MODULE 15. GAS TURBINE ENGINE

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European Aviation Safety Agency (EASA)

Part-66

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JAR 66 CATEGORY B1 MODULE 15 GAS TURBINE ENGINES

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1.1 WORK, POWER & ENERGY ... 1-1 1.1.1 Work ... 1-1 1.1.2 Power... 1-1 1.1.3 Energy ... 1-2 1.2 FORCE AND MOTION ... 1-3 1.2.1 Force... 1-3 1.2.2 Velocity ... 1-3 1.2.3 Acceleration ... 1-4 1.3 PRINCIPLES OF JET PROPULSION ... 1-4 1.3.1 Thrust Calculation. ... 1-4 1.4 GAS TURBINES ... 1-6 1.5 THE BRAYTON CYCLE ... 1-7 1.6 CHANGES IN TEMPERATURE, PRESSURE AND VELOCITY . 1-9 1.6.1 Temperature and Pressure ... 1-9 1.6.2 Velocity and Pressure ... 1-10 1.6.3 How The Changes are Obtained. ... 1-10 1.7 DUCTS AND NOZZLES ... 1-10 Continuity equation. ... 1-10 1.7.2 Incompressible fluid flow. ... 1-11 1.7.3 Bernoulli’s Theorem ... 1-11 1.7.4 Total energy. ... 1-12 1.8 CONTINUITY EQUATION AND BERNOULLI’S THEOREM ... 1-13 1.8.1 Incompressible fluid. ... 1-13 1.8.2 Gas Laws ... 1-15

1.9 SUBSONIC AIRFLOW THROUGH DIVERGENT AND CONVERGENT DUCTS 1-16

Divergent Duct ... 1-16 1.9.2 Convergent Duct ... 1-16 SONIC AIRFLOW THOUGH DIVERGENT AND CONVERGENT DUCTS 1-17 1.11 THE WORKING CYCLE ON A PRESSURE VOLUME DIAGRAM 1-18 1.12 ENGINE CONFIGURATIONS. ... 1-19

1.12.1 Reaction engines ... 1-19 1.12.2 Power Engines ... 1-21

2 ENGINE PERFORMANCE ... 2-1

2.1 METHOD OF CALCULATING THE THRUST FORCES ... 2-1 2.2 CALCULATING THE THRUST OF THE ENGINE ... 2-2 2.2.1 Comparison between thrust and horse-power ... 2-6 2.3 ENGINE THRUST IN FLIGHT ... 2-7 2.3.1 Effect of forward speed ... 2-9 2.3.2 Effect of afterburning on engine thrust... 2-11 2.3.3 Effect of altitude ... 2-11 2.3.4 Effect of temperature ... 2-13

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2.5 FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP 2-15 2.6 SPECIFIC FUEL CONSUMPTION ... 2-16

2.6.1 SPECIFIC FUEL CONSUMPTION – DEFINITION ... 2-16 2.7 FLAT RATING ... 2-16 2.8 PERFORMANCE RATINGS ... 2-16

3 INLET ... 3-1

3.1 INTRODUCTION ... 3-1 3.2 RAM COMPRESSION ... 3-1 3.2.1 Importance of Ram Compression ... 3-1 3.3 TYPES OF AIR INTAKES ... 3-2 3.3.1 PITOT INTAKES ... 3-2 3.3.2 DIVIDED ENTRANCE DUCT ... 3-3 3.4 IDEAL INTAKE CONDITIONS ... 3-4 3.5 INTAKE ANTI-ICING ... 3-5 3.5.1 Engine Hot Air Anti-icing ... 3-5 3.5.2 Engine Electrical Anti-icing ... 3-7 3.5.3 Oil Anti-ice ... 3-8

4 COMPRESSORS ... 4-1

4.1 COMPRESSORS GENERAL... 4-1 4.2 CENTRIFUGAL FLOW ... 4-1 4.2.1 Operation ... 4-3 4.3 THE AXIAL FLOW COMPRESSOR ... 4-5 Operation ... 4-6 4.4 COMPRESSOR STALL AND SURGE ... 4-13

