Mod 15 turbine engine technology - b1.018 part 1


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Mod 15 turbine engine technology - b1.018 part 1

  2. 2. © Air Service Training (Engineering) LtdAERONAUTICAL ENGINEERINGTRAINING NOTESThese training notes have been issued to you on the understanding that they areintended for your guidance, to enable you to assimilate classroom and workshoplessons and for self-study. Although every care has been taken to ensure that thetraining notes are current at the time of issue, no amendments will be forwarded toyou once your training course is completed. It must be emphasised that thesetraining notes do not in any way constitute an authorised document for use inaircraft maintenance.All Rights ReservedThe copyright in these technical training notes remain the physical and intellectualproperty of Air Service Training (Engineering) Ltd, (AST). Copying, storing in hardcopy or electronic format, transmission to third parties and use for teaching byestablishments other than AST is forbidden, except with the written permission of theAST General Manager.M HaufeTraining Manager March 2006
  3. 3. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1CONTENTSPAGECHAPTER 1 : GAS TURBINESSECTION 1 : Introduction 01.01.01SECTION 2 : Introduction to Gas Physics 01.02.01SECTION 3 : Jet Mechanics 01.03.01SECTION 4 : Compressors 01.04.01SECTION 5 : Types of Intake 01.05.01SECTION 6 : Combustion Systems 01.06.01SECTION 7 : Turbines 01.07.01SECTION 8 : Exhaust System 01.08.01SECTION 9 : Materials used in Engine Construction 01.09.01SECTION 10 : Bearings 01.10.01SECTION 11 : Gas Turbine Fuel Systems 01.11.01SECTION 12 : Engine Speed Governor 01.12.01Issued May 2006 Contents 1
  4. 4. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/0182 Contents Issued May 2006
  5. 5. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1CHAPTER 1 :GAS TURBINESSECTION 1 : INTRODUCTIONThe first known instance of ‘jet propulsion’ produced by man was when Hero, aGreek engineer living in Alexandria, made a machine as a toy in the year 120 BC.However, nature uses ‘jet propulsion’ in squid, octopus and jellyfish, so it can beseen that this means of getting about is not new.Nothing more was done about developing the jet principle until 1629 when an Italian,Giovanni Branca, produced a steam driven impulse turbine (now in the BritishMuseum).Issued May 2006 Chapter 1 : Section 1 1
  6. 6. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018Sir Isaac Newton produced his laws of motion from which an inventor, Gravinade,designed and produced a steam driven carriage - a horseless carriage, which washopelessly underpowered.In 1913 a French Engineer, René Lorin, patented a jet propulsion engine. This waswhat is known as an athodyd (an aero-thermodynamic-duct), but at that time wasimpossible to manufacture or use due to the lack of appropriate materials. However,it was very similar to a modern Ram-Jet.In the year 1928, a young flying officer in the RAF named Whittle, (later Sir FrankWhittle) first suggested using a gas turbine for jet propulsion and was granted hisfirst patent in 1930. Working in co-operation with Dr A A Griffith of the Royal AircraftEstablishment (RAE), Whittle eventually produced a pure jet engine, which wasbench tested in 1937.2 Chapter 1 : Section 1 Issued May 2006
  7. 7. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1Issued May 2006 Chapter 1 : Section 1 3
  8. 8. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018In the same year, a similar engine was tested at the Heinkel factory in Germany. Theengine, designed by Pabst Von Ohain, was fitted to a Heinkel 178 and this becamethe first jet-propelled aircraft to fly in August 1939.The first British jet aircraft was the Gloster/Whittle E28/39, which flew in May 1941.In man’s quest for greater altitudes and speeds, piston engines of increased size andpower and of various configurations (12, 24, 36 and 48 cylinder, ‘V’, ‘X’ and ‘H’engines) were developed. Superchargers, water-methanol injections systems wereintroduced to improve performance. The propeller was also developed to absorb andtransmit the increased power with the introduction of variable pitch, constant speedunits, three, four and five bladed propellers, ‘cropped’ blades and as torqueincreased, the contra-rotating and co-axial propellers.The demand for better performance, power altitude and speed was handicapped bythe lack of engines of high power/weight ratio and small bulk, also as altitudeincreased, a series of fall-off in power output and reduced propeller efficiencies were4 Chapter 1 : Section 1 Issued May 2006
  9. 9. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1encountered. Thus, a new type of power unit was needed and that is where the ‘jet’engine came to the fore.The piston engine and jet engine both produce thrust by reaction. The piston engineby driving a propeller, accelerates a relatively large cold mass to a moderate velocityand the jet by accelerating a small mass to a much higher velocity.The gas turbine is a very simple engine, with few moving parts, when compared witha piston engine, giving it a high reliability factor with less maintenance. A furtheradvantage is the high power/weight ratio, about three times better than the pistonengine.The working cycle of a gas turbine engine is similar to that of a piston engine, in thatthere is an induction, compression, combustion and exhaust.Issued May 2006 Chapter 1 : Section 1 5
  10. 10. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018With the piston engine, combustion (power) is intermittent, whereas in the gasturbine, each process is continuous. The gas turbine has a separate compressor,combustion system, turbine assembly and exhaust system, with each part concernedonly with its own function, with combustion taking place at a constant pressure. Theabsence of reciprocating parts, provides a much smoother running engine of lighterconstruction, enabling more energy to be released for useful propulsive work. Thepulse power produced by a piston engine leads to increased maintenance costsbecause of the increased wear on the bearings that are subject to severe impactloadings. Furthermore, the jet engine uses a cheaper, and less dangerous, fuel. Thepower to weight ratio of the average turbopropeller engine is three times better thanits equivalent piston engine; there is reduced frontal area; the efficiency is superior.This will quickly give the realisation that the turbine engine is far more suitable tocommercial aircraft use and has brought long distance travel from being the provinceof the very rich into the range of almost everyone.That said, at speeds below approximately 450 m.p.h. the pure jet engine is lessefficient than a propeller type engine. Since its propulsive efficiency depends largelyupon its forward speed the pure jet engine is, therefore, more suitable for highforward speeds and turbopropeller engines are more suited to short range, lowspeed, low altitude use. Turboshaft engines are used for helicopters in aviation andfor sundry used in industry ranging from air conditioning units and oil/gas pumpingthrough to electricity generation and marine use.Gas turbines have come a long way since Hero first made his ‘toy’ and advancedenormously from Whittle’s and Ohain’s first prototypes. Where they were consideredpointless they now impinge on almost every aspect of modern day life.6 Chapter 1 : Section 1 Issued May 2006
  11. 11. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1NOTES:Issued May 2006 Chapter 1 : Section 1 7
  12. 12. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 2 : INTRODUCTION TO GAS PHYSICSIn order to attain a fuller understanding of the functioning of gas turbines, it isessential to have an appreciation of the basic gas laws.BOYLE’S LAW‘The volume of a given mass of gas, whose temperature is maintainedconstant, is inversely proportional to the gas pressure’.What this means is that if the pressure of a given mass of gas is doubled, its volumeis halved, or if the pressure is halved, the volume will be doubled, provided that thetemperature of the gas remains constant.CHARLES’ LAW‘If the pressure of a given mass of a gas is maintained constant, the volume ofgas increases as its temperature is increased’.These historical laws are combined in what is now called the Ideal Gas Law whichgives the relationship:222111TVPTVP=This relationship is applied to heat engines in the following manner.EFFECT OF ADDING HEAT ENERGY AT CONSTANT VOLUMEIf we heat a mass of air without allowing its volume to change, its temperature willincrease and as shown from the above equation, there will be an INCREASE INPRESSURE. This is the condition that exists in the cylinder of a piston engine.EFFECT OF ADDING HEAT ENERGY AT A CONSTANT PRESSUREIf we heat a mass of air which is not confined in volume eg. Not in an enclosedcylinder, its temperature will rise and there will be a consequent INCREASE IN THEVOLUME of the gas. The pressure will remain approximately constant. This is whathappens in a gas turbine engine.Issued May 2006 Chapter 1 : Section 2 1
  13. 13. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018ENERGYThe above laws depend upon the usage of heat energy to a large extent. It must beremembered that energy can take many forms and that it can be changed from oneform to another but never destroyed or created. The Universe contains a finiteamount of energy that, in flowing from one point to another, does work. Ultimatelythe level of energy throughout the Universe will be the same (singular) which meansthat no flow of energy will take place, no work is possible and, therefore, everythingwill cease. This flow into an even level is called ‘entropy’.Here are some forms of energy:HEAT: Heat travels by radiation (from the sun, for example), Conduction(through a metal rod) and Convection (rising air currents). Heatis regarded as ‘disorderly energy’ because of its random patternsand the difficulty of channelling it efficiently.FUEL: Fuel (or ‘chemical’) is ‘orderly energy’ because it is stored neatlyand extremely controllable in most of its forms.KINETIC: Kinetic is the energy a body (or mass) possesses by virtue of itsmotion. This is closely tied in with ‘momentum’.POTENTIAL: Potential energy is caused by the position of a body. If a body isheld suspended above the ground then it has the potential todevelop into kinetic energy. If the body is not released then thatpotential has not been realised.ELECTRICAL: Electrical is the movement of atoms from a negative charge to apositive charge. Like heat – it travels from one point to anotheruntil it ‘levels out’.PRESSURE: A fluid under compression will attempt to flow from a highpressure to a low pressure and, in doing so, will – or can, dowork.2 Chapter 1 : Section 2 Issued May 2006
  14. 14. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1THERMODYNAMICSThermodynamics is the branch of physics which explores the relationship betweenheat and the other forms of energy. This, roughly, means the ability to convert heatinto useful work. There are two basic laws involved. The first law of thermodynamicssays that the energy in to a system equals the energy you get out plus the energythat remains stored in the system.Σ Energy in = Σ Energy out + Σ Energy stored(Σ = Energy)The second law says that there is a loss of energy from a system which reduces theability of the system to do work perfectly. This loss of energy (heat) is called‘entropy’.dS = dQ/T, Btu/RdS = Rate of EntropydQ = Heat RateT = TemperatureBtu = British Thermal Units (a measurement of the quantity of heat)R = Gas constantThis also means that the flow of heat energy will only be from one temperature to alower one.THERMODYNAMIC CYCLESThe piston engine works on what is called an ‘Otto’ cycle or, more accurately, a‘Modified Otto’ cycle. The gas turbine engine works on the basis of the ‘Brayton’cycle formulated by George Brayton, an American Engineer from Boston whoproposed the idea in the late nineteenth century.Issued May 2006 Chapter 1 : Section 2 3
