Module 15 Gas Turbine Engine Category B1.1 PDF

Bangladesh Airlines Training Center BATC Training Note for Module 15 Gas Turbine Engine Category B1.1 EASA – 147 Cours...

Bangladesh Airlines Training Center BATC Training Note for Module 15 Gas Turbine Engine Category B1.1 EASA – 147 Course Notes This training note has been prepared for Module 15 Gas Turbine Engine, Category B1.1 in accordance with the syllabus prescribed in EASA Part 66 to fulfill the requirements of EASA Part147. Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 1 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC INTENTIONALLY LEFT BLANK EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 2 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Index Chapter Subchapters Topics Page 15.1 Engine 15.1.1 Introduction 1.1 Fundamentals 15.1.2 Definitions and Terminologies 1.1 15.1.3 Forms of Mechanical Energy 1.3 15.1.4 Newton's laws of motion 1.3 15.1.5 The Brayton Cycle 1.5 15.1.6 Types of Gas turbine Engines 1.6 15.2 Engine 15.2.1 Introduction 2.1 Performance 15.2.2 Thrust 2.1 15.2.3 Engine efficiencies 2.4 15.2.4 By-pass ratio and engine 2.7 pressure ratio 15.2.5 Pressure, Temperature and 2.8 velocity of gas flow 15.2.6 Factors affecting thrust 2.9 15.3 Inlet 15.3.1 Introduction 3.1 15.3.2 Compressor inlet ducts and 3.1 Effects of various inlet configurations 15.3.3 Ice protection 3.5 15.4 15.4.1 Introduction 4.1 Compressors 15.4.2 Types of Compressors 4.1 15.4.3 Features of centrifugal 4.4 compressor 15.4.4 Features of axial flow 4.7 compressor 15.4.5 Fan balancing 4.13 15.4.6 Causes and effects of 4.15 compressor stall and surge 15.4.7 Method of air flow control 4.19 EASA – 147 Course Notes 15.5 15.5.1 Introduction 5.1 Combustion 15.5.2 Basic Construction 5.1 Section 15.5.3 Combustion Process 5.2 15.5.4 Types of Combustion Chamber 5.6 15.5.5 Materials of Combustion 5.9 Chamber 15.6 Turbine 15.6.1 Introduction 6.1 Section 15.6.2 Types of turbine 6.1 15.6.3 Nozzle guide vanes 6.8 15.6.4 Blade to disk attachment 6.9 15.6.5 Causes and effects of turbine 6.10 blade types Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 3 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Chapter Subchapters Topics Page 15.7 Exhaust 15.7.1 Constructional features and 7.1 principles of operation 15.7.2 Convergent, divergent and 7.3 variable area nozzles 15.7.3 Engine noise reduction 7.5 15.7.4 Thrust reversers 7.9 15.8 Bearing 15.8.1 Seals 8.1 and Seals 15.8.2 Bearings 8.4 15.9 Lubricants 15.9.1 Introduction 9.1 and Fuels 15.9.2 Properties 9.1 15.9.3 Fuel additives 9.3 15.9.4 Safety precautions 9.4 15.10 15.10.1 System components 10.1 Lubrication 15.10.2 Oil Sub Systems 10.6 Systems 15.11 Gas 15.11.1 Introduction 11.1 Turbine Engine 15.11.2 Hydro-mechanical Fuel Control 11.1 Fuel system 15.11.3 FADEC 11.20 15.11.4 System layout and component 11.24 15.12 AirS 15.12.1 Introduction 12.1 Systems 15.12.2 Operation of engine air 12.1 distribution 15.12.3 Anti-ice control system 12.6 15.12.4 Hot Air System 12.9 15.12.5 External cooling and ventilation 12.10 15.13 Strating 15.13.1 Introduction 13.1 and Ignition 15.13.2 Engine start systems and 13.1 components 15.13.3 Ignition system and component 13.10 15.13.4 Maintenance safety 13.14 EASA – 147 Course Notes requirements 15.14 Engine 15.14.1 Exhaust gas temperature 14.1 Indication 15.14.2 Engine thrust indication 14.4 systems 15.14.3 Oil pressure and temperature 14.5 15.14.4 Fuel pressure and flow 14.6 15.14.5 Engine Speed Indicator 14.7 15.14.6 Vibration measurement and 14.9 indication 15.14.7 Torque Indication 14.10 15.14.8 Power Indication 14.11 15.14.9 Engine Failure Detection 14.12 Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 4 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Chapter Subchapter Topics Page 15.15 Power 15.15.1 Operation and applications of 15.1 augmentation Water injection, water methanol systems 15.15.2 Operation and applications of 15.4 After burner system 15.16 Turbo- 15.16.1 Gas couple/free turbine and 16.1 prop engine gear couple turbine 15.16.2 Reduction gears 16.3 15.16.3 Integrated engine and propeller 16.5 controls 15.16.4 Propeller control 16.20 15.16.5 Over speed safety devices 16.24 15.17 Turbo 17.1 shaft Engine 15.18 Auxiliary 15.18.1 Introduction 18.1 Power unit 15.18.2 Purpose 18.2 15.18.3 General description 18.3 15.18.4 APU Protective System 18.8 15.19 Power 15.19.1 Fire walls 19.1 plant 15.19.2 Cowlings 19.2 Installation 15.19.3 Engine mounts 19.11 15.19.4 Hoses and Pipes 19.15 15.19.5 Engine Drains 19.21 15.20 Fire 15.20.1 Introduction 20.1 Protection 15.20.2 Fire detection systems 20.1 Systems 15.20.3 Fire extinguishants 20.8 15.21 Engine 15.21.1 Procedures for Starting and 21.1 monitoring and Ground Running ground 15.21.2 Starting 21.3 operation 15.21.3 Unsatisfactory Starts 21.6 EASA – 147 Course Notes 15.21.4 Engine Stopping 21.7 15.21.5 Engine Fires 21.8 15.21.6 Interpretation of Engine Power 21.8 Outputs and Parameters 15.21.7 Trend Monitoring 21.14 15.21.8 Inspection of Engine and 21.28 Components 15.21.9 Compressor Cleaning 21.42 15.21.10 Foreign Object Damage 21.44 15.22 Engine Preservation and de- 22.1 storage and preservation for the engine preservation Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 5 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC INTENTIONALLY LEFT BLANK EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 6 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC 15.1 Engine Fundamentals 15.1.1. Introduction An engine is a thermal device that converts heat energy into mechanical energy. Mechanical energy is principally derived in the form of torque on the output shaft of the engine and is utilized for necessary driving works. Some examples of such driving works by engine shafts are as follows: (i) Rotating the armature of a generator to produce electricity. (ii) Driving the propeller of an aircraft or a ship or driving the rotor of a helicopter. (iii) Driving wheel axles of locomotives. A gas turbine engine uses turbines to convert heat energy of a gas into torque; the gas is the combustion- product produced by burning fuel in compressed air in its combustion chamber inside the engine. A turbine is a rotary device with arrangement of series of blades around the periphery of a wheel mounted on a shaft so that energy of the working fluid, when impinged over blades, will rotate the wheel. In short, turbine is an energy transfer mechanism, transferring energy from working fluid to its shaft in the form of rotation or torque. In aircraft application, an engine is a propulsive device that provides propulsive force (thrust) to propel the aircraft forward overcoming atmospheric drag. Thus, as long as aircraft propulsion is concerned, the objective of the aircraft gas turbine engine is not directly the work output at its shaft but is the propulsive force. This is fundamental difference between the primary objective furnished by the gas turbine engines in industrial applications and in the aircraft applications. However, the basic aerodynamic and thermodynamic considerations are almost the same. 15.1.2 Definitions and Terminologies Force: A Force is that which, when acting on a body which is free to move, causes it to move, or conversely, that which stops, or changes the direction of a moving body. (Units: Newton or Pound force). Force is produced when a mass is accelerated. Force = Mass x Acceleration (F = M x a) i.e. a Force moves the Piston down the cylinder. Gravity, for example, is a force that attracts bodies towards the center of the earth at a rate that will cause the object to increase its velocity by 32.2 ft/sec (9.81 m/sec) for each second EASA – 147 Course Notes the object is falling i.e. at the end of 2 sec, the speed would be 64.4 ft/sec (19.62 m/sec) and so on. Figure 32.2ft/sec/sec. (32.2 ft/sec2 or 9.81 m/sec2) is called acceleration due to gravity and it is represented in formulas by the letter g. This value can be used to determine the amount of resistant an object of given weight offers to motion. When the weight is divided by the acceleration constant (g), the quotient is called the mass of the object i.e. Mass = w/g Force is a vector quantity; that is, it has both magnitude and direction. When we speak of 1000 lb (453.6 kg) of force acting on an object, we cannot know its effect unless we know the direction of the force. Two or more forces acting on a body will produce a resultant force. In Turbine Engine problems we often consider force to be produced by a fluid. In the English system of measurement, force may be found by the formula: F = P x A where, F = Force measured in pounds Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 1 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC P = Pressure measured in pounds per square inch A = Area on which the force is acting, measured in square inch Work: Mechanical work is done when a force acts on a body and causes it to move through a distance. In the English system of measurement, work may be found by the formula: W = F x d where, F = Force measured in pounds d = Distance through which the force acts, measured in ft W =Work done measured in ft-pounds Force is a scalar quantity; that is, it has only the magnitude. It is important to remember that work is accomplished only when an object is moved some distance by an applied force. For example, if an object that is pushed as hard possible fails to move, then by the definition, no work has been done. Figure 15.1.1: Works equals force times distance Example: A jet engine that is exerting 1000 lb of fo [44,480 N] moves an airplane 10 ft [3 m]. How much w is being accomplished? Solution: We know, W =Fxd = 1000 X 10 W = 10,000 ft. lb [13,560 N m] Energy: Energy is defined as the capacity for doing work. In the Gas Turbine Engine this energy produces both motion