4.4.1 Airflow Control System Principles ... 4-13 4.4.2 Compressor Characteristics ... 4-17

4.4.3 Effect of Temperature on the Operating Point of the Airflow Control System 4-18

4.5 AIR FLOW CONTROL SYSTEM – OPERATION ... 4-20

4.6 AEROFOIL THEORY AND THE AXIAL FLOW COMPRESSOR (CONTINUED) 4-25

4.6.1 Speed of Airflow Over Blades ... 4-25 4.6.2 Angle of Attack ... 4-25 Some Important Points about Angle of Attack ... 4-26 4.7 APPLICATION TO THE AXIAL FLOW COMPRESSOR ... 4-27 4.7.1 Compressor RPM... 4-27 4.7.2 Common Causes of Compressor Stall ... 4-27 4.7.3 Stagger Angle and End Bend ... 4-27 4.7.4 Recent innovations... 4-27 4.8 AIRFLOW CONTROL ... 4-29 4.9 AIR BLEED VALVES (SUMMARY) ... 4-29 4.10 VARIABLE INTAKE GUIDE VANES (SUMMARY) ... 4-29 4.11 MULTI-SPOOL COMPRESSORS (SUMMARY) ... 4-29

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4.12.1 Centrifugal ... 4-30 4.12.2 Axial Flow ... 4-30 5 COMBUSTION SECTION ... 5-1 5.1 INTRODUCTION ... 5-1 5.2 COMBUSTION PROCESS ... 5-1 5.3 FUEL SUPPLY ... 5-3 5.4 TYPES OF COMBUSTION CHAMBER ... 5-4 5.4.1 Multiple combustion chamber ... 5-4 5.4.2 Tubo-annular combustion chamber ... 5-6 (Also known as Can-annular or Cannular.) ... 5-6 5.4.3 Annular combustion chamber ... 5-7 5.4.4 Reverse Flow Combustion Chamber ... 5-9 5.5 COMBUSTION CHAMBER PERFORMANCE ... 5-10

5.5.1 Combustion intensity ... 5-10 5.6 COMBUSTION EFFICIENCY ... 5-11 5.7 COMBUSTION STABILITY ... 5-11 5.8 POLLUTION CONTROL ... 5-12 5.8.1 Introduction ... 5-12 5.8.2 Sources of Pollution ... 5-12 5.9 EMISSIONS ... 5-12 5.10 MATERIALS ... 5-14 6 TURBINE SECTION ... 6-1 6.1 INTRODUCTION ... 6-1 6.2 ENERGY TRANSFER FROM GAS FLOW TO TURBINE ... 6-5 6.3 CONSTRUCTION ... 6-8 6.3.1 Nozzle guide vanes ... 6-8 6.3.2 Turbine discs ... 6-9 6.3.3 Turbine blades ... 6-9 6.3.4 Dual alloy discs ... 6-11 6.4 COMPRESSOR-TURBINE MATCHING ... 6-11 6.5 MATERIALS ... 6-11 6.5.1 Nozzle guide vanes ... 6-11 6.5.2 Turbine discs ... 6-11 6.5.3 Turbine blades ... 6-12 6.6 DYNAMIC BALANCING PRINCIPLES ... 6-16 6.6.1 Introduction ... 6-16 6.6.2 Centrifugal Force ... 6-17 6.6.3 Causes of Unbalance ... 6-18 6.6.4 Objective of Balancing ... 6-20 6.6.5 Definition of Unbalance ... 6-20 6.6.6 Fan Balancing ... 6-23 7 EXHAUST ... 7-1 7.1 INTRODUCTION ... 7-1

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7.4 NOISE REDUCTION ... 7-8 7.4.1 Sources of Engine Noise ... 7-8 7.5 THRUST REVERSAL ... 7-18

7.5.1 Introduction ... 7-18 7.5.2 Requirement for Thrust Reversal ... 7-18 7.5.3 Layout and Operation of Typical Thrust Reversing Systems 7-19 7.5.4 Safety Features ... 7-22 CFM 56 Thrust Reverser for Boeing 737-300 ... 7-22