  15. 15. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018TEMPERATURE SCALESThe temperature scale normally used in thermodynamics is the Kelvin scale.American engine manufacturers commonly use degrees Fahrenheit when describingtheir engines. It will be useful, therefore, to remember the conversions:FOR INTERPOLATION - 1°C = 1·8°F°C = 5/9 (°F – 32) °K = °C + 273°F = (9/5°C) + 32 °R = °F + 460Where: °C = Centigrade°F = Fahrenheit°K = Kelvin°R = Rankine4 Chapter 1 : Section 2 Issued May 2006
  16. 16. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1THE BRAYTON CYCLEThe working cycle upon which the gas turbine engine functions is, in its simplest andmost ideal form, represented by the cycle shown on the pressure volume diagramabove. Point ‘A’ represents air at atmospheric pressure that is compressed along theline to ’B’. From ‘B’ to ‘C’ heat is added to the air by introducing and burning fuel at,ideally, constant pressure, thereby considerable increasing the volume of air. From‘C’ to ‘D’ the expanding gases from combustion pass through the turbine and jet pipenozzle back to atmosphere. During this part of the cycle, some of the energy in theexpanding gases is turned into mechanical power by the turbine; the remainder, onits discharge to atmosphere, provides a propulsive jet.Ideally, the compression and exhaust phases are isothermal in that the temperatureof the working fluid should remain constant. In practice the heat in the gas isconcentrated which increases the temperature of the mass flow; furthermore, thecompression and exhaust process is not 100% efficient leading to friction that willadd heat and thus raise the temperature.Adding energy to the working fluid should be adiabatic. This means that no heat islost from the mass flow ensuring that all the heat added remains to do useful work.Unhappily some heat is lost through transfer to atmosphere. The combustionprocess is also, ideally, isobaric: it will take place at a constant pressure. This is notpractical as a pressure drop is required to induce a flow of fresh air into the systemduring combustion to ensure complete fuel combustion and also for cooling thegases before they reach the exhaust section.Issued May 2006 Chapter 1 : Section 2 5
  17. 17. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018In both the ‘Otto’ and ‘Brayton’ cycle, the work done on the gas is subject to changethrough wear and thermodynamic cycling. The wear is self-evident but the cyclesmay be adequately compared to a light bulb. If an electric light bulb is turned ‘on’ -and left on, it will ‘live’ a long time but, if it is continually turned ‘off’ and ‘on’ thechange from cold to hot to cold to………….. will wear it out quite rapidly. Gas turbineengines, in particular, refer to ‘cycle’ times when their serviceability is beingexamined.Both piston and gas turbine engines have a need to compress the gas before energyis added. Since energy is being added there should be no need to compress the gasfor it to accomplish useful work. Additional pressure to the working fluid (air or air/fuelmix as appropriate) prior to the energy being added increases the efficiency of theengine. Since the majority of gas turbine engines run at much higher compressionratios they are already working at an advantage over piston engines.The limitation on compression for a piston engine is the fuel used. It has to burnrapidly without exploding but a higher compression will cause detonation (exploding)which will damage the engine. Thus the limit for a ‘normal’ aero-piston engine isaround 10:1 and the average engine will run at 7⋅5:1. Gas turbine engines have notheoretical limit. In practice there will come a point where the expense of addingpressure to the air compared to the energy released will make further compressionpointless. However, modern jet engines are running at compression ratios of over40:1. Industrial gas turbines can have lower ratios and remain efficient becausethere is no weight limit; this means that the heat left in the exhaust can be removedand put back into the intake thus increasing the efficiency. Industrial turbinescommonly run at compression ratios of 2 or 3:1.6 Chapter 1 : Section 2 Issued May 2006
  18. 18. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1JET ENGINE GAS FLOWSDuring the practical operation of the Brayton cycle the gas is variously heated andpressurised as it passes through the engine. It is this change of condition of the gasthat is realising the thrust produced by the engine. The thrust, and thrust distribution,will be looked at later. In the meantime it will be worthwhile to look at typical valuesfor the gas flow through the engine in terms of its temperature, pressure and velocity.The airflow through this particular engine, at maximum thrust, is:Dry Thrust: 164 lb/secReheated 172 lb/secThis correlates to a fuel flow and thrust of:Dry Thrust: 8,520 lb/hr (fuel) and 10,200 lb (thrust)Reheated 33,650 lb/hr (fuel) and 16,000 lb (thrust)The term ‘reheated’ refers to using the spare oxygen left over from the combustionprocess for burning in the jet pipe at the back of the engine to produce extra thrust.This compares to:The need for higher pressures and temperatures will be referred to in a later section.Issued May 2006 Chapter 1 : Section 2 7
  19. 19. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018This illustration will give an idea of the gas velocities through the engine as well asthe pressures and temperatures experienced at the various stages.ENGINE NOTATIONSIn order to simplify descriptions of the gas flow through an engine, the variousstations through the engine are annotated. At each point where the condition of thegas is changed, the pressure and temperature will be noted as a specific number.The number will change depending on the number of spools there are in the engineso that, while all engine inlet pressures and temperatures will be ‘1’, the exhaustcould be ‘4’, ‘6’ or ‘8’ for single, twin and triple spool engines respectively.The notations for a single spool are shown above and, for comparison purposes, thenotations for a low by-pass twin spool engine are shown in the next diagram.8 Chapter 1 : Section 2 Issued May 2006
  20. 20. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1With reference to the above illustration, the notation for gas leaving the second stageof the low pressure (LP) compressor would be P1·2 as the condition of the air is‘Engine Intake’ plus ‘Two Stages’ of compression. In the same way, air coming out ofthe 4thstage of the high pressure (HP) compressor will be known as P2·4.Issued May 2006 Chapter 1 : Section 2 9
  21. 21. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018BERNOULLI’S THEOREMAt any point in a tube (or a gas passage) through which liquid (or gas) is flowing, thesum of the pressure energy, the potential energy and the kinetic energy is constant.Thus, if one of the energy factors in a gas flow changes, one or both of the othervariables also changes so that the total energy remains constant.This theorem gives us the relationship between velocity and pressure of a stream ofair flowing through a tube, or duct, such as a gas turbine engine.CONVERGENT DUCTA convergent duct is one that has an area at the inlet greater than the area at theoutlet. When air flows through such a duct it incurs a velocity increase at theexpense of the static pressure and temperature.DIVERGENT DUCTA divergent duct is one which has an inlet area which is less than the outlet area.This gives a decrease in velocity with an increase in pressure and temperature.10 Chapter 1 : Section 2 Issued May 2006
  22. 22. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1CHOKED NOZZLEIn a turbo jet engine, the exit velocity of the exhaust gases is subsonic at low thrustconditions only. During most operating conditions, the exit velocity reaches thespeed of sound in relation to the Exhaust Gas Temperature (EGT) and the propellingnozzle is then said to be CHOKED; that is, no further increase in velocity can beobtained unless the temperature is increased.Issued May 2006 Chapter 1 : Section 2 11
  23. 23. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018NEWTON’S FIRST LAW OF MOTIONEvery object continues in its state of rest or uniform motion in a straight line unless itis compelled to change that state by an external force acting upon it.i.e., You cannot move, stop or steer anything unless you apply a force to it.NEWTON’S SECOND LAW OF MOTIONThe rate of change of momentum of a body is proportional to the total force actingupon it and occurs in the direction of the force.i.e., the effect of a force depends on its mass, speed and directionamf ×=12 Chapter 1 : Section 2 Issued May 2006
  24. 24. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1NEWTON’S THIRD LAW OF MOTIONOf Sir Isaac Newton’s three laws of motion, the third is most applicable to theoperation of gas turbine engines. This law states that:If body ‘A’ exerts a force on body ‘B’, then body ‘B’ exerts a force of the same size onbody ‘A’.ie., ‘for every action, there is an equal and opposite reaction’.In the operation of a gas turbine engine, the mass flow of gas is accelerated in arearward direction thus, by reaction, thrust is produced.DEFINITION OF MASSMass is a basic property of matter that (with length and time) constitutes one of thefundamental, undefined, quantities upon which all physical measurements arebased, and which is intuitively associated with the amount of matter a body contains.Generally it is associated with the force required to overcome the inertia of a bodyand is called weight when that inertia is associated with gravity.DEFINITION OF ACCELERATIONAcceleration is the rate of change of velocity of a body. Since velocity is the timetaken for a body to travel a given distance, then acceleration is the time taken tochange the velocity.TDV =2TDa =Issued May 2006 Chapter 1 : Section 2 13
  25. 25. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018DEFINITION OF MOMENTUMA measure of the quantity of motion. It is defined as product of the mass and thevelocity of a body and is determined by the length of time a constant force must acton that body to bring it to rest. Momentum is a vector quantity and is parallel to thevelocity vector. Loosely speaking, momentum is the force built up by a moving body.Momentum = Mass x Velocity14 Chapter 1 : Section 2 Issued May 2006