and heat. The energy that bodies possess can be classified into two categories: EASA – 147 Course Notes  Potential Energy  Kinetic Energy Potential Energy Potential energy exists within a body because of its configuration or its position. In a fluid, potential energy is often stored in the form of pressure. If the mass and height of aa object are known, the potential energy can be determined from the formula: EP = mgh or EP = wh (since, Weight of an object = Mass of the object x Acceleration due to gravity) where, w = Weight of the object in lb or Kg h = Height of the object in ft or m EP = Potential energy of the object in ft.lb or Kg.m Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 2 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Example: A 20,000 lb (9072 Kg) airplane is held 5 ft (1.52 m) off the floor by a jack. How much energy does this system possess? Solution: We know, EP = wh w = Weight of the object in lb or Kg h = Height of the object in ft or m So, EP = Potential energy of the object in ft.lb or Kg.m = 20,000 x 5 = 100, 000 ft.lb (13,830 Kg.m) Kinetic Energy Kinetic Energy is possessed by a body because of its motion. Gases striking the turbine wheel exhibit kinetic Energy. If the mass and speed of a body are known, the kinetic Energy can be determined from the formula: EK = mv2 or, EK = wv2 where, w = Weight of the object in lb or Kg v = Velocity of the object in ft/s or m/s EP = Kinetic energy of the object in ft.lb or Kg.m Example: An airplane weighing 6440 lb (2924 kg) has a velocity of 205 mph(300 ft/s) [330 km/h (91.6 m/s)].Find the Kinetic energy. Solution: We know, EK = wv2 Here, w = Weight of airplane = 6440 lb (2924 kg) v = Velocity of airplane = 205 mph (300 ft/s) [330 km/h (91.6 m/s)] So, EP = Kinetic energy of the airplane = X 6400 X 3302. =9,000,000 ft.lb [1,244,700 Kg.m] The Law of Conservation of Energy states that: "Energy can be neither created nor destroyed; only its form may be changed" i.e. the chemical energy of the fuel is converted to heat energy during combustion in the engine. The engine then converts this to mechanical energy for fulfilling its different purposes. Power: EASA – 147 Course Notes Nothing in the definition of work states how fast the work is being done. Power is the rate of doing work or work done per unit time. Work is done as the piston moves in the cylinder. It is moved so many times a minute, and so the power can be measured. Energy is found by the formula: P= i.e. P = Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 3 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.1.2: Rate of work done Where, P = Power in ft.lbs per seconds (ft.lbs per minute) [Jouls per seconds (Watt)] F = Force in seconds d = Distance through the force acts in ft or m t = Time in seconds or minutes Power may be expressed in any one of the several ways, depending on the units used for the force, the distance, and the time. Power is often expressed in units of Horsepower. One Horsepower (HP) is equal to 33,000 ft.lb/min [4554 Kg.m/min] or 550 ft.lbs/s [69Kg.m/s]. In other words, a I HP motor can raise 33,000 lbs a distance of 1 ft in 1 min or 550 lbs a distance of 1 ft in 1 s.(Figure 15.1.2).Therefore,. /. / 1 HP = = , Example: A 5000 lb [2250 Kg] weight is lifted a distance of 10 ft [3 min] in 2 min. How much horsepower is required? Solution: We know, P = = = 25,000 ft.lb/min [3457 Kg.m/min]. / ,. / Therefore, Power in Horsepower (HP) = = = 0.75 or , , EASA – 147 Course Notes Acceleration: Acceleration of a body in motion is defined as an increase of velocity with respect of time. However, since velocity is a vector, acceleration describes the rate of change of both the magnitude and the direction of velocity. The definition is not based on the distance travelled, but on the loss (Deceleration) or gain (Acceleration) of velocity with time. Acceleration is found by the formula: Acceleration = = 𝐯 𝐮 So, a = 𝐭 Where, a = Acceleration in (ft/s)/s [(m/s)/s] i.e. ft/s2 [m/s2] u = Initial velocity in ft/s [m/s] v = Final velocity in ft/s [m/s] Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 4 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC t = Time in seconds i.e. s Deceleration is the decrease of velocity with respect of time and it is found in the same way as acceleration, but by subtracting the final velocity from the initial velocity. Therefore, we can say “negative acceleration will be treated as deceleration”. Momentum: Mass times the velocity or mv defines momentum. It is the property of a moving body that determines the length of time required to bring it to rest under the action of a constant force. Large objectives with a lot of mass but very little velocity can have a much momentum as low-mass objectives with very high velocity. A boat must dock very slowly and carefully because if it touches the dock very gently, it may crush it. On the other hand, a bullet weighs very little but its penetrating power is very high because of its high velocity. 15.1.3. Forms of Mechanical Energy Mechanical energy may be subdivided into three different forms; potential energy (EP), strain energy and kinetic energy (EK). EP is the energy possessed by a body by virtue of its position, relative to some datum. The change in EP is equal to its weight multiplied by the change in height. Since the weight of a body is mg, then the change in EP may be written as: Change in ∆EP = mg∆h which of course is identical to the work done in overcoming gravity, So, the work done in raising a mass to a height is equal to the EP it possesses at that height, assuming no external losses. EK is energy possessed by a body by virtue of its motion. Translational EK i.e. the EK of a body travelling in a linear direction (straight line) is: Translational Kinetic Energy, EK = mv2 = (Mass X Velocity2) Or, EK = wv2 (since Mass of an object = ). Flywheels are heavy wheel-shaped masses fitted to shafts in order to minimize sudden variations in the rotational speed of the shaft, due to sudden changes in load. A flywheel is therefore a store of rotational EK. EASA – 147 Course Notes Rotational EK can be defined in a similar manner to translational EK, i.e. Rotational EK = Iɷ2 J Where, I =mass moment of inertia (which you met when we studied torsion) ɷ = Angular velocity of the rotating object in rad/s. 15.1.4. Newton's laws of motion The fundamental laws of jet propulsion were demonstrated many years ago by recognized scientists and experimenters. These laws, and the equations derived from them, must be discussed in order to understand the operating principle of the gas turbine engine. Foremost among these scientists was Sir Isaac Newton of England, who derived three laws pertaining to bodies at rest and in motion and the forces acting on these bodies. Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 5 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Newton's first law states that "A body (mass) in a state of rest tends to remain at rest, and a body in motion tends to continue to move at a constant speed, in a straight line, unless acted upon by some external force." The portion of the law that states "a body in a state of rest tends to remain at rest" is acceptable from our own experience. But the second part that states "a body in motion tends to remain in motion at a speed and in a straight line" is more difficult to accept. For example, the less friction that a bearing offers, longer a wheel will coast. Therefore, according to law, if all friction were removed, the wheel coast forever. Newton's second law of motion says that "An unbalance force on a body tends to produce an acceleration in direction of the force and that the acceleration, if is directly proportional to the force and inversely proportional to the mass of the body." 𝐅 a= 𝐦 Where, a = Acceleration of the moving body F = Force exerted on the moving body m = mass of the moving body A ball thrown with a force that accelerates it at the rate of 50 ft/s2 [15.24 m/s2] will need double this force accelerate the ball to 100 ft/s2 [30.48 m/s2]. On other hand, if the mass of the ball is doubled the rate of acceleration would be halved, or 25 [7.62 m/s2]. If each side of the equation is multiplied by m, then F = ma Newton's third law states that "For every acting force there is an equal and opposite reacting force”. The acting force means the force exerted by one body on another, while the reacting force means the EASA – 147 Course Notes Fig 15.1.3 applications of Newton’s third law of motion force second body exerts on the first. These forces always occur in pairs but never cancel each other although equal in magnitude; they always act on different objects. Examples of the third law are to found in everyday life (fig. 15.1.3). When a person jumps from a boat, it is pushed backward with same force that pushes the person forward. It should be noted that the person gains the same amount of momentum as the boat received, but in the opposite direction The equation for momentum equals mass times velocity. Since the momentum of both the person and the boat be equal, then m1v1 = m2v2 Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 6 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Example: A man weighing 150 lb [68.04 kg] jumps from his boat to shore at a velocity of 2 ft/s [0.61 m/s]. If the boat weighs 75 lb [34.02 kg], what will be its velocity? Solution: We know, m1v1 = m2v2 Here, m1 = Mass of man = w1/g, (Weight of man, w1 = 150 lb) v1 = Velocity of man to jump = 2 ft/s m2 = Mass of boat = w2/g, (Weight of boat, w2 = 75 lb) v2 = Velocity of boat =? Now, v2 = (m1/m2) v1 So, v2 = (w1/w2) v1 = (150/75) x2 = 4 ft/s 15.1.5. The Brayton Cycle (The Constant-Pressure Engine) A cycle is a process that begins with certain conditions and ends with those same conditions. The Brayton cycle