8 BEARINGS, SEALS AND GEARBOXES ... 8-1

8.1 BEARINGS ... 8-1 8.1.1 Introduction ... 8-1 8.1.2 Ball Bearings ... 8-1 8.1.3 Roller Bearings ... 8-1 8.1.4 Other types of bearings ... 8-1 8.2 BEARING CHAMBER OR SUMP ... 8-3 8.2.1 Lubrication ... 8-3 8.2.2 Sealing ... 8-3 8.2.3 Thread Seals ... 8-4 8.2.4 Carbon Seal ... 8-5 8.2.5 Spring Ring Seal ... 8-5 8.2.6 Hydraulic Seal ... 8-6 8.3 ACCESSORY DRIVE GEARBOXES ... 8-7 8.3.1 Introduction ... 8-7 8.3.2 Internal gearbox ... 8-7 8.3.3 Radial driveshaft ... 8-10 8.3.4 Direct drive ... 8-10 8.3.5 Gear train drive ... 8-10 8.3.6 Intermediate gearbox ... 8-10 8.3.7 External gearbox ... 8-11 8.3.8 Auxiliary gearbox ... 8-12 8.3.9 Construction and Materials ... 8-14

9 LUBRICANTS AND FUEL ... 9-1

9.1 GAS TURBINE FUEL PROPERTIES AND SPECIFICATION ... 9-1 9.2 FRACTIONAL DISTILLATION ... 9-1 9.3 PROPERTIES ... 9-3 9.3.1 Ease of Flow ... 9-3 9.3.2 Ease of Starting ... 9-3 9.3.3 Complete Combustion ... 9-3 9.3.4 Calorific Value ... 9-4 9.3.5 Corrosive Properties ... 9-4 9.3.6 Effects of By-Products of Combustion ... 9-5 9.3.7 Fire Hazards ... 9-5 9.3.8 Vapour Pressure ... 9-6 9.3.9 Fuel Boiling and Evaporation Losses ... 9-6 9.3.10 Methods of Reducing or Eliminating Fuel Losses ... 9-6 9.3.11 Fuel additives ... 9-8

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9.4.1 Viscosity ... 9-9 9.4.2 Hydro-Dynamics or Fluid Film Lubrication ... 9-9 9.4.3 Boundary Lubrication ... 9-10 9.5 LUBRICATING OILS ... 9-10 9.6 TURBINE OILS ... 9-11 9.6.1 First Generation Synthetic Oils ... 9-12 9.6.2 Second Generation Synthetic Oils ... 9-12 9.6.3 Third Generation Synthetic Oils... 9-12 9.6.4 Safety Precautions ... 9-13

10 LUBRICATION SYSTEMS ... 10-1

10.1 INTRODUCTION ... 10-1 10.2 BEARINGS ... 10-1 10.3 ENGINE LUBRICATION SYSTEMS ... 10-5 10.3.1 Pressure Relief Valve Re-circulatory System ... 10-5 10.3.2 Recirculatory Oil System – Full Flow Type ... 10-8 10.3.3 Advantages of Full Flow Lubrication ... 10-8 10.3.4 Expendable System ... 10-10 10.4 MAIN COMPONENTS ... 10-11 10.4.1 Oil Tank ... 10-11 10.4.2 Oil Pumps ... 10-12 10.4.3 oil cooling ... 10-14 10.4.4 Pressure Filter ... 10-15 10.4.5 Last Chance Filter ... 10-17 10.4.6 Scavenge Oil Strainers ... 10-17 10.4.7 Magnetic Chip Detector ... 10-18 10.4.8 De-aerator ... 10-18 10.4.9 Centrifugal Breather ... 10-19 Pressure Relief Valve ... 10-19 10.4.11 By-Pass Valve ... 10-20 10.5 INDICATIONS AND WARNINGS ... 10-21 10.5.1 Low Pressure Warning Lamp ... 10-21 10.5.2 Oil Pressure, temperature and quantity indication ... 10-21 10.6 OIL SEALS ... 10-21 10.7 SERVICING ... 10-21