  26. 26. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 3 : JET MECHANICSPOWERWhen describing a turbo-prop, turbo-shaft, or piston engine, the accepted unit formeasuring the rate of doing work is HORSEPOWER. Energy is the capacity forperforming work and power is the rate of doing work. Power is measured not by theamount of work done, but by units of accomplishment correlated with time. Onehorsepower is defined as 33,000 foot-pounds of work accomplished in one minute.A foot-pound being the ability to lift a one pound weight a distance of one foot. Thus,both time and distance are necessary to compute horsepower.TDFP×=where (in appropriate units)P = PowerF = ForceD = DistanceT = TimeThis means that force times distance is the work done and dividing the work by thetime taken gives the power. It is important to note this correlation. For a constantpower input, or output, if the work required becomes less then the rate, in terms ofrpm, will increase and vice versa.When a turbo-prop or a piston engine perform work by driving a shaft that turns apropeller, TORQUE and RPM can be used to determine the horsepower that theengine is developing. Torque, in this case, is the twisting, or rotary force exerted bythe engine to turn the propeller against its drag and RPM is the number ofrevolutions per minute that the engine crankshaft is making.It is interesting to note, at this point, that, while energy is regarded in different forms,power is also seen in the same way. Electrical power and mechanical power, forexample, is the same thing and can be directly converted from one to the other.Equally, work done can be converted from heat to electrical to mechanical work.Electrical power is measured in Watts:745.7 Watts = One Horsepower (Roughly: One Horsepower is ¾ of aKilowatt)Similarly, it takes 778 ft lb of work to raise the temperature of 1 lb of water by 1°F.this amount of heat increase in the water s the equivalent of 1 British Thermal Unit(Btu) and 1 ft lb is the same as 1.356 Joules of electrical energy.Issued May 2006 Chapter 1 : Section 3 1
  27. 27. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018THRUSTThrust is the reaction to the acceleration of a mass. In the gas turbine, the mass isthe weight of air consumed by it. It now becomes clear that the working fluid of agas turbine engine is, indeed, air. It is the air flowing through the engine that has theenergy added to it so that its velocity can be increased. This acceleration is thevelocity change between the inlet and the exhaust outlet. These factors are used tocalculate the thrust developed by the engine.)( 12 VVMT −×=Where: T = Thrust in pounds1V = Initial velocity of a mass of air in ft/sec2V = Final velocity of a mass of air in ft/secM = Weight of airTHRUST AND SHAFT HORSEPOWERThe performance of the turbo-jet engine is measured in thrust, produced at thepropelling nozzle or nozzles, and that of the turbo-propeller engine is measured inShaft Horse Power (SHP) produced at the propeller shaft. However, both types aremainly assessed on the amount of thrust or SHP they develop for a given weight,fuel consumption and frontal area.Since the thrust or SHP developed is dependent on the mass of air entering theengine and the acceleration imported to it during the engine cycle, ie.THRUST = Mass of air x acceleration of airIt is obvious that such variables as the aircraft forward speed, altitude and climaticconditions will influence the value of this thrust/SHP since these are the variableswhich will affect the mass of the air entering the engine. This aspect will beexamined in more detail later. Here it is sufficient to say that the thrust developedwith the aircraft stationary is the maximum that can be obtained. As soon as theaircraft begins to move forward the value of acceleration in the equation falls and thethrust will decrease in direct proportion to the forward speed.It is possible to equate, roughly, the horsepower produced by a turbopropeller orturboshaft engine with the thrust produced by a turbofan or turbojet engine. Someturbopropellers have a small amount of residual thrust produced by the propeller. Ifthe jet thrust is divided by 2.6 it will give an equivalent horsepower figure that can beadded to the shaft horsepower to give an equivalent horsepower figure that can beadded to the shaft horsepower to give an overall power figure for the engine.This is called EQUIVALENT SHAFT HORSEPOWER (ESHP)2 Chapter 1 : Section 3 Issued May 2006
  28. 28. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1CHOKED NOZZLE THRUSTIf the air is allowed to escape in an unrestricted form from the exhaust there will be acertain amount of thrust developed by the engine. To increase the useful thrustobtained it becomes necessary to create a pressure drop at the exhaust nozzle.This is done by passing the air through the nozzle at sonic velocity. Once the airflowreaches this speed there can be no further increase in velocity. Since:Mass Flow = Air Density x Velocity)( VW ×= ρthen there can be no increase in mass flow either. This will create a pressure build-up in the exhaust which can be released across the shock wave at the nozzle andgive additional thrust.The design of the nozzle requires great care since too small a nozzle will give toogreat a pressure in the exhaust and too great a nozzle will give insufficient restrictionto create the extra thrust.The exhaust nozzle on all turbo-jet engines runs in a choked condition for most of itsoperating range. It is only ‘unchoked’ around idle rpm.The extra thrust can be calculated by determining the pressure drop across thepropelling nozzle and comparing it with the area of the nozzle.APPTp )( 0−=Where: pT = Pressure ThrustP = Pressure (in the exhaust)0P = Ambient PressureA = Nozzle AreaPOWERThe thrust or power produced by an engine is not necessarily proportional to gastemperature and RPM as factors such as ambient air temperature and pressure andforward speed also affect the power produced. Power is measured in the case of theturbo-jet in thrust and in a turbo-prop or turbo-shaft in torque. Power is indicated ona gauge in the cockpit and is also used as an indication of power deficiency. A trueindication of thrust in lbs or Kg can only be seen on the test bed. At all other timesthe indication of thrust is shown as a comparison value.For example:EPR – Engine Pressure RatioIEPR – Integrated Engine Pressure RatioIn all cases coupled with shaft RPM x Exhaust Gas Temperature.Issued May 2006 Chapter 1 : Section 3 3
  29. 29. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018VARIATIONS OF THRUST WITH SPEED, HEIGHT AND TEMPERATURESpeedNormally, if the jet velocity remains constant, independent of aircraft speed, then asthe aircraft speed increases, the thrust would decrease in direct proportion.However, due to ‘ram effect’ from aircraft forward speed, extra air is taken into theengine so that the mass airflow and also the jet velocity increase with aircraft speed.The effect of this tends to offset the extra intake momentum drag due to the forwardspeed so that some thrust is recovered as aircraft speed increases.Note that the ram effect begins to level out at around 700 knots but that the totalthrust will never achieve the same level, at subsonic velocities, as that developedwith the engine static. However, a recovery of thrust near to static levels isachievable at supersonic aircraft speeds, ie. Concorde. Allowing for the effects ofaltitude.ALTITUDEWith increasing altitude the ambient air pressure and temperature are reduced. Thisaffects the engine in two ways:4 Chapter 1 : Section 3 Issued May 2006
  30. 30. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1o The fall in pressure reduces the air density andhence the mass airflow into the engine for a givenengine speed. This causes the thrust or SHP tofall.o The fall in air temperature increases the airdensity, so that the mass airflow into the engine fora given engine speed is greater. This causes themass airflow to reduce at a lower rate and socompensates to some extent for the loss of thethrust due to atmospheric pressure.At altitudes above 36,000 ft and up to 65,500 ft however, the temperature remainsconstant, and the thrust or SHP is affected by pressure only.TEMPERATURE (CLIMATE)On a cold day the density of the air increases so that the mass of air entering thecompressor for a given engine speed is greater, hence the thrust or SHP is higher.On a hot day the reverse will occur.HUMIDITYBecause the density of water vapour is less than that of dry air, the mass flow of airentering the engine will be less in a humid climate than in a dry climate. Thedifference is fairly small and, although engine manufacturers will take humidity intoaccount when testing engines, engineers will only take into account altitude andtemperature when carrying out power/thrust checks on installed engines.Issued May 2006 Chapter 1 : Section 3 5
  31. 31. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018THRUSTDistribution of ThrustThe thrust developed within a gas turbine engine is, as previously stated, theacceleration of a mass airflow within the engine. The acceleration can be rearwardas well as forward. The total forward thrust is the gross thrust developed by theengine but the forward thrust less the rearward thrust is the net thrust developed.Since the rearward thrust can be over 60% of the total this means that, with otherlosses, there is less than 40% left over to drive the aeroplane forwards.The force of the thrust is felt against all the static and rotating parts of the engine.The thrust felt against the rotating parts has to be transferred to the static part andthen passed to the airframe to drive the aircraft forwards.Not all of the thrust is felt in the same place, it is spread variously along the engineso that some is developed at the compressor and some at the combustion section,and so on.6 Chapter 1 : Section 3 Issued May 2006
  32. 32. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1CLASSIFICATION OF PROPULSION UNITSBasic Construction and Operation of the Pure JetThe modern jet engine is basically cylindrical in shape as it is essentially a duct intowhich the necessary parts are fitted.The parts from front to rear are the:o Compressoro Combustion Systemo Turbine Assemblyo Exhaust SystemA shaft connects the turbine to the compressor and fuel burners are positioned in thecombustion system. Initial ignition is provided once the airflow is produced byrotation of the compressor, the pressure of the mass ensures the expanding gastravels in a rearward direction. Once ignition is achieved, the flame will becontinuous, providing fuel is supplied and the ignition device can be switched off.The hot gases crossing the turbine produce torque to drive the compressor,therefore the starter can also be switched off.Issued May 2006 Chapter 1 : Section 3 7
  33. 33. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018As the demand for more and more thrust increased in order to lift higher and higherpayloads, the size of the spool to drive the required mass airflow became prohibitive.The heavier spools were taking too long to accelerate (spool up) and therefore themass flow of air was also taking too long to reach the combustion section. In orderto make the spool lighter it was split into two parts. Furthermore, the gas could alsobe split so that only a portion of the air would now pass through the combustor, therest could go around the ‘core’ engine as bypass air and mix after the combustionprocess. This would give a greater mass of air travelling more slowly – to reducekinetic friction losses, and improve the pressure thrust at the nozzle by giving alarger area from which the cooler mass exits.A further advance was to make the spool even lighter by splitting it into three parts.Add a higher bypass ratio, the engine shown above is a ‘low bypass’ of around 1:1,to make the engine more efficient at lower altitudes and speeds and the turbofancomes into existence. These have ratios of over 1:1 to more than 6.5 and are said tobe ‘high bypass’ engines. Note that the British term is a ‘twin spool’ engine but theAmericans use the phrase ‘dual axial’. On some high bypass engines the ‘cold’ airand the ‘hot’ air emerge as separate streams; on others, the streams are combinedin what is called a ‘common nozzle’.The high bypass turbofan engine is essentially a fixed pitch, multi-bladed, ductedpropeller. It shifts a very large mass of air faster than a conventional propeller butmuch more slowly than a pure jet. It is able to move this large mass of air quitequickly because, unlike a propeller blade, the fan blade can cut through the air atsupersonic velocities – the tip is usually quoted as moving at around Mach 1.3.8 Chapter 1 : Section 3 Issued May 2006