describes the events that take place in a turbine engine as the fuel release its energy. When energy is added, the air remains at a relatively constant pressure, but its volume is increased which increases the velocity of the air as it leaves the engine. EASA – 147 Course Notes Figure 15.1.4: The principle of operation of a GTE- Suck, Squeeze, Bang, and Blow Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 7 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Actual Brayton cycle: Figure 15.1.5: Idealized Brayton cycle (No Inlet duct and Exhaust pipe is installed) 1. Adiabatic compression (in compressor). Process: 1-2 2. Isobaric (constant pressure) process heat addition (in combustion chamber). Process: 2-3 3. Adiabatic expansion (in turbine). Process: 3-4 4. Isobaric (constant pressure) heat rejection (exhaust). Process: 4-1 Brayton cycle for a Gas Turbine Engine: The four continuous events shown on the pressure volume graph of (fig 15.1.6) are intake, compression, expansion, and exhaust. Intake: The air entering the inlet duct of the engine is at essentially ambient pressure. Ultimate goal of an Air Inlet Duct designer is to recover the total pressure of the air passing thru’ it. The static pressure thru’ the Air Inlet Duct increases slightly in the process 0-2. Compression: The thermodynamic (adiabatic compression) process 2-3 shows that the air pressure rises from ambient as the compressor does work (-ve) on the air. Here work is done on the system (air) by the surrounding (compressor). Compressor increases the pressure and decreases volume of the air passing thru’ it. EASA – 147 Course Notes Expansion: When energy is added to the air from the fuel burned in the combustion chamber, the pressure remains relatively constant, but the volume increases greatly in accordance with isobaric process 3-4. It is because of this characteristic that the Brayton cycle is called a constant pressure cycle. But in practice the exit of a combustion chamber is slightly narrow to speed up the flowing gases which cause decrease in pressure. Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 8 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.1.6: Brayton cycle (Inlet duct and Exhaust pipe is installed) Exhaust: When the heated air leaves the combustion chamber, it passes through the turbine where the pressure drops, but its volume continues to increase and it is expressed as thermodynamic (adiabatic expansion) process 4-5. The burning gases have heated the air and expanded it greatly, and since there is little opposition to the flow of these expanding gases as they leave the engine, they are accelerated greatly. Here work is done by the system (air) on the surrounding (turbine). So the work done is +ve. Some of the energy is extracted from the exiting gases by the turbine and this is used to drive the accessories and the various engine accessories (Figure 15.1.7). Because the turbine and compressor are on the same shaft, the work done on the turbine is exactly equal to the work done by the compressor and, ideally, the temperature change is the same. The nozzle (converging exhaust system where the pressure continues to drop to ambient and the velocity continues to increase) then brings the flow isentropically back to free stream pressure from station 5 to station 8. Externally, the flow conditions return to free stream conditions, which complete the cycle. In this cycle, work is accomplished by increasing the velocity of the air as it passes through the engine. The area under the T-s diagram is proportional to the useful work and thrust generated by the engine. EASA – 147 Course Notes Figure 15.1.7: Gas flow diagram Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 9 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC After the air leaves the turbine, it passes through a converging exhaust system where the pressure continues to drop to ambient and the velocity continues to increase. In this cycle, work is accomplished by increasing the velocity of the air as it passes through the engine. 15.1.6 Construction and Types of Gas turbine Engines A typical gas turbine engine consists of: 1. An air inlet, 2. Compressor section, 3. Combustion section, 4. Turbine section, 5. Exhaust section, 6. Accessory section, and 7. The systems necessary for starting, lubrication, fuel supply, and auxiliary purposes, such as anti- icing, cooling, and pressurization. Gas turbine engines can be classified according to the type of compressor used; the path the air takes through the engine, the way the power produced is extracted or used. Compressor types fall into three categories: 01. Centrifugal Flow Compressor 02. Axial Flow Compressor 03. Axial-Centrifugal Flow Compressor In addition, power usage produces the following engine divisions: 01. Turbojet 02. Turbofan 03. Turboprop 04. Turbo shaft EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 10 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC EASA – 147 Course Notes Turbojet Engine The modern turbojet engine is built with many variations, but the basic components are still the compressor, the combustor, and the turbine. A modern turbojet engine (fig-15.1.8) produces its thrust from the acceleration of the flow of hot gases. Air enters the engine inlet and flows into the compressor where its pressure is increased. Fuel is added in the combustor where it is ignited and burns, expanding the gases. As the expanded gases flow out of the engine, they pass through the turbine where part of their energy is given up to spin the turbine which drives the compressor. Energy that remains in the gases as they leave the tail pipe is in the form of velocity energy, and this produces the reaction we know as thrust. Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 11 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.1.8: Turbojet engine A turbojet derives its thrust by highly accelerating a small of air, all of which goes through the engine. Since a high “jet” velocity is required to obtain an acceptable amount of thrust, the turbine of a turbojet is designed to extract only enough power from the hot gas stream to drive the compressor and accessories. All of the propulsive force produced by a jet engine is derived from the imbalance of forces within the engine itself. The turbojet characteristics and uses are as follows: 01. Low thrust at low forward speeds. 02. Relatively high, thrust-specific fuel consumption (TSFC) at low altitudes and airspeeds, a disadvantage that decreases as the altitudes and airspeeds increase. 03. Long takeoff roll. 04. Small frontal area, resulting in low drag and reduced ground-clearance problems. 05. Lightest specific weight (weight per pound of thrust produced) 06. Ability to take advantage of high ram-pressure ratios. 07. The turbojet engine has problems with noise and fuel consumption in the speed range that airliners fly (0.8 Mach). EASA – 147 Course Notes The characteristics suggest that the turbojet engine would be best for high-speed, high-altitude, long- distance flights. Turbofan Engine The turbofan engine has a duct-enclosed fan mounted at the front or rear of the engine and driven either mechanically geared down or at the same speed as the compressor, by an independent turbine located to the rear of the compressor drive turbine. There are two methods of handling the fan air. Either the fan air can exit separately from the primary engine air (short duct), or it can be ducted back to mix with the primary engine’s air the rear (long duct). On some long duct engines the primary and secondary airflow may be mixed internally and then exit from a common nozzle, or the two gas streams may be kept separate for the entire length of the engine. If Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 12 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC the fan air is ducted to the rear, the total fan air pressure must be higher than the static gas pressure in the primary engine’s exhaust, or air will not flow. By the same token, the static fan discharge pressure must be less than the total pressure in the primary engine’s exhaust, or the turbine will not be able to extract the energy required to drive the compressor and fan. By closing down the area of flow of the fan duct, the static pressure can be reduced and the dynamic pressure increased. A turbofan engine (Figure 15.1.9) consists basically of a multi-bladed ducted propeller driven by a gas turbine. The fan has a compression ratio somewhere in the range of 2:1, or it produces a compression pressure of approximately two atmospheres. The ducted fan gives a turbofan engine cruise speed capabilities similar to those of a turbojet engine, yet its greater propulsive efficiency at low speeds gives an airplanes equipped with turbofan engines far better short-field takeoff characteristics than one having turbojet engines. The fan has a diameter and mass flow much less than that of a propeller, but it moves air from its convergent exhaust nozzle with a greater velocity. On the other hand, the fan discharge is much slower than the exhaust of a comparable turbojet engine, but it has a greater mass flow. There are several fan configurations. The fan can be bolted directly to the compressor and rotate at the same speed, or it can be connected through a reduction gear system to the compressor. The fan on some engines is driven by a separate turbine and rotates independently of the compressor and in some engines the fan is mounted in the turbine section as an extension of the turbine wheel blades. An engine with the fan in the turbine section is called an aft-fan engine, and those with the fan in front are called forward-fan engines. The aft-fan configuration is not a popular design today because the fan does not contribute to the compressor pressure ratio. Turbofans are usually grouped into three classifications depending upon the amount of air the bypass around the gas generator i.e. Low Bypass, Medium Bypass and High Bypass Engines. The efficiency of the fan engine is increased over that of the pure jet by converting more of the fuel energy into pressure energy rather than kinetic (dynamic) energy of a high velocity exhaust gas stream. Pressure times the area equals a force. The fan produces this additional force or thrust without increasing fuel flow. The turbofan characteristics and uses are as follows: EASA – 147 Course Notes 01. Increased thrust as forward speeds similar to a turboprop results in a relatively short takeoff. However, unlike the turboprop, the turbofan thrust is not penalized with increasing air speed, up to approximately Mach 1 with current fan design. 