11 ENGINE FUEL CONTROL SYSTEMS ... 11-1

11.1 INTRODUCTION ... 11-1 11.2 PURPOSE OF THE ENGINE FUEL SYSTEM ... 11-1 11.3 LAYOUT OF TYPICAL SYSTEM COMPONENTS ... 11-3 11.3.1 Aircraft Mounted Components ... 11-3 11.3.2 The Engine LP fuel system ... 11-3 11.3.3 The Engine HP Fuel System ... 11-3 11.4 FACTORS GOVERNING FUEL REQUIREMENTS ... 11-5 11.5 REQUIREMENTS OF THE ENGINE FUEL SYSTEM ... 11-5 11.6 ENGINE FUEL SYSTEM COMPONENTS ... 11-5

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11.7.2 Plunger-type Fuel Pump ... 11-6 11.7.3 Gear-Type Fuel Pump ... 11-7 11.8 FUEL FLOW CONTROL ... 11-7 11.8.1 Basic Flow Control System ... 11-8 11.9 HYDRO-MECHANICAL CONTROL UNITS ... 11-10

11.9.2 Barometric Controls ... 11-11 11.9.3 Proportional Flow Control. ... 11-13 11.9.4 Acceleration Control Units ... 11-14 11.10 ENGINE PROTECTION DEVICES ... 11-18 11.10.1 Top Temperature Limiter. ... 11-18 11.10.2 Power Limiter. ... 11-18 11.10.3 Overspeed Governor. ... 11-19 BURNERS ... 11-21 11.11.1 Atomiser Burners ... 11-21 11.11.2 Vaporising Burners... 11-26 11.11.3 Combustion and Airflow ... 11-28 11.12 ELECTRONIC ENGINE CONTROL SYSTEMS ... 11-30 11.12.1 Supervisory Electronic Engine Control ... 11-30 11.12.2 FUEL CONTROL ... 11-32 11.12.3 General ... 11-32 11.12.4 Full-Authority Digital Electronic Control (FADEC) ... 11-37

12 AIR SYSTEMS ... 12-1

12.1 INTRODUCTION ... 12-1 12.2 INTERNAL COOLING AIRFLOW ... 12-2 12.2.1 Low Pressure Air ... 12-2 12.2.2 Intermediate Pressure Air... 12-2 12.2.3 High Pressure Air ... 12-2 12.2.4 Differential Pressure Seals ... 12-3 12.3 SEALING ... 12-3 12.4 COOLING. ... 12-5 12.5 TURBINE CASE COOLING – DESCRIPTION AND OPERATION 12-9

12.5.1 Passive Clearance Control System. Figure 12.7. ... 12-9 12.5.2 Active Clearance Control System. Figure 12.8. ... 12-10 12.5.3 Low Pressure Turbine Clearance Control Valve ... 12-11 12.6 EXTERNAL COOLING ... 12-13 12.6.1 External skin of aero-engine. ... 12-13 12.6.2 Cooling of Accessories ... 12-14 12.7 HP AIR FOR AIRCRAFT SERVICES. ... 12-15 External Air Tappings ... 12-15 12.8 ANTI-ICING SYSTEMS ... 12-18

13 STARTING AND IGNITION SYSTEMS ... 13-1

13.1 BASIC PRINCIPLES OF GAS TURBINE ENGINE STARTING SYSTEMS 13-1 13.1.1 Purpose ... 13-1

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13.2.2 Electric Starter/Generator... 13-3 13.2.3 Safety Interlocks ... 13-4 13.2.4 Air Turbo Starters ... 13-5 13.3 A300 STARTING SYSTEM ... 13-8 13.3.1 GE 6-50 Starting Procedure ... 13-8 13.4 IGNITION SYSTEMS ... 13-12