  34. 34. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1It is for this reason that some turbofans have mid-span shrouds (snubbers orclappers) at some stage along their length to support them and stop them ‘whipping’in the airflow. Wide chord fan blades do not require these supports.As stated above, the air mass flow is cooler leaving a bypass engine so that thethrust is achieved by moving the air through a larger area of nozzle. By making acomparison between pure jet engines and bypass engines more differences can befound.Issued May 2006 Chapter 1 : Section 3 9
  35. 35. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018The turbines of pure jet engines are heavy because they are dealing with the wholemass flow of air through the engine. Bypass engines are only using a proportion ofthe air through the turbines which means, in turn, that the HP compressor, and thecombustor can now be made smaller and, therefore, lighter. In order to obtain thesame power at the turbine to drive the compressors and the accessories, the turbineinlet temperature is elevated and the pressure ratio is also increased. The coreengine is not only narrower but shorter and the use of modern materials andimproved gas flow characteristics makes for a considerably lighter engine. Theweight reduction on a typical low bypass engine over a pure jet of similar mass flowis around 20%.Curiously, the number of parts in a triple spool engine is less than those in a twinspool. This is brought about by having smaller overall spool sizes and permits acloser matching between the components. This, in turn, leads to less stages in boththe compressor and the turbine to perform the same tasks.For a given mass flow of air through the engine, a bypass engine produces lessthrust due to the lower exit velocity. To obtain the same thrust, a bypass enginemust be scaled to move a greater mass flow of air than a pure jet engine. Theweight of the engine is still less than the equivalent pure jet engine because of thereduced size of the HP section which still gives an improvement in the power/weightratio as well as a lower specific fuel consumption (this will be dealt with later).10 Chapter 1 : Section 3 Issued May 2006
  36. 36. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1BASIC CONSTRUCTION AND OPERATION OF THE TURBO-PROP ENGINEThe turbo-prop (turbo-propeller engine) is a combination of a gas turbine and apropeller. Turbo-props are basically similar to turbo-jet engines in that both have acompressor, combustion chamber(s), turbine and jet nozzle, all of which operate inthe same manner on both engines.However, the difference is that the turbine in the turbo-prop engine usually has morestages than that in the turbo-jet engine. In addition to operating the compressor andaccessories, the turbo-prop turbine transmits increased power forward, through ashaft and a gear train, to drive the propeller. The increased power is generated bythe exhaust gases passing additional stages of the turbine.The exhaust gases and reaction within the engine also contribute to engine poweroutput through jet reaction, the amount of energy available for jet thrust is roughly10% on most modern engines at ISA (SL).the turbopropeller pictured above is known as a ‘fixed turbine’ unit. This is becausethe turbine drives the compressor, accessory gearbox and reduction gearbox(propeller) as one mechanically coupled unit. This is a very simple system. It is lightfor the power output obtained and relatively simple to maintain.Most turbopropeller engines are now ‘free turbine’ units. This is a design wherethere is one turbine to drive the compressor and the accessory gearbox and anotherturbine to drive the reduction gear and propeller. The only link between the ‘core’engine (the turbine, compressor and accessory gearbox) and the propeller drive isenergy rich gas.Issued May 2006 Chapter 1 : Section 3 11
  37. 37. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018The turbopropeller engine shown above is a combination of the two because it has afree turbine that also drives a LP compressor and a HP compressor driven by its ownturbine. The accessories are driven from the gas generator (‘core’) – the HP section.The typical turbo-prop engine can be broken down into assemblies as follows:o The power section assembly which contains theusual major components of gas turbine engines(compressor, combustion chamber, turbine andexhaust system).o The reduction gear or gearbox assembly whichcontains those sections peculiar to turbo-propconfigurations.o The torque meter assembly which transmits thetorque from the engine to the gearbox to thereduction section.o The accessory drive housing assembly.The turbo-prop engine can be used in many different configurations. It is often usedin transport aircraft, but can be adapted for use in single-engined aircraft.12 Chapter 1 : Section 3 Issued May 2006
  38. 38. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 4 : COMPRESSORSINTRODUCTIONA gas turbine engine, to produce thrust, requires a mass airflow, so some type of airpump is needed. The compressor provides this mass flow and, at the same time,increases its pressure. Although the pressure has little bearing on the mass, it doesimprove the rate at which energy is released from the burning fuel. To a largeextent, high pressure ratios have a similar effect as high compression ratios in pistonengines.Basically two types of compressor are used in gas turbines, centrifugal vane typesand axial flow types.With regard to the advantages and disadvantages of the two types, the centrifugalcompressor is usually more robust than the axial flow types and also easier todevelop and manufacture. The axial flow type, however, consumes far more air thanthe centrifugal compressor of the same frontal area and can also be designed forhigh pressure ratios more easily. Since the airflow is an important factor indetermining the amount of thrust, this means that the axial compressor engine willalso give more thrust for the same frontal area.With higher pressure ratios, there is a higher engine efficiency and performance dueto an improved specific fuel consumption and thrust.Issued May 2006 Chapter 1 : Section 4 1
  39. 39. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018CENTRIFUGAL TYPEThis type is similar to the supercharger used in piston engines, consisting of a discon which is formed a number of radially spaced vanes. Around the disc, or impeller,is a ring of stationary vanes formed with divergent cross section between them.DOUBLE ENTRY CENTRIFUGAL COMPRESSORWhen driven at high speed, the air at the disc centre is forced radially outwardsalong the vanes of the disc. The rotational energy of the disc imparts velocity energyto the air, but, because the disc vanes have a divergent passage, some of theenergy is converted into pressure and temperature.2 Chapter 1 : Section 4 Issued May 2006
  40. 40. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1Leaving the impeller tip at high speed, the air enters the ‘diffuser’ ring and passesthrough its divergent passages that cause most of the remaining velocity energy tobe converted into pressure and temperature.Because of the high speed of rotation and drastic changes in the airflow direction,the temperature increase is high and this tends to lower this type of compressor’sefficiency. Furthermore, when impeller tip speeds reach sonic values, no furtherpressure increase is possible. The impeller tips will often reach speeds of 1600ft/sec. This limits the pressure ratio of this type of compressor to about 4.5:1.Issued May 2006 Chapter 1 : Section 4 3
  41. 41. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018There are single sided impeller and double-sided impellers. The double-sidedimpeller will give a greater air mass flow for an equivalent frontal area than a single-sided impeller but it has certain disadvantages that will be discussed elsewhere.SINGLE-SIDED AND DOUBLE-SIDED IMPELLERSIn spite of the adoption of the axial flow type compressor, some engines still retainthe centrifugal type because it is:o Simple and comparatively cheap to manufactureo Robust in construction and less vulnerable todamageIts main disadvantages are:o The high speed of rotation requiredo Large frontal areao Limited pressure ratioo High temperature increase4 Chapter 1 : Section 4 Issued May 2006
  42. 42. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1Issued May 2006 Chapter 1 : Section 4 5
  43. 43. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018MATERIALS AND CONSTRUCTIONThe impeller of a centrifugal compressor is of unit construction, being forged from analuminium alloy. The radial vanes are an integral part of the impeller, beingmachined from the forged blank. The rotating guide vanes are often made out ofsteel to resist impact damage and are splined onto the aluminium impeller. On otherimpellers the rotating guide vanes are aluminium and forged at the same time, andout of the same ‘block’, as the impeller disc and radial vanes.Some modern impellers are made out of titanium for lightness and impact resistance.These are constructed in one piece so that the impeller disc, vanes and rotatingguide vanes are integral.6 Chapter 1 : Section 4 Issued May 2006
  44. 44. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1AXIAL FLOW COMPRESSORIn the axial flow compressor, many stages of moving and stationary blades areneeded. These are positioned alternately so that each row of rotating (rotor) bladesis followed by a row of stationary (stator) blades. A row of rotors and a row of statorsform a stage.Because the air leaving each stage is at a higher pressure, it occupies a smallerspace. Therefore, each stage of the compressor is smaller than the preceding one,giving the casing a convergent passage. This maintains uniform axial velocity.Both the rotors and stators are of aerofoil section, between each adjacent rotor bladeand each adjacent stator blade the cross sectional area is ‘divergent’. Duringrotation the rotors act similarly to a propeller blade and accelerate the air rearwards,velocity energy is converted into pressure and temperature. Leaving the rotor, theair passes across the stator, the divergence here causes the remaining velocityenergy to be converted into pressure and temperature.The stators are angled so as to pass the air into the next stage of rotor at the correctangle of attack. Each stage will increase the pressure by about 1.1 – 1.2:1 and thetemperature by approximately 25°C. In order to increase the pressure more at eachstage there would need to be a greater divergence; this would risk the airflowbreaking away from the blades. If this occurs, the blades are no longer doing workon the air and are considered to be in a state of ‘stall’.Issued May 2006 Chapter 1 : Section 4 7