02. Ground clearances are less than turboprop but not as good as turbojet. 03. TSFC and Specific Weight fall between the Turbojet and Turboprop resulting in increased operating economy and aircraft range over the Turbojet. 04. Considerable noise level reduction of 10 to 20 percent over the Turbojet reduces acoustic fatigue in surrounding aircraft parts and is less than objectionable to people on the ground. Also, No noise suppressor is needed. On newer fan engines, such as the General Electric CF6 and Pratt & Whitney 4000 series and others, Inlet Guide Vanes have been eliminated to reduce the fan noise, which is considered to be a large problem for high-bypass-ratio fan engines. The noise level is reduced by elimination of the discrete frequencies that are generated by the fan blades cutting Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 13 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC through the wakes behind the vanes. Other fan-noise-reducing features are also incorporated with the Exhaust Systems. 05. The Turbofan is superior to Turbojet in “Hot day” performance. 06. Two thrust reversers are required if the fan air and primary engine air exit through separate fan nozzles, the advantage of which is the short fan duct with corresponding low duct loss. The above characteristics show that the Turbofan engine is suitable for long-range, relatively high-speed flight and has a definite place in the prolific gas turbine family. In a low by-pass engine, the fan and compressor section utilize approximately the same mass airflow, but the fan discharge will generally be slightly greater than that of the compressor. The fan discharge air may be ducted directly over-board from a short fan duct or it may pass along the entire length of the engine in what is called a long fan duct. In either case the end of the duct has a converging discharge nozzle to produce a velocity increase and reactive thrust. EASA – 147 Course Notes Figure 15.1.9: Double spool hi-bypass Turbofan engine In a fully ducted fan engine, the hot and cold air streams generally mix before they are discharged into the atmosphere and this mixing helps attenuate, or lessen, the noise. The air in the core engine is compressed, burned, and discharged in the normal manner and the thrust ratio of the two gas streams is approximately 1:1. Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 14 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.1.10: Triple-spool hi-bypass Turbofan Engine A medium, or intermediate, by-pass engine has an airflow by-pass ratio of between 2:1 and 3:1, and has a thrust ratio that is approximately the same as its by-pass ratio. The fan used on these engines has a larger diameter than that on a low by-pass engine of comparable power and its diameter is determined by both the by-pass ratio and the thrust output of the fan compared with the thrust obtained from the core engine. This latter ratio is often called the cold-stream to hot-stream ratio. A high by-pass turbofan (fig 15.1.9) engine has a fan ratio of 4:1 or greater and has an even wider diameter fan in order to move more air. The turbofan engine has become the most widely used engine type because it offers the best fuel economy. This economy is obtained by increasing the total mass airflow while decreasing the velocity of the hot exhaust wake. Turbo-shaft Engine A gas turbine engine that delivers power through a shaft to operate something other than a propeller is referred to as a turbo-shaft engine. These are widely used in such industrial applications as electrical EASA – 147 Course Notes power generating plants and surface transportation systems, while in aviation, turbo-shaft engines are used to power many modern helicopters. The turbo-shaft power takeoff may be coupled to and driven directly by the turbine that drives the compressor, but it is more likely to be driven by a turbine of its own. Engines using a separate turbine for the power takeoff are called free-turbine engines or free power turbine-type turbo-shaft engines. Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 15 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.1.11: Turbo-shaft engine A free-turbine turbo-shaft engine (Figure-15.1.11) has two major sections, the gas generator and the free turbine section. The function of the gas generator is to produce the required energy to drive the free turbine system, and it extracts about two thirds of the energy available from the combustion process leaving the other one third to drive the free power turbine. Turboprop Engine A turboprop engine is similar in design to a turbo-shaft engine except that it uses a reduction gear system to drive the propeller. Fig 15.1.11 Turboprop engine EASA – 147 Course Notes Figure 15.1.12: Turboprop engine Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 16 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC The propeller may be driven from the gas generator turbine (fig-15.1.12), or it may use its own free turbine in the same way as the turbo-shaft engine. The free turbine arrangement allows the propeller driving turbine to seek its optimum speed while the compressor turbine operates at the best speed for the compressor efficiency. Propulsion in a Turboprop engine is accomplished by the conversion of the majority of the gas-stream energy into mechanical power to drive the compressor, accessories, and the propeller load. Unlike a Turbojet engine, propeller of a Turboprop engine imparts a small acceleration to a large mass of air. Only a small amount (approximately 10 percent) of “jet” thrust is available in the relatively low-pressure, low- velocity gas stream created by the additional turbine stages needed to drive the extra load of the propeller. The turboprop characteristics and uses are as follows: 01. High propulsive efficiency at low speeds, which results in shorter takeoff rolls but falls off rapidly as airspeed increases (shown in Figures 15.1.13 (a) & 15.1.13 (b)). The engine is able to develop high thrust at low airspeeds because the propeller can accelerate large quantities of air at zero forward velocity of the airplane. EASA – 147 Course Notes Fig 15.1.13 (a): Thrust compare to airspeed at sea level and at 30,000 ft Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 17 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Fig 15.1.13 (b): Thrust specific fuel consumption (TSFC) versus airspeed at sea level and at 30,000 ft 02. More complicated design and heavier weight than a Turbojet engine. 03. Lowest TSFC. 04. Large frontal area of propeller and combustion that necessitates longer landing gears for low- wing airplanes but does not necessarily increase parasite drag. 05. Possibility of efficient reverse thrust. These characteristics show that Turboprop engines are superior or for lifting heavy loads off short and medium-length runways. Turboprops are currently limited in speeds to approximately 500 mph [805 km/h], since propeller efficiencies fall off rapidly with increasing airspeeds of shock wave formations. EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |1- 18 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Chapter 02 Engine Performance EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 1 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC INTENTIONALLY LEFT BLANK EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 2 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Index Chapter Subchapters Topics Page 15.2 Engine Performance 15.2.1 Introduction 2.1 15.2.2 Thrust 2.1 15.2.3 Engine efficiencies 2.4 15.2.4 By-pass ratio and engine pressure ratio 2.7 15.2.5 Pressure, Temperature and velocity of 2.8 gas flow 15.2.6 Factors affecting thrust 2.9 EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 3 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC INTENTIONALLY LEFT BLANK EASA – 147 Course Notes Module – 15 | Gas Turbine Engine Category| B1.1 Page |0- 4 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC 15.2 Engine Performance 15.2.1. Introduction The performance requirements of an engine are obviously dictated to a large extent by the type of operation for which the engine is designed. The power of the turbojet engine is measured in thrust, produced at the propelling nozzle or nozzles, and that of the turbo-propeller engine is measured in shaft horse-power (S.H.P.) produced at the propeller shaft. However, both types are in the main assessed on the amount of thrust or S.H.P. they develop for a given weight, fuel consumption and frontal area. Since the thrust or S.H.P. developed is dependent on the mass of air entering the engine and the acceleration imparted to it during the engine cycle, it is obviously influenced, as subsequently described, by such variables as the forward speed of the aircraft, altitude and climatic conditions, These variables influence the efficiency of the air intake, the compressor, the turbine and the jet pipe; consequently, the gas energy available for the production of thrust or S.H.P. also varies. In the interest of fuel economy and aircraft range, the ratio of fuel consumption to thrust or S.H.P. should be as low as possible. This ratio, known as the specific fuel consumption (S.F.C.), is expressed in pounds of fuel per hour per pound of net thrust or S.H.P. and is determined by the thermal and propulsive efficiency of the engine. In recent years considerable progress has been made in reducing S.F.C. and weight. 