13.4.1 High Energy Ignition Unit... 13-12 13.4.2 Igniter Plug ... 13-14 13.4.3 Servicing the Ignition System ... 13-14

14 ENGINE INDICATION SYSTEMS ... 14-1

14.1 INTRODUCTION. ... 14-1 14.2 ENGINE SPEED INDICATORS. ... 14-3 14.3 THRUST INDICATION... 14-7 14.3.1 Engine Pressure Ratio.EPR. ... 14-7 14.3.2 Torque indication ... 14-9 14.3.3 Phase comparison Torquemeter ... 14-12 14.4 EXHAUST GAS TEMPERATURE ... 14-13 14.4.1 Thermocouples ... 14-13 14.5 FUEL FLOW METERING ... 14-17 14.6 OIL ... 14-20 14.6.1 The Oil Pressure Indicator ... 14-20 14.6.2 Oil pressure warning light ... 14-21 Oil Temperature. ... 14-22 14.6.4 Oil Quantity ... 14-23 14.7 VIBRATION ... 14-24 14.8 WARNING LIGHTS ... 14-24 15 THRUST AUGMENTATION ... 15-1 15.1 INTRODUCTION ... 15-1 15.2 WATER INJECTION ... 15-1 15.2.1 Effects on Engine Power ... 15-1 15.2.2 Methods of Applying Water/Methanol ... 15-1 15.2.3 Compressor Intake Injection (Turbo Prop) ... 15-2 15.2.4 Combustion Chamber Injection System ... 15-4 15.3 RE-HEAT (AFTER BURNING) ... 15-6 15.3.1 Purpose ... 15-6 15.3.2 Revision of Thrust ... 15-6 15.3.3 Re-heat and By-pass Engines ... 15-6 15.3.4 The Advantage of Re-heat ... 15-6 15.3.5 The disadvantages of Re-heat ... 15-7 15.3.6 Propelling Nozzles ... 15-7 15.3.7 Re-heat Nozzles ... 15-8 15.3.8 The Re-heat Jet Pipe ... 15-10

16 TURBO-PROP ENGINES ... 16-1

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16.2.2 Free Power Turbine ... 16-2 16.2.3 Compounded Engine ... 16-3 16.3 REDUCTION GEARING ... 16-3 16.3.1 Simple Spur ‘Epicyclic’ ... 16-4 16.3.2 Compound Spur Epicyclic ... 16-6 16.3.3 Gear Train/Epicyclic ... 16-7 16.4 TURBO-PROP PERFORMANCE ... 16-7 16.5 TURBO-PROP ENGINE CONTROL ... 16-7 16.5.1 Integrated Control of RPM and Fuel Flow ... 16-8 16.5.2 Direct Control of Fuel Flow ... 16-8 16.5.3 Direct Control of Blade Angle (Beta Control) ... 16-8 16.6 ENGINE AND PROPELLER CONTROLS... 16-9 16.7 CONTROL OUTSIDE NORMAL FLIGHT RANGE ... 16-9 16.8 PROPELLER CONTROL ... 16-9 16.8.1 Constant Speed Unit ... 16-10 16.8.2 Manual and Automatic Feathering Controls ... 16-10 16.8.3 Fixed and Removable Stops ... 16-15 16.9 OVERSPEED SAFETY DEVICES ... 16-16

17 TURBOSHAFT ENGINES ... 17-1

17.1 INTRODUCTION. ... 17-1 17.2 FUEL CONTROL SYSTEM ... 17-4 17.3 ARRANGEMENTS ... 17-6 17.4 DRIVE SYSTEMS ... 17-10 17.5 COUPLINGS ... 17-13

18 AUXILLIARY POWER UNITS ... 18-1

18.1 INTRODUCTION ... 18-1 18.2 GENERAL ARRANGEMENTS AND CONFIGURATION ... 18-3 18.2.1 Inlet Duct Arrangement ... 18-7 18.2.2 Exhaust Duct Arrangement ... 18-9 18.3 THE APU ENGINE ... 18-10 18.4 FUEL CONTROL ... 18-12 Mechanical Fuel Control ... 18-12 18.4.2 Speed Control ... 18-18 18.4.3 Mechanical Fuel Control Unit Operation ... 18-19 18.4.4 Electronic APU Fuel Control ... 18-20 18.4.5 Electro/mechanical Fuel Control (FIGURE 18.26) ... 18-21 18.5 APU OIL SYSTEM ... 18-23 18.6 APU BLEED AIR SYSTEMS ... 18-25 18.6.1 direct from engine compressor ... 18-25 18.6.2 SEPARATE LOAD COMPRESSOR ... 18-27 18.7 BAY COOLING ... 18-28 18.7.1 Ram Air Cooling ... 18-28 18.7.2 Fan Air Cooling ... 18-28

