc., as soon as possible followed by copious washing of the skin with COLD WATER ONLY. This can be followed later by the use of soap and water using normal washing facilities.

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Fuel Control

Power Management Control (PMC) and Main Engine Control (MEC)

Fig. PMC and MEC

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Full Authority Digital Engine Control A FADEC system performs the functions necessary to control the engine during start up and in flight. It eliminates the problems associated with hydromechanical engine control systems such as wear in control rods, valves and shafts. FADEC is connected to the electrical Flight control system and AFS via the digital data bus and fully compatible with FBW systems. Basically fuel flow is controlled to achieve the required thrust (as determined by the throttle settings) by monitoring and controlling either: Engine pressure ratio (EPR) or N1 low pressure compressor or fan speed On modern high by engines the relationship between EPR and thrust is more linear than that between N1 and thrust and is more likely to be used. However, with the N1/thrust relationship known, a modern computer can easily use N1 as the controlled parameter. In some systems EPR is the primary controlled parameter with N1 as the back up. The FADEC system is fully redundant and hence is failure tolerant. This basic operation is illustrated by the following much simplified closed loop system block diagram. EPR or N1

REDUNDANT FAD EC SYST EM

T LA

FUEL FLO W CO NT ROL

ENG INE

The thrust lever is set by the pilot and a signal representing the thrust lever angle (TLA) fed to the FADEC as a demand signal. The FADEC computer calculates the required fuel flow and feeds this control signal to the engine. The closed loop control is achieved by feeding the engine speed back to the FADEC computer. More of the complexity of the system is shown in the following diagram:

AIR DAT A

FADEC A

G EOMET RY CO NT RO L

CO MPRESSOR CO NTR OL

FADEC B

FUEL FLOW CO NT RO L

FUEL MET ERIING UNIT

TLA AIR DAT A TLA

N1 O R EPR

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ENGINE

N1 O R EPR

Since the altitude, Mach number, airspeed and total air temperature all affect engine thrust these quantities are fed to the FADEC system from the Air Data Computer (ADC) to modify the control signals. For maximum efficiency the compressor air flow geometry can also controlled by automatically adjusting variable inlet guide vanes and variable geometry vanes within the first few stages of the compressor. Also automatic adjustment of blade clearance can be effected for compressor and/or turbine efficiency. is achieved by monitoring, and hence controlling, engine speed or EPR. Use of FADEC reduces crew workload, achieves optimum operation for efficiency and economy and prevents over speed and over temperature conditions. On modern engines the FADEC system is the normal method of control. A TYPICAL FADEC SYSTEM Overview (A typical system) This example is of an aircraft fitted with FADEC controlled twin high by (5:1), two spool, axial flow turbofan engines. The fan acts as a single stage LP compressor directing the incoming air into a by stream and a core stream. The core stream is fed through variable inlet guide vanes (IGV) into a 14 stage HP compressor. The first five stages of the compressor employ variable geometry vanes. Each engine has two FADECs designated ‘A’ and ‘B’. All signals between the FADECs and the corresponding engine and between the FADECs and the airplane are completely redundant. FADECs are interconnected by a Cross-Channel Data Link (CCDL) to share engine data and FADEC status. One FADEC is automatically in standby mode. While in standby a FADEC will still monitor all inputs, perform all computations and perform built in test and fault detection routines however the command output drivers are de-energized. The active FADEC has complete control over the engine response to thrust lever settings and both external and engine conditions. A diagram of the FADEC interfaces for one engine and one FADEC is shown on the next page.

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T O O T HER ENG INE FADEC A

ARINC-429 ENG INE INDICAT ING & CREW ALERT ING SYSTEM (EICAS)

ARINC-429

FRO M O THER ENG IN E FADEC A

CCDL FADEC B

AIR DATA CO MPUT ER (ADC)

ARINC-429 ENG INE SENSO RS A

T HRUST LEVER

FUEL PUMP AND METERING UNIT (FPMU) CO CKPIT DISCRETES

FADEC A

CO MPRESSO R CON TRO L

AIRCRAFT DO SRET ES

IGNIT IO N EXCITER A

AIRCRAFT RELAYS

FADEC ID JUMPERS

PERMANEN T MAG NET ALT ERNATO R (PMA)

AIRCRAFT PO W ER SUPPLY

FADEC INPUTS The following inputs are fed to the FADECs: Internacelle

Air data

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ARINC429 bus to exchange data between active FADECs for thrust reverser interlock and thrust control functions. • • • •

Altitude (P ALT) Mach Number (MN) Indicated Airspeed (IAS) Total Air Temperature (TAT) from an ADC via an ARINC 429 bus 43

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Thrust lever angle

Cockpit discrete

A resolver analogue feed. The resolver excitation is supplied by the corresponding FADEC.