  45. 45. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018COMPRESSOR STALL AND SURGEAs with centrifugal flow compressors, the compressor acts as a pump to drive airthrough the engine and to compress it before energy is added. It is important toremember that the working fluid of the engine is air. It is the mass flow of air throughthe engine, and the acceleration of that mass flow, that gives thrust or power. Tothis end it must also be remembered that any amount of air entering the front of theengine must come out of the back. This may seem obvious but is the root of thedevelopment of thrust by a jet engine. If an engine takes in a mass flow (thewords ‘mass’ and ‘flow’ are vital here) of, say, 10lb of air per second at the front, thenthere must be 10lb per second of air emerging from the back. However, the 10lbcoming out at the back is going to be of a much greater volume than the 10lb goingin the front. This means that the velocity is much greater. It is the job of thecompressor to organise the mass flow into manageable units for the engine tohandle. If the rotation of the engine (compressor) does not match closely the massflow of air passing through it, then there will be problems that could, potentially,damage the engine.This situation is much more critical with axial flow compressors than it is withcentrifugal compressor. The relationship between the rpm of the compressor andthe mass flow can be seen clearly on a graph that is called the WORKING LINE ofthe engine.As has already been stated, the mass of air passing through the engine is affectedby altitude, temperature and forward speed. This means that the mass flow is not aconstant for any given rpm – hence the rpm line crossing the working line will be acurve to demonstrate this.Because the rotors and stators are of aerofoil shape, the airflow reacts in a similarway to the airflow over a wing. Because of this, the air must pass over the rotorsand stators at the correct angle. If not, they will cause the airflow to becometurbulent and stalling takes place. The angle at which the airflow strikes the bladesis dependent upon the rotational speed and the rate of linear flow, therefore, an axialflow compressor can only be designed to have these correct flows at one particularRPM and mass flow.Below this RPM or above it, the angles are incorrect. The further from the designedRPM, the more incorrect are angles and consequently, the more turbulent theairflow. These compressors therefore, are designed to be most efficient at theengine’s maximum cruise RPM Limitation of the maximum engine RPM to a little8 Chapter 1 : Section 4 Issued May 2006
  46. 46. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1above this, prevents stall at the higher RPM range, but during acceleration up tocruise, stalling is always possible and therefore, some control of airflow below thedesigned RPM is usually necessary.If stalling becomes excessive, the mass flow leaving the compressor is greatlyreduced and as approximately 60% of this air is used to keep combustion chambertemperature within limits, the temperature rises and can cause serious damage tothe engine. If combustion pressure also increases due to excessive temperature,above compressor outlet pressure, the airflow will reverse in direction and surgeforward through the compressor, with possible risk of damage to the engine.A compressor stall can be recognised by the following:o Vibrationo Rumbling noiseo Inability of the engine to accelerateo Rapid rise in exhaust gas temperatureOnce all the stages in the compressor have stalled they will be doing no work on theair. At this point the air at high pressure at the back of the compressor will travel tothe low pressure zone at the front of the engine. It will do this very rapidly –explosively, almost. So, when all stages of an axial flow compressor have stalled, asurge will develop which will be recognised by a loud bang, or banging, in the intake.Issued May 2006 Chapter 1 : Section 4 9
  47. 47. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018AIRFLOW CONTROLTo ensure efficient operation of a gas turbine engine over a large rpm range and toretain the safety margin between normal operating conditions and those conditionscausing compressor stall and engine surge, a system of airflow control is used. Thissystem usually consists of a row of intake guide vanes arranged so that their angel isautomatically adjusted by a control that is sensitive to engine speed, to prevent orminimise compressor stall in the first intake stage.To further improve the smooth airflow through the compressor, valves are fitted tobleed away air from selected intermediate stages of the compressor – the air bled offpasses to atmosphere. In addition to bleed valves and variable angle intake guidevanes, some recent engine designs include variable angle stator blades. Variableangle stator blades are automatic in action and are fitted to the compressor stagesmost likely to stall.The variable angle intake guide vanes give ‘whirl’ to the air entering the front of thecompressor and the amount is adjusted to suit engine running conditions. The angleof the intake guide vanes is altered mechanically by an actuator and suitable linkage.The actuator control is sensitive to engine rpm and intake temperature.10 Chapter 1 : Section 4 Issued May 2006
  48. 48. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1The amount of air bled from the compressor is controlled by bleed valves that areopen at idle rpm, and low pressure ratios, and automatically closed at higher rpmwhen the airflow conditions are more stable. This is because the conditions mostlikely to bring about compressor stall and engine surge are those encountered whenthe engine is idling in flight or during MAXIMUM ACCELERATION. Compressorstall may also be induced by ice formation in the air intake and by certain aircraftmanoeuvres. The air bleed valves do most to reduce the risk of compressor stall as,during the critical periods, these valves are open to dump the excess airflow from theearly stages to reduce the airflow over the rear stages. This prevents ‘choking’ ofthe rear stages with subsequent stall at the front leading to engine surge.Issued May 2006 Chapter 1 : Section 4 11
  49. 49. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018TYPES OF AXIAL FLOW COMPRESSORThe most simple type is the single spool compressor which consists of one rotorassembly and stators with as many stages as necessary to achieve the desiredpressure ratio and all the airflow from the intake passes through the compressor.Because of the large number of stages required to produce a high compression ratio,as the number of stages increases so it becomes more difficult to ensure that eachstage will operate efficiently over the engine speed range.To make the engine more efficient and achieve a higher pressure ratio multi-spoolcompressors, consisting of two, or more rotor assemblies, each driven by their ownturbine at optimum speed, were introduced.In a further development, the air from the LP compressor (the first spool) is split intotwo parts. Only a percentage of the air from the LP compressor passes into the HP12 Chapter 1 : Section 4 Issued May 2006
  50. 50. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1compressor, the remainder of the air, the bypass flow, is ducted round the HPcompressor. Both flows mix in the exhaust system before passing to the propellingnozzle.Afan may be fitted to the front of a single or twin spool compressor and, on thesetypes of engine, the fan is driven at the same speed as the compressor to which it isfitted. Very often, the LP compressor to which a fan is fitted may be called a‘booster’ rather than a LP compressor.Issued May 2006 Chapter 1 : Section 4 13
  51. 51. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018On engines of the triple spool type the fan is the LP compressor and is driven by itsown turbine separately from the intermediate pressure compressor and the HPcompressor.The LP compressor has large rotor (fan) blades and stator blades and is designed tohandle a far larger mass airflow than the other two compressors, each of which hasseveral stages of rotor blades. A large proportion of the air, from the outer part ofthe fan and known as the ‘cold’ stream, by-passes the other two compressors and isducted to atmosphere through the cold stream nozzle. The smaller air flow, from theinner part of the fan and known as the ‘hot’ stream, passes through the intermediateand HP compressors where it is further compressed before passing to thecombustion system.14 Chapter 1 : Section 4 Issued May 2006
  52. 52. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1MATERIALS AND CONSTRUCTIONThe construction of the axial flow compressor centres around the rotor assembly andcasings. The casing assembly consists of a number of cylindrical casings, some ofwhich are in two halves to facilitate engine assembly and maintenance. When thehousings are bolted together, they completely enclose the rotor. It is usually made ofmagnesium or aluminium alloy at the front and steel in the rear.The rotor assembly consists of a drum or series of discs mounted in tandem on acentral shaft. Fastened to the rotor are rows of blades made of light alloy or steel.The blades vary in length, the longest at the front, usually made of a light alloy andthe shortest at the rear, usually made of steel.Issued May 2006 Chapter 1 : Section 4 15
  53. 53. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018Methods of securing the blades to the rotor vary with different designs shown in thediagrams. The blades are sometimes free to rock slightly so that much of the stressconcentration near the blade root is relieved.The rotor blades are of aerofoil section and are usually designed to give a pressuregradient along their length to ensure that the air maintains a fairly uniform axialvelocity. The higher pressure at the tip balances out the centrifugal action of therotor on the airstream. To obtain these conditions it is necessary to twist the bladefrom root to tip to give the correct angle of incidence at each point.The stator blades are again of aerofoil section and are secured into the compressorcasing or into stator blade retaining rings, which are themselves secured into the16 Chapter 1 : Section 4 Issued May 2006
  54. 54. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1casing. The blades are often mounted in packs in the front stages and may beshrouded at their tips to minimise the vibrational effect of flow variations on thelonger blades. It is also necessary to lock the stator blades in such a manner thatthey will not rotate around the casing.FAN BLADESThe high by-pass ratio fan can produce 75% of engine thrust, have a diameter of 97inches and be capable of pumping air at the rate of 1670 lbs/sec with tips speeds ofup to 1500 ft/sec. The relative speed of air at the tips is 1600 ft/sec. A blade thatweighs 15 lbs imposes a load of 60 tons on its root and the disc that supports it. Theforces generated by the blade whilst rotating means that the blade must be as lightas possible to reduce the forces to an acceptable level and keep out of balanceforces to a minimum in the event of blade failure. It is a requirement that if a bladebecomes detached it must not penetrate the engine casing and hazard the aircraft soa containment ring is necessary around the fan and the heavier the blade - theheavier and stronger the ring needed. Furthermore, there is an aerodynamic elementin that the containment ring is shaped to ‘flip’ the blade so that it will be swept downthe fan duct before the next fan blade can hit it.TYPICAL FAN BLADESIssued May 2006 Chapter 1 : Section 4 17