15.2.2. Thrust EASA – 147 Course Notes Figure 15.2.1: Forces on aircraft Thrust is a mechanical force, so the propulsion system must be in physical contact with a working fluid to produce thrust. Thrust is generated most often through the reaction of accelerating a mass of gas. Since thrust is a force, it is a vector quantity having both a magnitude and a direction. The engine does work on the gas and accelerates the gas to the rear of the engine; Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 1 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC the thrust is generated in the opposite direction from the accelerated gas. The magnitude of the thrust depends on the amount of gas that is accelerated and on the difference in velocity of the gas through the engine. Thrust is measured in "pounds of thrust" in the U.S. and in Newton under the metric system (4.45 Newton of thrust equals 1 pound of thrust). A pound of thrust is the amount of thrust it would take to keep a 1-pound object stationary against the force of gravity on Earth. Gross and Net Thrust Gross thrust, is the thrust produced when the engine is not in motion. The true acceleration of the gas within the engine is the difference in velocity between the air in the inlet duct, and the air as it leaves the exhaust nozzle, and this difference is used in computing net thrust. The loss in thrust involved in taking the air in, at the front of the engine is known as the ram drag. The net thrust is the gross thrust minus the ram drag. Therefore, Fn = Fg - Fr Where, Fn = Net thrust Fg = Gross thrust Fr = Ram Drag (Momentum drag) Engine inlet air velocity times the mass of airflow is the initial momentum. The faster the aircraft goes, the greater the initial momentum and the less the engine can change the momentum. We can calculate the net thrust by the formula: Fn = mvj – mvi = the acceleration of the gases through the engine Where, mvj = gross thrust or momentum of exhaust gases (i.e. Fg) mvi = ram drag or momentum of incoming air due to aircraft speed (i.e. Fr). Thrust Formula Let us consider the aircraft speed is vi ft/see and jet velocity is vj, ft/see and weight of air flow EASA – 147 Course Notes through engine is wa lb/sec. The initial momentum is (wa/g)vi and the final momentum is (wa/g)vj. Therefore, the net momentum thrust i.e. Fn = outgoing momentum – incoming momentum Fn = vj - vi … … …... (i) i.e. Fn = (vj – vi) … … … … (ii) Where, wa = weight of air flow in pounds vi = inlet velocity of air in feet per second vj = final (Jet) velocity of air in feet per second g = acceleration due to gravity (32.2 feet/second2) Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 2 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Since fuel adds some mass to the air flowing through the engine, the same formula must be applied to the weight of the fuel as was applied to the weight of the air , and this must be added to the basic thrust equation(i). So we will get: Fn = vj - vi + vf = vj + vf ) - vi = vj - v i) + vf Notice that because the fuel is carried along with the engine it will never have any initial velocity relative to the engine. Some formulas do not consider the fuel flow effect when computing thrust because the weight of air leakage (bleed) through the engine is approximately equivalent to the weight of the fuel added. So we can apply equation (ii) as a thrust formula to compute the thrust roughly for jet engines, and this formula was complete until the development of the “choked” nozzle. Choked nozzle thrust When the jet nozzle is chocked, the pressure is such that the gases are travelling through it at the speed of sound and cannot be further accelerated. Any increase in internal engine pressure will pass out through the nozzle still in the form of pressure. Even though this pressure energy cannot be retuned into velocity energy, it is not lost. The pressure inside the nozzle is pushing in all directions, but when the neck is open, the air can no longer push in the direction of the nozzle. The pressure in the other direction continues undiminished, and as a result the pressure of the gases will push the engine forward (Figure 15.2.2). EASA – 147 Course Notes Figure 15.2.2: The chocked nozzle Any ambient air (air outside the nozzle) is in the way and will cancel out part of the forward thrust. The completed formula for a Turbojet engine with choked nozzle is: Fn = vj - v i) + vf + AJ(PJ - Pam) Where, Aj = Area of jet nozzle Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 3 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Pj = Static pressure of the nozzle Pam = Ambient pressure at the exhaust nozzle AJ(PJ - Pam) = Choked nozzle thrust By neglecting the fuel flows and losses, we can write the formula for net thrust as: Fn = vj - vi) + AJ(PJ - Pam) which is mostly used by the maintenance personnel. Actual thrust produced for a particular engine will be observed in the manufacturer’s test cell. Thrust distribution The net thrust of an engine is a result of pressure and momentum changes within the engine. Some of these changes produce forward forces while others produce rearward forces (Fig 15.2.3). Whenever there is an increase in total heat energy by burning fuel, or in total pressure energy by compression, or by a change from kinetic energy to pressure energy, as in the diffuser, forward forces are produced. Conversely, rearward forces or thrust losses result when heat or pressure energy decreases or is converted into kinetic energy, as in the nozzle. The rated net thrust of any engine is determined by how much the forward thrust forces exceed the rearward thrust forces (i.e. subtracting the sum of all of the rearward thrust from the sum of all of the forward thrust within the engine). The compressor, diffuser, combustion chamber and tail (exhaust) cone exit areas all exert forward thrust while both the turbine and the tail pipe exit areas exert a rearward thrust. If the areas, pressure acting across these areas, velocities, and mass flows are known at any point in the engine, the forces acting at the point can be calculated. For any point in the engine, the force would be the sum of Fn = vj - vi) or mass x acceleration = ma Plus Fn = AJ(PJ - Pamb) or pressure x area = PA EASA – 147 Course Notes The completed formula for net thrust would then be read: Fn = ma + PA Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 4 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.2.3: Thrust distribution of a typical single-spool axial flow engine Thrust Compared with Horsepower Trust and Horsepower cannot be directly compared because, by definition, power is a force applied through a distance in a given period of time. All of the power produced by a jet engine is consumed internally to turn the compressor and drive the various engine accessories. Therefore, the jet engine does not develop any horsepower in the normally accepted sense but supplies only one of the terms in the horsepower formula. The other term is actually provided by the vehicle in which the engine is installed. To determine the Thrust Horse Power (THP) of the jet engine or for a comparison of thrust and horsepower it is possible to convert gas turbine engine thrust to horsepower by using the following formula: EASA – 147 Course Notes / THP = = / / / From this it can be seen that at 375 mph each pound of thrust will be converted to one horsepower, and that for each speed of the aircraft there will be a different THP. The constant 375 is the miles-pound/hour equivalent of one horsepower. This formula shows that the horsepower produced by a GTE increased as the airspeed increases. Thrust Horse Power may be used for Turbojet and Turbofan engines but it does not apply to power produced by turbo-shaft engines. The formula of THP indicates that it (THP) can only be calculated in flight. When the aircraft is stationary the THP is zero, as the aircraft speed is zero. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 5 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Shaft Horsepower Turbo shaft engines are rated in Shaft Horse Power (SHP) which is measured on a Dynamometer Test. It is the power delivered to the propeller shaft to produce propulsive force is known as Shaft Horse Power (SHP); sometimes it is also called Brake Horse Power (BHP), which means that the amount of power required to brake the shaft. Equivalent Shaft Horsepower Turboprop engines are rated in Equivalent Shaft Horse Power (ESHP). This is the combination of the exhaust thrust converted into THP, and the SHP determined by the dynamometer test. Under static conditions (i.e. Turboprop engine is running but the aircraft is not moving and one Shaft Horse Power equals approximately 2.6 pounds of thrust), THP may be found by dividing the pounds of jet thrust by 2.6. When we add this to the SHP measured by the dynamometer we have the Equivalent Shaft Horse Power of the engine. So we have: ESHP = SHP + (Under static condition, Fn = Fg). If the aircraft is moving, the Equivalent Shaft Horse Power (ESHP) of the engine can be found as follows: / ESHP = SHP + = SHP +. /. / Example: Find the ESHP of a Turboprop engine that produces 187.5 pounds of thrust and develops 580 shaft horsepower. Solution: We know, ESHP = SHP +.. = 580 + 652.1. EASA – 147 Course Notes Specific Fuel Consumption One of the principal measures of jet engine efficiency is called Specific fuel consumption (SFC) or Thrust Specific Fuel Consumption (TSFC). SFC is the ratio between the fuel flow (in pound per hour) and the thrust of the engine (in pounds) i.e. SFC = The amount of fuel is burned per hour per unit of thrust is known as Thrust Specific Fuel Consumption (TSFC). Obviously the more thrust we can obtain per pound of fuel, the more efficient the engine is. The term Equivalent Specific Fuel Consumption (ESFC) is used to compare fuel consumption between Turboprop engines. ESFC = Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 6 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC 15.2.2. Engine efficiencies Mechanical Efficiency The mechanical efficiency of an engine is measured by the ratio of the shaft output or brake horsepower to the indicated horsepower or power developed in the cylinders. For example, if the ratio of the BHP to the IHP is 9:10, then the mechanical efficiency of the engine is 90 percent. In determination of mechanical efficiency, only the losses suffered by the energy that has been delivered to the pistons is considered. The word efficiency may be defined as the ratio of output to input. Volumetric Efficiency V olumetric efficiency is the ratio of the volume of fuel-air charge, burned by the engine, at atmospheric pressure and temperature to the piston displacement. If the cylinder of an engine draws in a charge of fuel and air having a volume at standard atmospheric pressure and temperature which is exactly equal to the piston displacement of the cylinder, the cylinder has a volumetric efficiency of 100 percent. Thermal Efficiency Thermal efficiency (called internal efficiency) is one of the important considerations on GTE performance. It is the ratio of the net amount of work produced by an engine to the amount of energy in the fuel burned, and can be computed by the formula: ηth = x 100% EASA – 147 Course Notes Fig 15.2.4 (a) Fig 15.2.4 (b) Figure 15.2.4: Thermal efficiency The three most important factors affecting the thermal efficiency are turbine inlet temperature, compression ratio, and the component efficiencies of the compressor & turbine. Other factors that Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 7 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC affect thermal efficiency are compressor inlet temperature and combustion efficiency. Fig 15.2.4(a) shows the effect that changing compression ratio (compressor pressure ratio) has on thermal efficiency when compressor inlet temperature and the component efficiencies of the compressor and turbine remain constant. The highest thermal efficiency is produced with a high TET and as technological developments bring out new metal that have high strength at high temperatures, we can allow TET to reach higher values. The effects that compressor and turbine component efficiencies have on thermal efficiency when turbine and compressor inlet temperatures remain constant are shown in Fig 15.2.4(b). Propulsive efficiency A measure of the effectiveness with which an aircraft engine converts the fuel it burns into useful thrust. It is the ratio of the thrust horsepower produced by a propeller to the torque horsepower of the shaft turning the propeller. The efficiency of conversion of kinetic energy to propulsive work is termed as the propulsive or external efficiency. Propulsive efficiency depends on the amount of kinetic energy wasted by propelling mechanism that, in turn, depends on the mass airflow multiplied by the square of its velocity. From this relationship we can see that the high-velocity, relatively low-weight jet exhaust wastes considerably more energy than the propeller with its low-velocity, high-weight airflow (figure 15.2.5). This condition changes, however, as the aircraft speed increases because although the jet stream continues to flow at high velocity from the engine, its velocity relative to the surrounding atmosphere is lessened by the forward speed of the aircraft. Thus energy wasted is reduced. The formula for determining propulsive efficiency (ηp) is: Propulsive efficiency (ηp) = 2/[ + 1] x 100% = x 100% Waste energy dissipated in the jet wake, which represents a loss can be expressed as: (v - u)2 where (v - u) is the waste velocity. EASA – 147 Course Notes Figure 15.2.5: Comparison between turboprop, by-pass and pure turbo-jet engine Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 8 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC From the formula of propulsive efficiency it can be seen that 100% propulsive efficiency would occur if the aircraft speed equaled the exhaust velocity. On some fan engine the fan air is mixed with the exhaust gases and exits from a common nozzle, thereby reducing the exit velocity and consequently increasing the propulsive efficiency of the system, which in turn is reflected in lower SFC. 15.2.4. By-pass ratio and engine pressure ratio By-pass ratio The by-pass ratio is a ratio of the mass of air (secondary air) that goes through the fan duct to the mass of air (primary air) that goes through the engine core (shown in Figure Fig 15.2.5) i.e. By-pass ratio = Turbofan engines are usually grouped into the following three classifications depending upon the amount of air the bypass around the gas generator: (a) Low by-pass engine: In a low by-pass engine, the fan and compressor section utilize approximately the same mass airflow, but the fan discharge will generally be slightly greater than that of the compressor. The fan discharge air may be ducted directly over-board from a short fan duct, or it may pass along the entire length of the engine in what is called a long fan duct. In either case the end of the duct has a converging discharge nozzle to produce a velocity increase and reactive thrust. In a fully ducted fan engine, the hot and cold air streams generally mix before they are discharged into atmosphere and this mixing helps attenuate, or lessen, the noise. The air in the core engine is compressed, burned, and discharged in the normal manner and the thrust ratio of two gas streams is approximately 1:1. (b) Medium by-pass engine: A medium, or intermediate, by-pass engine has an airflow by-pass EASA – 147 Course Notes ratio of between 2:1 and 3:1, and has a thrust ratio that is approximately the same as its by-pass ratio. The fan used on these engines has a larger diameter than that on a low by-pass engine of comparable power and its diameter is determined by both the by-pass ratio and the thrust output of the fan compared with the thrust obtained from the core engine. This later ratio is often called the cold-stream to hot-stream ratio. (c) High by-pass engine: A high by-pass turbofan engine has a fan ratio of 4:1 or greater and has an even wider diameter fan in order to move more air. A typical Turbofan engine (GE 90-115B fitted in B777-300ER aircraft) has a by-pass ratio 9:1 (fan diameter 128 inch or 3.52 meter) produces 115,300lb of Takeoff thrust. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 9 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Very high by-pass ratios, in the order of 15:1, are achieved using prop-fans. These are a variation on the turbo-propeller theme but with advanced technology propellers capable of operating with high efficiency at high aircraft speeds. The Turbofan engine has become the most widely used engine type because it offers the best fuel economy. This economy is obtained by increasing the total mass airflow while decreasing the velocity of the hot exhaust wake. The thrust produced by the Pratt & Whitney JT3D, one of the first fan engines, is approximately 50% greater than the thrust produced by the core turbojet engine while the fuel flow remains basically the same. Engine pressure ratio (EPR) The engine pressure ratio (EPR) is defined to be the total pressure ratio across the engine. Mathematically, 𝐓𝐮𝐫𝐛𝐢𝐧𝐞 𝐃𝐢𝐬𝐜𝐡𝐚𝐫𝐠𝐞 𝐓𝐨𝐭𝐚𝐥 𝐏𝐫𝐞𝐬𝐬𝐮𝐫𝐞 𝐩𝐭𝟓 EPR = 𝐂𝐨𝐦𝐩𝐫𝐞𝐬𝐬𝐨𝐫 𝐈𝐧𝐥𝐞𝐭 𝐓𝐨𝐭𝐚𝐥 𝐏𝐫𝐞𝐬𝐬𝐮𝐫𝐞 𝐩𝐭𝟐 On this slide we show how the flow pressure varies through a typical turbojet engine. The pressure is color-coded with blue indicating the lowest pressure and white the highest pressure. Air is brought into the turbojet through the inlet at the left of the computer drawing. At the rear of the inlet, the air enters the compressor. The compressor acts like many rows of airfoils, with each row producing a small increase in pressure. At the exit of the compressor, the air is at a much higher pressure than free stream. In the burner a small amount of fuel is combined with the air and ignited at near constant pressure. EASA – 147 Course Notes Figure 15.2.6: Flow of Cold-stream and Hot-stream for a typical engine. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 10 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Leaving the burner, the hot exhaust is passed through the turbine. Energy is extracted from the flow by the turbine to turn the compressor, which is linked to the turbine by a central shaft. Some pressure is lost in the hot exhaust during this process, but the pressure entering the nozzle is still greater than free stream. The nozzle then converts the high pressure and temperature into high velocity. Because the exit velocity is greater than the free stream velocity, thrust is created as described by the thrust equation. EPR can be easily measured on an operating engine and displayed to the pilot on a cockpit dial. That is why the ratio is not defined in terms of free stream conditions. Total pressure losses in the inlet are not contained in the EPR. But if we know the EPR, the inlet losses, and the corresponding engine temperature ratio, ETR, we can easily determine the thrust of an engine using the nozzle performance information and the thrust equation. The EPR is simply the product of the pressure ratio across all of the engine components. EPR = pt5 / pt2 = (pt3 / pt2) * (pt4 / pt3) * (pt5 / pt4) EPR = compressor pressure ratio * burner pressure ratio * turbine pressure ratio For a given engine design, we can determine the pressure ratio of each component as given on each of the component thermodynamic slides. EASA – 147 Course Notes Figure 15.2.7: ERP Computation for a typical turbojet engine. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 11 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC 15.2.5. Pressure, Temperature and velocity of gas flow Figure 15.2.8 (a) & (b) illustrate pressure, temperature, and velocity changes through typical gas turbine engines. Figure 15.2.8 (a): For a typical turbojet engine with and without afterburner operation EASA – 147 Course Notes Figure 15.2.8 (b): For a typical turboprop engine Pressure Changes: Air usually enters the front of the compressor at a pressure that is less than ambient, indicating that there is considerable suction at the inlet to the engine. This somewhat negative pressure at Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 12 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC the engine inlet may be partly or completely overcome by ram pressure as the engine speed increases. From this point on, there is a considerable pressure total rise through the successive compression stages, with the rate of rise increasing in the later stages of compression. The exit area of turbine nozzle, and rate of fuel flow all can determine the compression ratio (Compressor discharge static pressure/Ambient static pressure) of the compressor. A final static pressure rise is accomplished in the divergent section of the diffuser. From the diffuser, the air passes through the combustion section where a slight pressure loss is experienced. The combustion-chamber pressure must be lower than the compressor discharge pressure in all phases of engine operations in order to establish direction of airflow toward the rear of the engine and allow the gases to expand as combustion occurs. A sharp drop in pressure occurs as the air is accelerated between the converging passages of the turbine nozzle. The pressure continues to drop across the turbine wheel as some of the pressure energy in the hot gas is converted to a rotational force by wheel. If the engine is equipped with more than one turbine stage, a pressure reduction occurs across each turbine wheel. Pressure changes after the turbine depend upon the type of exhaust nozzle used and whether the nozzle is operating in a choked (gas velocity at the speed of the sound) or non-choked condition. When the gases leave the exhaust nozzle, the pressure continues to drop to ambient. Temperature Changes: Air entering the compressor at sea level on a standard day is at a temperature of 590F [150C]. Due to compression, the temperature through the compressor gradually climbs to a point that is determined by the number of compressor stages and its aerodynamic efficiency (see Compressor Aerodynamics topic). On some large commercial engines, the temperature at the front of the combustion section is approximately 8000F [4270C]. As the air enters the combustion chambers, fuel is added and the temperature is raised to about 35000F [19270C] in the hottest part of the flame. Since this temperature is above the melting point of most metals, the combustion chamber and surrounding parts of the engine are protected by a cooling film of air that is established through proper design of the combustion chamber (see combustion chamber topic). Because of this cooling film, the air entering the turbine section is considerably cooler. The acceleration of EASA – 147 Course Notes air through the turbine section further reduces the temperature. If the engine is operating without the use of an afterburner, there is a slight temperature drop through the exhaust pipe. If the engine is operating with the use of the afterburner, there will be a sharp temperature rise through the exhaust pipe. Velocity Changes: Since a jet engine derives its thrust mainly from the reaction to the action on a stream of air as it flows through the jet engine, the pressure and temperature changes just discussed are important only because they must be present to accomplish the action-reaction process. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 13 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC The velocity of the air at the front of the compressor must be less than sonic for most present- day compressors. In order to achieve this goal, the design of the engine inlet duct of the aircraft is of paramount importance (see the Inlet duct). If ambient air velocity is zero (aircraft is stationary), the air velocity in front of the duct increases as it is drawn into the compressor. Because the incoming air at zero aircraft forward velocity has no kinetic energy relative to the engine intake before entering, it does not contribute to the total compression ratio. This situation changes as the ram recovery point (see the Inlet duct) of the inlet is reached. From this point on, the relative kinetic energy does contribute to the total pressure ratio in the form of ram compression. In a good inlet duct, this compression will occur early and efficiently, with a minimum temperature rise. On the other hand, if the aircraft speed is high subsonic or supersonic, the air’s velocity is slowed in the duct. Airflow velocity through the majority of compressors is almost constant, and in most compressors may decrease slightly. A fairly large drop in airspeed occurs in the enlarging diffuser passage. The turning point where flow velocity starts to increase is in the combustion chamber as the air is forced around the forward end of the combustion chamber inner and through the holes along the sides. A further increase occurs at the rear of the combustion chamber as the hot gases expand and are forced through the slightly smaller area of the transition liner. An extremely sharp rise in velocity, with a corresponding loss of pressure, occurs as the air passes through the converging partitions of the turbine nozzle. This exchange of pressure for velocity is very desirable, since the turbine is designed to operate largely on a velocity drop. As previously explained, the velocity increase is accompanied by a temperature and pressure drop. A large portion of the velocity increase through the nozzle is absorbed by the turbine wheel and applied to drive the compressor and engine accessories. Velocity changes from this point on depend upon the design of the engine. As the air discharges through the orifice formed by the exhaust nozzle, the velocity increases sharply. After the air leaves the turbine, it passes through a converging exhaust system where the pressure continues to drop to ambient and the velocity continues to increase. 15.2.6. Factors affecting thrust EASA – 147 Course Notes The jet engine is much sensitive to operating variables than is the reciprocating engine. Such variables can be divided into following two groups: (a) Design factors: (i) Engine rpm (weight of air) (ii) Size of nozzle area (iii) Weight of fuel flow (iv) Amount of air bled from the compressor (v) Turbine Inlet Temperature (vi) Use of water injection. (b) Non-design factors: Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 14 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC (i) Speed of the aircraft (ram-pressure rise) (ii) Density effect: Temperature and pressure of the air (iii) Amount of humidity. RPM Effect: Fn = vj - v i) + vf + AJ(PJ - Pam) Engine speed in revolutions per minute (rpm) has a very great effect on the thrust developed by a jet engine. Figure 15.2.9 (B- Ram Effect) shows that very little thrust is developed at low rpm as compared with the thrust developed at high engine rpm and that a given rpm change has more effect on thrust at higher engine speeds than at lower engine speeds. The weight of air pumped by a compressor is a function of its rpm. Recalling the formula Fn = vj - v i) Figure 15.2.9: A non-linear relationship exists between engine thrust and RPM. it is evident that increasing the weight of air being pumped will result in an increase in Fn or thrust. As we shall see when we get to the section on compressor, engine speed may not be indiscriminately varied but must be controlled within very close limits, and this limit is imposed by the aerodynamics of the compressor. As the tips of the compressor blades reach the speed of EASA – 147 Course Notes sound, Mach one or slightly higher, their efficiency drops off. The design of the compressor is such that the blades are operated at such a speed that the airflow over them is not allowed to get into a shock stall condition. We can compute the compressor Mach number by using this formula: Tip speed = ft/s and Mach number = Where, π = Pi, a constant 3.1416 d = Compressor Diameter in feet prm = Compressor revolutions per minute Aircraft Speed Effect: Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 15 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC The formula Fn = vj - vi) shows that any increase in the forward velocity of the aircraft will result in a decrease in thrust. The faster the aircraft goes, the greater will be the initial momentum of the air in relation to the engine (v1 increasing). But the jet nozzle velocity is generally fixed by the speed of the sound. Obviously, the vj - vi) difference or momentum change will become smaller as aircraft speed increases and thrust will fall as shown in Figure 15.2.10. Figure 15.2.10: Thrust loss is due to vj - vi) difference becoming smaller The loss of thrust will be partially offset by the increase in the Wa due to ram shown in Figure 15.2.11. EASA – 147 Course Notes Figure 15.2.11: Combining thrust loss due to vj - vi) difference decrease with thrust gain due to ram pressure rise Therefore, two opposing trends occur when an aircraft’s speed is increased. What actually takes place is the net result of these two different effects. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 16 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Figure 15.2.12: Thrust loss is due to vj - vi) and thrust gain due to ram-pressure rise. Not as much thrust is recovered due to the ram-pressure rise as would seem to be indicated on first examination. At high aircraft speeds there is a considerable temperature rise in addition to the rise in pressure shown in Figure 15.2.13. EASA – 147 Course Notes Figure 15.2.13: The ram effect-comes from two factors. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 17 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC The actual weight of airflow increase into the engine will be directly proportional to the rise in pressure and inversely proportional to the square root of the rise in temperature i.e. Wa = δt/√θt where, δt = Total pressure and θt = Total temperature. Losses may also occur in the duct during high speeds as a result of air friction and shock-wave formation. Temperature Effect: The gas turbine engine is very sensitive to variations in the temperature of the air. Many engines are rated with the air at a standard temperature of 590F [150C], although some manufacturers will “Flat rate” their engines to a higher temperature (FRT or CP); that is, the engine is graduated to produce a minimum specific thrust above 590F [150C] up its design FRT or CP. Figure 15.2.14: The effect of OAT on thrust output. Careful Thrust lever manipulation is required at lower temperatures. In any case, if engine operates in air temperatures hotter than standard, there will be less thrust produced. Conversely, EASA – 147 Course Notes engine operates in air temperatures colder than standard day conditions will produce a greater than rated thrust. A rise in temperature will cause the speed of the molecules to increase so that they run into each other harder and move farther apart. When they are farther apart, a given number of molecules will occupy more space. And when a given number of molecules occupy more space, fewer can get into the engine inlet area. This result in a decrease in Wa into the engine with a corresponding decrease in thrust. Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 18 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC Pressure Effect: An increase in pressure results when there are more molecules per cubic foot. When this situation occurs, there are more molecules available to enter the engine inlet area, and as a result, an increase in Wa occurs through the engine (shown in Figure15.2.15), ultimately it will increase the engine thrust as per Fn = vj - vi). Figure15.2.15: Effect of air pressure on thrust. Altitude Effect: The atmosphere that surrounds the earth is a compressible fluid whose density is directly proportional to its pressure and inversely proportional to its temperature. Air under standard conditions, at sea level has a pressure of 14.69 psi, and this pressure decreases as the altitude increases. At 20,000 feet it is down to 6.75 psi, and at 30,000 feet it has dropped to 4.36 psi, and it continues to drop at a relatively uniform rate. As the pressure of the air decreases, so does its density. EASA – 147 Course Notes Figure 15.2.16 Effect of altitude on thrust Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 19 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC The temperature of the air also drops with an increase in altitude from its standard sea level condition, and as the temperature drops, the air becomes more dense. So, the thrust increases with the increase of altitude as the air density increases. The decrease in thrust caused by the dropping pressure is more than the increase caused by the lowering temperature. When we combine the loss from pressure drop with the gain from temperature drop, net effect is that thrust decreases with the increase of altitude. At about 36,000 feet, essentially the beginning of the tropopause, an interesting happens. The temperature of the air stops falling i.e. stabilizes at -69.7°F (-56.5°C) and remains constant while pressure continues to fall. As a result, thrust will drop off more rapidly above 36,000 feet because the thrust loss due to the air pressure drop will no longer be partially offset by the thrust gain due to temperature drop. Thus 36,000 feet is the optimum altitude for long-range cruising, because at this altitude, even though the engine’s thrust is reduced, the relationship between the thrust produced by the engine and the diminished drag on the aircraft is most favorable. Limitations As previously stated the engine is susceptible to temperature, speed, altitude and engine rotational speed. To prevent damage occurring to the engine, safety devices are fitted to prevent limitations being exceeded. To protect engine from the effects of temperature two governors are normally fitted. These are the power limiter, which limits the power output in low temperatures preventing over stressing of the engine, and the EGT limiter which protects the turbine assembly due to excessive temperatures. Engine speed is limited by the over-speed governor(s) to protect the rotational components from excessive centrifugal forces being imposed. It is also limited to prevent the rotating component’s tip speeds from achieving sonic velocities and the accompanying performance losses associated with them. EASA – 147 Course Notes 15.2.7. Engine Thrust Ratings The thrust an engine can generate has a range from idle to the maximum certified thrust (when the TLA is at the full forward stop). Within this range the main thrust levels for flight operation are designated. These designated thrust levels of an engine are called thrust ratings. The thrust level selected by the pilot, also between the designated ratings, is called thrust setting. Rating Structure of a Commercial Turbofan Engine: The highest thrust ratings on the rating scale are the two certified operating limits of the engine. These are maximum take-off thrust and maximum continuous thrust. On engines certified according to EASA CS-E or FAR33 the use of maximum take-off thrust is limited to duration of Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 20 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC 5 minutes per flight cycle. The thrust levels of the ratings below maximum continuous thrust are determined in accordance with the requirements of the aircraft. These ratings are:  Maximum climb thrust  Maximum cruise thrust  Maximum reverse thrust The maximum reverse rating is mainly an N1 speed limit. In Figure 15.2.17 the ratings are shown on a scale with idle ratings at the lower end of the scale. Several different idle ratings are used on turbofan engines. They are no thrust ratings because the idle ratings are shaft speed related. In the idle range the control system uses the HP shaft speed as the control parameter. One rating all engines have in common is the minimum idle rating. At this rating the HP shaft runs at its lowest possible speed to minimize the generated thrust. This shaft speed is slightly higher than the minimum speed from which the engine can accelerate. The idle speed in general is not a constant value. It increases with decreasing air density. The minimum idle speed on FADEC-controlled engines is often a fixed value over a broad OAT range. For the acceleration from minimum idle speed to 95% of maximum take-off thrust a time limit is established in the certification standards. This limit is necessary to ensure the immediate availability of the thrust during a balked landing maneuver. The acceleration time most turbofan engines need from minimum idle to 95% of maximum take-off thrust is too long to comply with certification standards. To fulfill this requirement, a higher idle speed is used during flight, the so-called flight idle or high idle speed. EASA – 147 Course Notes Figure 15.2.17: Thrust ratings and idle speeds Module – 15 | Gas Turbine Engine Category| B1.1 Page |2- 21 Issue No. 01 | Revision No. 05 Date: 16 October 2019  Bangladesh Airlines Training Center BATC On ground the minimum idle speed is used. This ensures that the aircraft can be taxied without an excessive use of brakes. In the simplest system designs the switching to flight idle is done by the air/ground logic of the aircraft. Thus this higher idle speed is valid during the whole flight (idle control system of the B737-300/-400/-500). The higher flight idle thrust is a disadvantage during the descent of the aircraft in clean configuration. To solve this problem, more sophisticated systems also use the flap position as a signal for switching to flight idle. With such a system the flight idle is only used in the approach configuration. It is then called approach idle. Due to the fact that a FADEC system can process more inputs, such systems use some additional idle ratings. Their use depends on the condition of the related airframe systems. One of these ratings is Ps3 idle (but not limited to FADEC engines). At minimum idle during flight, the supply pressure for the aircraft pneumatic system can be too low. To keep the air pressure for the supply of the environmental control system high enough, the Ps3 must be held on a sufficient level. This is ensured by the use of Ps3 as the control parameter during idle setting while the pneumatic system is supplied by the engine. The EEC calculates the required Ps3 from the bleed air demand of the aircraft pneumatic system. The N2 speeds reached during Ps3 idle are lower than the approach idle values. The other additional idle rating used is the one influenced by the oil temperature as used on the CFM56-5A/-5B/-5C and the V250