• • • • •

Thrust ratings Thrust mode selector (take off, climb, cruise, maximum continuous) Engine start/stop Ignition on/off Anti-ice on/off FADEC reset (to clear FADEC faults) Alternate FADEC select

Aircraft discrete

• • • •

Air/Ground (Weight-on-wheels – WOW) Landing gear down Thrust reverser deployed Thrust reverser stowed

Alternate FADEC

• • •

FADEC in control FADEC capable Shut off interlock. these being hardwired status signals duplicating data fed on the CCDL.

Engine sensors

FPMU

ID Jumpers

• •

N1 primary fan speed N1 secondary fan speed N2 primary HP rotor speed P2.5 compressor inlet pressure T2.5 compressor inlet temperature ITT inter Stage turbine temperature P0 ambient static air pressure Compressor variable geometry (CVG) (VBV) position Main Metering Valve (MMV) position from a liner variable induction transducer (LVIT) • • • • • • • •

Selection of: • • •

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Engine left/right FADEC A/B Parity

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EEC

Programming Plug

A/C Power supply PMA

• • • •

Engine Serial Number EPR modification Data Engine Performance package Variable Stator Vane Schedule VSV VIGV

28 VDC essential bus • •

3 phase AC N2 secondary HP rotor speed

All discrete signals have a 10kΩ resister across the s and are supplied with a current limited excitation voltage from the FADEC so allowing the FADEC to detect open and short circuit failure conditions in the aircraft engine wiring. Since the discrete signals and the TLA resolver excitation is fed from the FADEC and an engine driven PMA (Permanent Magnet Alternator) is used for power then, with the exception of the ADC input, the aircraft/FADEC interface is independent of the aircraft electrical power supply. So, even with complete aircraft power failure, the pilot would still control the engines in the normal way, i.e. through the FADECs. FADEC Outputs The following outputs are fed from the FADECs: Internacelle

ARINC429 bus to exchange data between active FADECs for thrust reverser interlock and thrust control functions.

EICAS

ARINC429 bus carrying • • •

engine parameters engine status maintenance fault data

Alternate FADEC

• • •

FADEC in control FADEC capable Shut off interlock. these being hardwired status signals duplicating data fed on the CCDL.

A/C Relays

• •

Starter shut off (N2 speed switch) Environmental Control System (ECS) off (reduced bleed request)

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FPMU

• • • •

Engine fuel flow rate (WF) torque motor Latching shut off valve (LSOV) start/stop CVG (Compressor Variable Geometry) torque motor Air vent solenoid

Ignition Exciter

•

Ignition relay energise (off)

Active Clearance Control

•

Signal to direct cooling air to expand or contract the turbine casing to maintain turbine tip clearance blade.

CONTROL HARNESSES Each engine control system includes four harnesses, two designated internal, clamped to the right and left side of the high pressure compressor and two designated external, routed right and left of the engine. The harnesses are labelled A or B and are dedicated to the corresponding FADEC. POWER SUPPLY Each FADEC has two power supplies one from a permanent magnet alternator (PMA) and the other from 28VDC essential bus. During engine starts the essential bus provides power but as engine speed increases the PMA output increases and is automatically connected to the FADEC power supply input at 50% N2 and above (in some installations it is 15% or less RPM). With engine shutdown, low speed and alternator failure FADEC is powered from aircraft electrical power supply. The PMA is driven by the accessory drive gearbox. It has four separate windings, two 3 phase and two single phase. One 3 phase winding provides power to channel ‘A’ the other to channel ‘B’ of the FADECs. The single phase windings feed the redundant exciters in the ignition system. One phase of the 3 phase supply is used to derive a secondary N2 signal. CONTROL FUNCTIONS

EN GINE PARAMET ER S

FAD EC P ALT

IG N ITIO N R ELAY

MN TAT IAS

T HR UST SCH ED ULING

T LA

N 1 REQ UEST

MMV PO SIT IO N

EN GIN E C ON T RO L LO G IC W F DEMAN D

AIRC RAFT D ISCR ET ES

EN G INE WF

FPMU C VG DEMAND

C VG

The simplified diagram above illustrates the principle of engine control in this representative system. The controlled parameters and actions are: Training Purposes Only

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Fan speed (N1)

Closed loop control to set the fan speed equal to that computed by the thrust scheduling logic. Any difference between desired and actual fan speed produces an error signal which commands the fuel flow (WF) so as to reduce the error to zero. The requested fan speed is scheduled taking into TLA, air data and aircraft discrete signals.

Engine transient rate limiting

Engine acceleration and deceleration limits are programmed into the FADEC logic. These restrict the rate of change of commanded fuel flow to prevent a surge on acceleration or a lean blowout on deceleration.