  55. 55. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018These factors imposed a weight limitation on the blade which meant that even usingsolid titanium the blade had to be a relatively narrow chord to keep the weight down.The narrow chord necessitated the use of mid span shrouds or snubbers to preventblade flutter due to aerodynamic instability. The snubbers, also known as clappers,form a virtually solid ring around the fan when rotating to act as a stiffener, but theyimpose a penalty by causing a pressure loss and reduced flow area, and obviouslythis reduces engine efficiency.Developments by Rolls Royce (after 20 years of work) have produced a wide chordlightweight blade which because of its wide chord does not need snubbers. Thisimproves the efficiency and, of course, fuel consumption. The blade consists of ahoneycomb core of titanium enclosed in a titanium skin. This construction producesa strong lightweight blade, allowing a wider chord and eliminating the need forsnubbers.18 Chapter 1 : Section 4 Issued May 2006
  56. 56. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 5 : TYPES OF INTAKEThe main requirement of an air intake is that under all operating conditions, deliveryof the air to the engine is achieved with the minimum loss of energy occurringthrough the duct. To enable the compressor to operate satisfactorily, the air mustreach the compressor at a uniform pressure distributed evenly across the whole inletarea.The ideal air intake for a turbo jet engine fitted to an aircraft flying at subsonic or lowsupersonic speed is a short, pitot-type circular intake. This type of intake makes thefullest use of the ram effect on the air due to forward speed, and suffers theminimum loss of ram pressure with changes of aircraft attitude. However, as sonicspeed is approached the efficiency of this type of intake begins to fall because of theformation of shock waves at the intake lip.The pitot-type intake is sometimes used on single engined aircraft, but in this casethe intake is divided at the wing roots and is not circular.Issued May 2006 Chapter 1 : Section 5 1
  57. 57. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018At higher supersonic speeds the air flow velocities met are much higher than thecompressor can efficiently use, therefore the intake air velocity must be decreased.In this case it is necessary to have an intake that has a variable throat area and spillvalves to accommodate and control the changing volume of air.The engine intake is the biggest single source of drag on the aircraft and sodesigners have to be very careful when deciding where to install engines. Enginesthat are far from the centre line of the aircraft will impose considerable turningmoments on the aircraft when under power or when shut down in flight. Having theengines close to the fuselage is more practicable but gives rise to the possibility ofturbulent air entering the intake from the skin of the aircraft and can create thepossibility of ‘handling stalls’ where the attitude of the aircraft can ‘blank off’ an2 Chapter 1 : Section 5 Issued May 2006
  58. 58. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1intake. Clearly, the best location is at the nose of the aircraft but, for a four-enginedcommercial transport aircraft, this is entirely impracticable!NOTES:Issued May 2006 Chapter 1 : Section 5 3
  59. 59. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 6 : COMBUSTION SYSTEMSINTRODUCTIONThe combustion system is designed to burn the fuel as efficiently as possible overthe whole range of engine operating condition. It must do so without any increase inpressure, all the energy released by the fuel is converted into heat and velocityenergy. Very high temperatures exist in the combustion system, the burningtemperature of the fuel being in the region of 2,000°C. To protect the material fromwhich the system is manufactured, about 60% of the air flow is used for cooling.FUEL/AIR RATIOSCombustion must be completed rapidly and in a comparatively small part of thecombustion chamber; therefore the mixing of the air and fuel must be good. Thecombustion process requires 15 lbs of air to 1 lb of fuel for COMPLETE combustion.This is known as an air/fuel ratio of 15:1 by weight.As a large proportion of the air is used for cooling this means that a much greaterproportion of air is passing through the combustion system that is required forcombustion alone. Thus, the ratio of air to fuel can be between 45:1 and 130:1 overthe engine RPM range.AIR FLOWThe air flow leaving the compressor is first split into two, approximately 25% and40% being used for combustion, the other 60% - 75% is further divided, the greaterproportion for gas cooling.These three flows are known as:-Issued May 2006 Chapter 1 : Section 6 1
  60. 60. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018o Primary air flow, for mixing with the fuel and tosupport combustion.o Secondary air flow to shape the flame andcomplete combustion.o Tertiary air flow, to cut off the flame and reducegas temperature to a figure acceptable to theturbine.Secondary and tertiary air also forms a boundary flow on the inside and outside ofthe flame tube, which is manufactured from special heat resistant steel (nimonic), theair casing is normally aluminised mild steel.TYPE OF SYSTEMThere are basically three types of combustion systems:-(a) Multiple chamber(b) Tubo-annular(c) AnnularType (a) has a number of interconnected chambers in a circle around the spine ofthe engine. Except for fuel drains and ignitors, each chamber is identical on anyparticularly mark of engineType (b) the tubo-annular, is the ‘half-way house’ in design between the separatechambers of the multiple and single chamber of the annular type. It uses an annularair casing around the engine spine and individual interconnected flame tubes, fittedwithin the casing. It is also known as a ‘can-annular’ system.Type (c) the annular combustion system, is a single chamber surrounding theengine. Annular inner and outer air casings form a tunnel around the spine of theengine, in the space between the inner and outer casings is fitted an inner and outerflame tube.2 Chapter 1 : Section 6 Issued May 2006
  61. 61. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1Issued May 2006 Chapter 1 : Section 6 3
  62. 62. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018MULTIPLE CHAMBERTUBO-ANNULAR4 Chapter 1 : Section 6 Issued May 2006
  63. 63. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1ANNULARThe three designs can be summed up simply by saying that the multiple type wasused on earlier engines and, although somewhat bulky, it was simple to dismantleand service. The tubo-annular has some of the advantages of the multiple but ismore compact – less frontal area, with a smooth exterior and reduced weight, it wasused on later engines. The annular system, as used on the latest engines, providesa much more compact system and, for the same power output and mass flow, amuch shorter one. Annular systems also give a better flow into the turbine byallowing more burners for the fuel flow required; multiple types were, generally,limited by having one burner per can. Tubo-annular combustion chambers weredeveloped to have multiple burners per chamber but this still did not give a smoothand even flow into the turbines(s).Issued May 2006 Chapter 1 : Section 6 5
  64. 64. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/0186 Chapter 1 : Section 6 Issued May 2006
  65. 65. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1The cooling flow of air into the combustion liner is through several different types ofboundary layer cooling chutes. This is a double-skinned combustion liner with tinyholes in each skin. The air is drawn from the air casing into the combustor throughthese tiny holes which induce a cool layer of air over the metal of the liner.The entrained air flows to support combustion and shape the flame are throughlarger holes carefully designed with regard to size and placement.This is to ensure that:o The flame is not distortedo The pressure drop across the liner is maintained ata satisfactory level at all operating ranges of theengine.COMBUSTION EFFICIENCYThe combustion efficiency of most gas turbine engines at sea level conditions is100%.CARBON FORMATIONHigh pressure ratio engines tend to produce exhaust smoke at maximum powerconditions. This indicates that carbon particles are being formed in over-rich regionsof the primary zone in conditions of less turbulence at high temperature andpressure. However, smoke represents an almost negligible loss in combustionefficiency of less than 0.3%.Issued May 2006 Chapter 1 : Section 6 7
  66. 66. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018BURNERSAtomisingThis type atomises the fuel in preparation for combustion and its efficiency isproportional to gas velocity. Therefore, it is very efficient at high engine rpm butrelatively inefficient at slow rpm. This requires that a high pressure fuel system mustbe used if starting problems are to be avoided, but, in general, it is simpler and lesscomplex than the vaporiser type in starting the engine.Vaporising TypeThis type vaporised the fuel by pre-heating it prior to combustion. This raisesproblems when starting the engine in that some form of artificial heat must beapplied.8 Chapter 1 : Section 6 Issued May 2006
  67. 67. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1GENERALThe burner type is largely a manufacturer’s choice and many factors are taken intoaccount before a decision is taken as to which type is finally used on a particularengine.Burners will be covered in greater detail in a later phase.Issued May 2006 Chapter 1 : Section 6 9
  68. 68. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018NOTES:10 Chapter 1 : Section 6 Issued May 2006
  69. 69. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 7 : TURBINESTwo basic types of turbine are used to produce torque, the impulse type, whichutilises the velocity energy of a moving stream, ie. a water wheel and the reactiontype using the pressure energy in a stream, ie. high pressure stream turbines.Those used in gas turbine engines combine both and are referred to asimpulse/reaction turbines.Before each turbine wheel is placed a set of nozzle guide vanes (NGVs). These areused to accelerate the gas stream to as high a value as possible and to direct thegas at the most efficient angle of attack. Therefore, there is a convergent crosssection passage between each vane.TURBINE CONSTRUCTIONA turbine consists of a disc on which is mounted a number of blades, between which,at some point, a convergent cross section area is formed. Torque is produced eitherby the gases impinging on the blade and/or accelerating between the blades.The impulse/reaction turbine will extract energy from the heat, velocity and pressureenergies of the gas stream. As heat energy is the most important, the higher theinlet temperature to the turbine, the greater will be its efficiency; this efficiency islimited by the materials available.Issued May 2006 Chapter 1 : Section 7 1
  70. 70. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018The blades of an impulse/reaction turbine are of aerofoil shape with a ‘cup shape’ ateither the forward part of the blade or at the root. This cup shape is the impulsesection and, in short blades, is formed at the forward section of the blade. With longblades, the impulse section is formed at the root or base of the blade. The cup orimpulse section uses the velocity energy. In both cases heat energy is used up.2 Chapter 1 : Section 7 Issued May 2006
  71. 71. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1TURBINE ASSEMBLIESThe simplest form of assembly consists of NGVs followed by a turbine wheel. Thisis known as a ‘single stage’ turbine assembly. An assembly of this type, to producemore torque, must have the inlet temperature increased, its diameter enlarged or itsr.p.m. increased. In each case a limit is imposed by the materials. To overcome thismaterial limitation, the modern turbines are ‘multi-staged’ ie. wheels of smallerdiameter are used in tandem, allowing increased r.p.m. without increased centrifugalloads, which also allows the temperature to be increased. Up to three stages arenormally used.THREE STAGE SINGLE SPOOL TURBINEThe first stage is known as the H.P. stage, the second stage is the intermediatestage and the third is referred to as the L.P. stage. Where only two stages are used,they are known as the H.P. and L.P. stages.Issued May 2006 Chapter 1 : Section 7 3
  72. 72. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018A further advance in temperature increase has been achieved by blade cooling. Apercentage of the mass flow is passed through holes formed in the blades, whichreduced the blade surface temperature. This allows the inlet temperature to beincreased without affecting the material.4 Chapter 1 : Section 7 Issued May 2006
  73. 73. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1A one, two or three stage assembly can be used as a drive for the different sectionsof the engine, the compressors, the fan or the power ‘take-off’ to drive propellers,rotors, etc., so some modern engines have three turbine assemblies, with possibly atotal of six or seven turbines. Each of these assemblies is identified by its position inrelation to the combustion system outlet, ie. first the H.P. assembly, second theintermediate assembly and finally the L.P. assembly.TWIN SPOOL TURBINEIssued May 2006 Chapter 1 : Section 7 5
  74. 74. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018TURBINE EFFICIENCYThe losses which prevent the turbine from being 100% efficient are due to a numberof reasons. A typical three stage turbine would suffer a 3.5% loss because ofaerodynamic losses in the turbine blades. A further 4.5% loss would be incurred byaerodynamic losses in the NGVs, gas leakage over the rotor blades and exhaustsystem losses. These losses are of approximately equal proportions. If the totallosses are deducted, an overall efficiency of 92% is the result.COMPRESSOR - TURBINE MATCHINGThe flow characteristics of the turbine must be very carefully matched with those ofthe compressor to obtain the maximum efficiency and performance of the engine. If,for example, the nozzle guide vanes allowed too low a maximum flow, then a backpressure would build up causing the compressor to surge, too high a flow wouldcause the compressor to choke. In either condition, a loss of efficiency would veryrapidly occur.THE TURBINE DISCThe turbine disc is machined forging with a flange onto which the shaft may bebolted. The disc also has around its perimeter provision for the attachment of theturbine blades. The disc is manufactured from a high heat resistant alloy steel,usually from the ‘nimonic’ range.TURBINE BLADESThe turbine blades are of an aerofoil shape but the profiles do not follow anyparticular class of aerofoils. The actual area of each blade cross section is fixed bythe permitted stress in the material used and by the size of any holes, which may berequired for cooling purposes. High efficiency demands thin trailing edges to thesections but a compromise has to be made so as to prevent the blades cracking dueto temperature changes during engine starting and stopping. To cope with the highthermal stresses during operation, the blades are made of ‘nimonic alloys’.The method of attaching the blades to the turbine disc is of considerable importance,since the stress in the disc around the fixing or in the blade root has an importantbearing on the limiting rim speed. The majority of gas turbine engines use the ‘firtree’ fixing. This type of fixing involves very accurate machining to ensure that theloading is shared by all the serrations. The blade is free in the serrations when theturbine is stationary and is stiffened in the root by centrifugal loading when theturbine is rotating.MATERIALS AND STRESSESAmong the obstacles in the way of using higher turbine entry temperature havealways been the effects of these temperatures on the nozzle guide vanes andturbine blades and the high speed of rotation which impart tensile stress to theturbine disc and blades.The highly stressed turbine blades therefore make it necessary to restrict the turbineentry temperature so that they and the nozzle guide vanes may do their job for asatisfactory length of working life without reaching the end of their useful creep life.Turbine discs also operate for a satisfactory length of working life without reaching6 Chapter 1 : Section 7 Issued May 2006
  75. 75. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1the end of the useful creep life of the material. Any increase in turbine entrytemperature must be accompanied by appropriate increases in material thicknessand an increase of the cooling air flow.The blades, while glowing red hot, must be strong enough to carry the centrifugalloads due to rotation at high speed. A blade weighing only two ounces may exert aload over two tons at top speed and it must withstand the high bending loads appliedby the gas to produce the many thousands of turbine horse-power necessary to drivethe compressor. Blades must also be resistant to fatigue and thermal shock, so thatthey will not fail under the influence of high frequency fluctuations in the gasconditions and they must also be resistant to corrosion and oxidisation.All the rotating turbine blades are subject to creep. This is the growth of the bladeover its operational life.The stresses on the blade are:• aerodynamic• centrifugal• thermal.Centrifugal force is the greatest by far of these forces. Every time a blade is used it isalways longer when it has cooled than when it was at the same temperature beforeuse. This growth is ‘creep’. The initial creep is called primary creep; most of its life iswhere growth occurs very gradually, this is ‘secondary creep’; and the final growthrate, much accelerated prior to final fracture (catastrophic failure) is ‘tertiary creep’.Issued May 2006 Chapter 1 : Section 7 7
  76. 76. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 8 : EXHAUST SYSTEMAero gas turbine engines have an exhaust system that passes the turbine dischargegases to atmosphere at a velocity, and in the required direction, to provide theresultant thrust. The velocity and pressure of the exhaust gases create the thrust inthe turbo-jet engine, but in the turbo-prop engine only a small amount of thrust iscontributed by the exhaust gases, because most of the energy has been absorbedby the turbine for driving the propeller.The design of the exhaust system, therefore, exerts a considerable influence on theperformance of the engine. The areas of the jet pipe and propelling or outlet nozzleaffect the turbine entry temperature, the mass air flow and the velocity and pressureof the exhaust jet.The temperature of the gas entering the exhaust system is between 550° and 850°according to the type of engine. Turbo propeller and by-pass engines have thecoolest flow. With the use of after-burning (reheat) the temperature in the jet pipecan be 1700°C or higher.Issued May 2006 Chapter 1 : Section 8 1
  77. 77. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018EXHAUST GAS FLOWGas from the turbine enters the exhaust system at velocities from 750 - 1200ft/sec,but, because of high friction losses the speed of flow is decreased by diffusion. Thisis accomplished by having an increasing passage area between the exhaust coreand outer wall. This core also prevents the hot gases from flowing across the rearface of the turbine disc. It is usual to hold the outlet gas velocity at approximately950ft/sec. Additional losses occur due to residual whirl velocity in the gas streamfrom the turbine. To reduce these losses the core support fairings are designed tostraighten out the gas flow.The exhaust gases pass to atmosphere through the propelling nozzle, which forms aconvergent duct, thus increasing gas velocity. The exit velocity of the exhaust gasesis subsonic at low thrust conditions only. During most operating conditions, the exitvelocity reaches the speed of sound in relation to the exhaust gas temperature, andthe propelling nozzle is then said to be choked; ie. no further increase in velocity canbe obtained unless the temperature is increased. As the upstream total pressure isincreased above the value at which the propelling nozzle becomes ‘choked’ thestatic pressure of the gases at exit increases above atmospheric pressure. Thispressure difference across the propelling nozzle gives what is known as pressurethrust, and is effective over the nozzle exit area.The propelling nozzle size is extremely important and must be designed to obtain thecorrect balance of pressure, temperature and thrust. With a small nozzle thesevalues increase, but there is a possibility of the engine surging, whereas with a largenozzle the values obtained are too low.2 Chapter 1 : Section 8 Issued May 2006
  78. 78. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1VARIABLE AREA NOZZLESOn some engines a variable area propelling nozzle is used. When this type ofnozzle is used, an increase in the flow area through the nozzle enables easierstarting to be made at low r.p.m. and temperature because of the reduction in turbineback pressure; with a reduced area, the thrust is increased. The variation in nozzlearea also enables low specific fuel consumption to be attained during some part ofthe engine operating range.Issued May 2006 Chapter 1 : Section 8 3
  79. 79. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018OTHER DESIGNSThe by-pass engine has two gas streams to eject to atmosphere, the cool by-passair flow and the turbine discharge gases. The two flows are combined by a mixerunit which allows the by-pass air to flow into the turbine exhaust gas flow in amanner that ensures thorough mixing of the two streams.On front fan engines the hot gas and cold air streams are exhausted separately.The hot and cold nozzles are co-axial and the area of each nozzle is designed toobtain maximum efficiency.4 Chapter 1 : Section 8 Issued May 2006
  80. 80. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1MATERIALSThe materials used in the manufacture of jet pipes and exhaust systems are similarto those used in combustion chambers ie. nimonic with mild steel or alloy outercasings.As a precaution against fire, and to provide some degree of cooling for the jet pipe,lagging of heat resisting material, or a cooling air flow, is provided.Issued May 2006 Chapter 1 : Section 8 5
  81. 81. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018THRUST REVERSALThe high landing speeds and landing weights of modern aircraft places tremendousloads on braking systems. This, in turn, causes rapid brake and tyre wear with theincreased risk of brake failure, fade and tyre burst through overheating. To reducethese possibilities and to lessen the loads on brakes and tyres another means ofslowing the aircraft was needed. Some military aircraft use braking parachutes butthis is not practicable on civil aircraft. The most common method is thrust reversal,which diverts all or a proportion of the exhaust jet forward, so reversing the thrust.REQUIREMENTSThe system must:-• Not affect engine operation, either when in use or not.• Be able to withstand high temperatures.• Be mechanically strong but light in weight.• Be ‘fail safe’.6 Chapter 1 : Section 8 Issued May 2006