CVG scheduling and control

The FADEC stores a schedule of CVG position versus corrected N2 so as to achieve optimum compressor efficiency. CVG position , fed from the CVG actuator on the engine, is compared with the CVG demand to achieve closed loop control. The IGVs and first 5 stages of CVG are controlled.

Engine steady state limiting

Fuel flow is limited to prevent overspeed of the LP and HP rotors. There is also an HP underspeed logic to prevent fuel flow being reduced to a value which would result in an N2 underspeed. A fuel flow limiting function is also incorporated which prevents the ITT exceeding a maximum allowable value of 1690 oF (921oC).

Automatic engine starting sequencing

Engine starts are automatically controlled by the FADEC system which sequences the ignition system and introduction of fuel to give light off and acceleration to idle. The sequence is initiated by a start discrete from the powerplant control .

Engine protection shut down

Should N1 exceed 105% (approx 9135 rpm) or N2 exceed 105.5% (approx 16750 rpm) the overspeed logic shuts the engine down if both FADECs agree. Engine shutdown also occurs if N2 is less than 54% (approx 8500 rpm.)

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IGNITION

A simplified diagram of the FADEC controlled ignition circuit is shown below.

EXCITER

FADEC RECTIFIER AND STORAGE CAPACITOR

PMA IGNITER LEAD ASSEMBLIES

EXCITER

FADEC RECTIFIER AND STORAGE CAPACITOR

IGNITER

For normal operation the igniters are only required to operate during engine start after which combustion is self sustaining. However, further operation will be necessary for a restart and continuous operation may be selected for flight in rain, hail or snow where combustion is more likely to cease. Each engine, in our representative system, has a dual redundant ignition system comprising two exciters, two high-tension ignition leads and two igniters. Each exciter is controlled by a separate FADEC. For ground starts only one exciter is used being selected alternatively for consecutive ground starts. For in-flight restarts both ignition systems are energized. The primary supply for the ignition system is a single phase winding in the PMA which provides 20-40 VAC at all N2 speeds above 10%. Power is stored in a capacitor and discharged four to seven times per second. The relays are shown de-energised which gives an ignition on condition. This is a fail safe arrangement. The ignition switches on the Powerplant Control , mounted on the overhead , can be set to one of three positions: ON

Power is removed from the exciter relay and continuous ignition operation takes place as a safety precaution when flame out is possible.

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AUTO

The exciter relay is de-energised during start up to provide the spark necessary to begin combustion. Once combustion is established the relay is energised so removing the supply to the igniters. If combustion fails in the air, the FADEC adjusts the fuel flow and compressor variable geometry then energizes the ignition for an automatic relight. On the ground the FADEC will initiate an automatic shutdown if no combustion is detected.

OFF

The FADEC energizes the exciter relay so stopping ignition.

An engine start is initiated by momentarily holding the STOP-RUN-START switch in the start position. ENGINE THRUST RATING MODES There are four thrust rating modes controlled by buttons on the overhead and on the control pedestal. They define the available thrust at the existing ambient and airspeed conditions. T/O-1

maximum take off

(100% N1)

CON

maximum continuous

(86% N1)

CLB

maximum climb

(76% N1)

CRZ

maximum cruise

(70% N1)

If the thrust levers are positioned at THRUST SET then the FADECs command the maximum N1 associated with the selected mode. If the thrust levers are set between THRUST SET and IDLE then the FADECs command an intermediary thrust. Both engines operate normally in the same rating mode. If the engines do not agree on thrust mode selection for a given time interval the maximum take off mode will be selected for both engines. FADEC Switching

The following controls, one for each FADEC, are on the Powerplant Control (P): RESET/ALTN

RESET clears recorded faults unless a fault is currently being detected in which case it is not cleared. ALTN manually switches the active and standby FADECs unless the selected FADEC is not capable of controlling the engine. This overrides the automatic switching from active to standby when a fault is detected.