  82. 82. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1Issued May 2006 Chapter 1 : Section 8 7
  83. 83. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018METHODSCLAMSHELL DOOR SYSTEMThis system is pneumatically operated. On selection of reverse thrust, the doorsrotate to uncover the ducts and close the normal gas stream exit. Cascade vanesthen direct the gas stream in a forward direction, so that the jet thrust opposes theaircraft motion. Ideally the gas should be directed in a completely forward position.It is not possible to achieve this, however, mainly for aerodynamic reasons, and adischarge angle of 45° is chosen. Reverse thrust power is approximately half that ofmaximum engine power in forward thrust.RETRACTABLE EJECTOR SYSTEMThis system is both pneumatically and hydraulically operated and uses bucket typedoors to reverse the exhaust gas stream. The ejector is mounted on a track thatextends rearwards from the combustion system to the propelling nozzle. Onselection of reverse thrust, hydraulic pressure moves the ejector rearwards over thepropelling nozzle. When the ejector is extended and latched the buckets are rotatedby a pneumatic actuator into the gas stream to deflect the stream in a forwarddirection.COLD STREAM REVERSER/HOT STREAM SPOILER SYSTEMOn a front fan engine, it is necessary to reverse the cold stream air flow, andbecause this air flow provides sufficient reverse thrust power, it is only necessary to‘spoil’ the hot stream gas flow to prevent it cancelling out the effect of the reversethrust obtained from the cold stream air flow.This system is operated by an air motor, the output of which is converted tomechanical movement through a series of flexible drives, gearboxes and screwjacks.On selection of reverse thrust, the actuation system moves the movable cowlrearwards and at the same time fold the blocker doors to blank off the cold streamfinal nozzle, thus diverting the air flow through the cascade vanes. Simultaneously,the spoiler doors move rearwards and swing across the hot stream to spoil the flow.TURBO-PROPELLER REVERSE PITCHThis system is operated by oil pressure. Movement of the throttle or power controllever directs oil from the control system to the propeller mechanism to reduce theblade angle to zero, and then through to negative (reverse) pitch.8 Chapter 1 : Section 8 Issued May 2006
  84. 84. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 9 : MATERIALS USED IN ENGINECONSTRUCTIONINTAKE AND COMPRESSOR SECTION CASINGSMagnesium or aluminium alloy at front with steel or nickel based alloys at the rearbecause of higher temperatures, with titanium coming into more common use.CENTRIFUGAL IMPELLERSAluminium forgings with some rotating guide vanes made of steel. Later types ofimpeller can be made of titanium.ROTOR, DRUMS, DISCSEarlier engines used aluminium alloy and steel but titanium is preferred for lightnessand strength. Titanium used for blades and discs is of higher heat resistance alloy.Rotor has to be as light as possible because of centrifugal forces.ROTOR BLADESAluminium alloy, steel and titanium. As higher temperature titanium alloys aredeveloped they are displacing nickel alloys used in high temperature areas.STATOR BLADESSteel or nickel based alloys having a high fatigue strength to resist affect of FOD.Earlier engines used aluminium alloy, steel and titanium.NOTE: Titanium can ignite due to frictionHIGH BY-PASS RATIO FAN BLADESTitanium is used to produce a light solid blade but requires a mid-span support(‘snubber’ or ‘clapper’). Later blades are made of skins of titanium with ahoneycomb core which enables a wider chord to be used with low weight andavoiding the need for ‘snubbers’.Issued May 2006 Chapter 1 : Section 9 1
  85. 85. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018COMBUSTION CHAMBERSFlame tubes of high heat resistant alloy with ‘nimonic’ a nickel-chromed alloy used agreat deal. The flame tubes may also have a heat resistant coating.BLADES Nickel based alloys used, though some small blades can bemade up of ceramicDISCS Nickel base alloys in current use or a more expensive productis a powder metallurgy disc which gives greater strength.NGV Nickel alloys with, possibly, ceramic coatingTURBINES Must resist creep, corrosion and oxidisation.EXHAUST SYSTEMNickel or titanium to resist high temperature. Insulation blanket consisting of innerlayer of fibrous material with a thin outer layer of dimpled (for strength) stainlesssteel. After-burning jet pipe made of heat resistant nickel alloy.GEARS AND GEARBOXESInternal gearbox casing – aluminium alloy. External can be magnesium alloy. Gearsmade of corrosion resistant steel, case hardened to resist wear.THE SHAFT AND BEARINGSThe shaft is transmitting a great deal of torque from the turbine into the compressorand accessories plus, in some cases, to the propeller shaft or the shaft to therotor-head. It is made of a chrome-vanadium-molybdenum steel alloy on olderengines, but modern engines are inclined to use marage steel. Maraging is aprocess where the carbon is removed from the steel and other elements are insertedin its place to restore the qualities required. This makes for a very light but strongalloy which is ideal for an engine shaft. Marage steel is, then, a carbon free steel –containing a maximum of 0.03% carbon.Along the shaft at various intervals will be bearing assemblies, roller bearings whichtake loads radiating from the centre of the shaft – used because of their greatersurface contact area, and at least one deep-grooved ball bearing which will transmitthrust loads to the carcass of the engine and then, through the engine mounts, to theairframe. These bearings comprise an inner race, an outer race and several rollingelements kept apart by a cage made of bronze. The bronze cage may be silver linedto reduce wear but is softer than the case-hardened high carbon steel that the rollingelements are made of. They are splash lubricated with oil and vented to the engineoil system, sealing the oil in is the responsibility of labyrinth seals backed with LP airfrom the compressor. Cooling is by the same LP airflow that is used to seal them.A mechanical take-off to the gearbox is by means of a radial drive shaft at an internalwheelcase. The radial drive shaft has a bevel gear at each end in order to turn thedrive through an angle from the axis of the main shaft – in this case an angle of 90°.2 Chapter 1 : Section 9 Issued May 2006
  86. 86. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 10 : BEARINGSINTRODUCTIONA bearing is any surface that supports or is supported by another surface. Bearingsare designed to produce a minimum of friction and a maximum of wear resistance.Bearings must reduce the friction of moving parts and also take thrust loads or acombination of thrust and radial loads. Those which are designed primarily for thrustloads are called thrust bearings. The two different types of bearings used on gasturbines are ball and roller.BALL BEARINGSA ball bearing consists of an inner race, and outer race and one or more sets of ballsand bearings which are designed for dismantling, a ball retainer or cage. Thepurpose of the retainer or cage is to prevent the balls touching one another. Ballbearings are used for radial and thrust loads, a ball bearing specially designed forthrust loads would have very deep grooves in the races.ROLLER BEARINGSThe bearings are manufactured in various shapes and sizes, and can be adapted toboth radial and thrust loads.The bearing race is a guide or channel along which the rollers travel, the roller issituated between an inner and outer race, both of which are made of case hardenedsteel. When the roller is tapered it rolls on a cone shaped race inside an outer race.Straight roller bearings are used only for radial loads and taper roller bearings willsupport both radial and thrust loads. Roller bearings will withstand greater radialloads than ball bearings because of a greater contact area.Issued May 2006 Chapter 1 : Section 10 1
  87. 87. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/0182 Chapter 1 : Section 10 Issued May 2006
  88. 88. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1SECTION 11 : GAS TURBINE FUEL SYSTEMSINTRODUCTIONThe complete engine fuel system consists of two sub-systems, a low pressure (LP)fuel system and a high pressure (HP) fuel system. The aircraft fuel system suppliesfuel at low pressure to the engine where the engine LP system will then supply fuelin a satisfactory state to the HP system. The HP system converts the fuel from aliquid state into an atomised state suitable for burning, at the correct rate to sustain aselected RPM.REQUIREMENTSThe fuel system must also ensure a fuel supply, at all times, correctly atomised forcombustion and also to satisfy the basic requirements for fuel control:o Provide a fuel flow relative to the selected positionof the cockpit lever thus allowing manual selectionof RPM.o Prevent excessive overfuelling during rapidmovements of the cockpit lever.o Correct the fuel flow for variations in engine intakeconditions to maintain a selected RPM.o Ensure that the maximum limits for the engine arenot exceeded.LP COCKPurpose To isolate the engine system from the a/c system in the eventof an emergency or for maintenance purposes.Type Normally a ball-type valve mechanically or electrically (via arotary actuator) operated through 90°.Operation They are rotated from fully closed to full flow position (and viceversa) mechanically or electrically and will have mechanicalstops, as well as limit switches if electrically actuated.Issued May 2006 Chapter 1 : Section 11 1
  89. 89. © Air Service Training (Engineering) LimitedTurbine Engine Technology Part 1 Part 66 – B1/018LP FUEL PUMPPurpose To supply a constant head of fuel to the HP fuel system in theevent of LP booster pump (a/c tanks) failure and thereby toprevent cavitation at the HP fuel pump.Type An engine driven flow-type pump, normally employing acentrifugal impeller.Operation Accepts fuel from a/c fuel system and passes it to the engineHP system at an acceptable rate to maintain a constant, andpositive, head of pressure.2 Chapter 1 : Section 11 Issued May 2006
  90. 90. © Air Service Training (Engineering) LimitedPart 66 – B1/018 Turbine Engine Technology Part 1FUEL COOLED OIL COOLER AND FUEL HEATERPurpose To raise the temperature of the fuel to a level where ice particleswill not form. These particles would otherwise block the filter(s)and interfere with the HP controlling functionsType A radiator-type matrix, one side containing fuel and the other sidecontaining the hot fluid (engine oil or compressor air).Operation Oil heated fuel heaters are continuous flow and non-controlledheaters. Those using compressor air are either manually orautomatically controlled non-continuous types and are generallyused in conjunction with oil heater types – in which case they willbe arranged in series with the oil heated assembly being first inline.As a minimum requirement there will be a temperature bulb inthe outlet from the heaters sending a signal to the appropriateindicator on the flight deck.Issued May 2006 Chapter 1 : Section 11 3