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REVERSE THRUST The FADEC system interfaces with the thrust reverser system of the corresponding engine. Each FADEC receives the following information: Stowed If all thrust reverser doors are stowed Deployed If all thrust reverser doors are deployed The FADEC enables reverse thrust depending on the position of the reverser doors and the positioning of the thrust lever. Engine thrust is reduced to IDLE if there is an indication of inadvertent thrust reverser deployment. FUEL CONTROL SYSTEM The fuel control system supplies filtered and measured fuel for combustion. It also supplies pressurised fuel to operate the CVG system. See figure on following page. The heart of the system is the electro-mechanical Fuel Pump and Metering Unit (FPMU) which is controlled by a FADEC. Pressurised and filtered fuel is fed to the servo operated main metering valve (MMV). The servo valve is biased towards the closed position so, in the absence of a torque motor signal from either FADEC, the fuel flow to the engine will be cut off. The FADEC controlled metered fuel flow is fed to the fuel nozzles via the pressure raising valve (PRV) which, by spring loading, ensures a minimum gear pump pressure at low engine speeds. is from a linear variable inductance transducer (LVIT). A latching two position shut off valve (LSOV) is STOP controlled by two dual-coil torque motors. At 23% N2 in the engine start sequence the LSOV START solenoid is energised and fuel pressure from the LSOV to the PRV is removed so opening the PRV and allowing fuel to be supplied from the MMV to the fuel nozzles. An air vent valve automatically vents entrapped air or fuel vapour during engine starting and motoring. A FADEC controlled solenoid valve is commanded to open at the initiation of the start sequence. The valve remains closed whenever the solenoid is not energised so preventing fuel leakage from the system if the aeroplane booster pumps are turned on when the engine is not running.

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A

FADEC

B

LVDT COMPRESSOR VARIABLE GEOMETRY ACTUATOR

FADEC A

FADEC B

B

A

TORQUE MOTOR FUEL PUMP AND METERING UNIT

COMPRESSOR VARIABLE GEOMETRY CONTROL VALVE TORQUE MOTOR

DRIVE FROM ACCESSORY GEARBOX

MAIN METERING VALVE (MMV)

FUEL PUMP ASSEMBLY

PRESSURE RAISING VALVE

TO FUEL NOZZELS

LVIT

AIR VENT VALVE

LATCHING SHUT OFF VALVE

SOLENOID

FUEL INLET

A

B

A

FADEC

B FADEC

START

STOP

A B FADEC

A B FADEC

The CVG actuator assembly comprises a series of turnbuckles, a torque tube assembly and hydraulic actuator controlled by fuel pressure from the fuel pump. The hydraulic actuator rotates the torque tube which, in turn, by means of levers and turnbuckles, moves the variable vanes in unison. The fuel supply to the actuator is fed from the CVG control valve which is operated by a torque motor under FADEC control. The fuel pressure fed to the actuator will determine the setting the angle of the inlet guide vanes and the first five variable geometry compressor vanes. is from a liner variable differential transducer (LVDT) on the CVG actuator. Lane Controller The FADEC computer has a unit inside it that decides which of the two lanes or channels will control the fuel flow and other outputs. Should the lane change element detect an error between the demand made by the thrust lever angle and the output signals from the engine, it will automatically switch lanes to disable the Training Purposes Only

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lane that is giving incorrect signals and allow the other lane to control fuel by enabling its command output drivers and disabling the drivers in the original lane. This is effected by ing demand and engine through a comparator unit and it is this unit that will signal the lane change element to switch lanes if a fault is detected.

FULL AUTHORITY FUEL CONTROLLER LANE CHANGE PROCESS The system consists of several parts: The Throttle Lever – throttle control unit transduces mechanical movement into electrical signal. It encloses the drive mechanism for potentiometers and resolvers which are protected within two cases.TLA operates in two of four quadrants: The first quadrant serves for positive angles. The fourth quadrant serves for negative angles. The other two quadrants are not used.

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The pilot will have initial control of the engine through the throttle lever. Movement of the lever will be no different from the more ‘traditional’ form of control. The difference is that, instead of operating a rod or a cable, the throttle lever is connected to a variable resistor – normally a tri-lane potentiometer, which will send information to a computer attached to the engine. A tri-lane potentiometer is used in case of failure. The signal will be sent to a ‘majority wins’ circuit’ so that, if one of the potentiometers fails, the other two will still send valid information through to the computer(s). FAFC. The Full Authority Fuel Controller (FAFC) receives the input signal from the throttle lever, The FAFC will process this information in one of its computers and compare the pilot’s requirement to existing engine parameters. Once the FAFC decides that it is possible to feed more fuel into the burners it will then open up the Fuel Metering Unit (FMU) to permit this to happen. As soon as more fuel is fed into the combustor the FAFC will measure the effect this action has had upon the engine and take the appropriate action. The action taken may be to increase the fuel feed until the selected rpm has been attained or it may be that one of the maximum operating parameters of the engine has been reached and the fuel flow will be trimmed to prevent exceeding that parameter. On many units the FAFC will also send a signal to other systems so that they can operate at their most efficient point – the airflow control system and the active clearance control systems are two of these. Other interactions are, for example, with the thrust reverser to prevent it operating in flight, the ignition system linked with the start circuits and stall warning system as well as a manual selection from the flight deck and an inhibiting circuit which will prevent the igniters firing if the engine is wind-milling in the reverse direction of rotation on start select; if idle fuel flow will be revised depending on whether the aircraft is in flight or on the ground.

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