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19 POWERPLANT INSTALLATION ... 19-1 19.1 NACELLES OR PODS... 19-1 19.1.1 Cowlings ... 19-1 19.1.2 Firewalls ... 19-4 19.1.3 Cooling ... 19-6 19.1.4 Acoustic Linings ... 19-8 19.1.5 Abradable Linings ... 19-11 19.2 ENGINE MOUNTS ... 19-12 19.2.1 Wing Pylon Mounted Engine (Turbofan) ... 19-12 19.2.2 Wing Mounted Engine (Turboprop) ... 19-14 19.2.3 Rear Fuselage Engine Turbofan.(Figure 19.14/15.) ... 19-16 19.3 ENGINE DRAINS. ... 19-18 19.3.1 Controlled Drains ... 19-18 19.3.2 Uncontrolled Drains ... 19-20 19.4 ENGINE CONTROLS ... 19-22 19.4.1 Throttle Control Mechanical ... 19-22 19.4.2 Turbofan Engine Controls. ... 19-22 19.4.3 Turboprop Engine Controls ... 19-24 19.5 ENGINE BUILD UNIT ... 19-29 19.5.1 Turbofan Engine ... 19-29 19.6 FIRE PREVENTION – BAYS OR ZONES... 19-38 19.7 INSTALLING AND REMOVING ENGINES. ... 19-40 19.7.1 Removal ... 19-40 19.7.2 Fitting ... 19-48 19.7.3 Turbo Prop Engine Removal/Fit. ... 19-48 19.7.4 Flight Transit ... 19-48

20 FIRE PROTECTION SYSTEMS ... 20-1

20.1 FIRE DETECTORS ... 20-1 20.2 FIRE WIRE SYSTEMS ... 20-3 20.2.1 Resistance Type ... 20-3 20.2.2 Capacitance Type ... 20-3 20.2.3 Gas Operation Fire Wire ... 20-4 20.2.4 Single Loop ... 20-5 20.2.5 Dual Loop ... 20-5 Dual Loop Systems ... 20-6 20.3 FIRE AND LOOP FAULT INDICATION (E.C.A.M.) ... 20-8 20.4 FIRE SUPPRESSION ... 20-9 20.4.1 Types of Fire Suppression System ... 20-11 One Shot System ... 20-11 20.4.2 Two Shot System (shared extinguishers) ... 20-12 20.4.3 Two Shot System (Single Head extinguishers) ... 20-14 20.5 EXTINGUISHERS ... 20-16 20.5.1 Operating Head ... 20-17 20.5.2 Safety Discharge ... 20-17 20.5.3 Discharge Tube Configuration ... 20-18 20.5.4 Operating Time ... 20-19

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20.6 INDICATIONS OF FIRE DETECTION ... 20-20 20.6.1 Fire T Handle ... 20-20 20.6.2 Fire Bell ... 20-20 20.6.3 Fire Detection Test ... 20-22 20.7 DISCHARGE INDICATORS ... 20-23 20.7.1 Mechanical Indicators ... 20-23 20.7.2 Electrical Indicators ... 20-23 20.8 CARTRIDGES OR SQUIBS ... 20-24 20.8.1 Life Control of Squibs ... 20-24 INTENTIONALLY BLANK ... 20-26

21 ENGINE MONITORING AND GROUND OPERATIONS. ... 21-1

21.1 PROCEDURES FOR STARTING AND GROUND RUNNING. .... 21-1 21.2 STARTING ... 21-3 21.3 UNSATISFACTORY STARTS ... 21-7 21.4 ENGINE STOPPING. ... 21-8 21.5 ENGINE FIRES ... 21-9

21.6 INTERPRETATION OF ENGINE POWER OUTPUTS AND PARAMETERS. 21-10 21.7 TREND MONITORING. ... 21-22

21.7.1 On Ground Monitoring ... 21-24 21.7.2 Air Washed Components ... 21-24 21.7.3 Oil Washed Components ... 21-32 21.7.4 Inspections ... 21-36

22 ENGINE STORAGE AND PRESERVATION. ... 22-1

22.1 STORAGE AND TRANSIT ... 22-1 22.1.1 Fuel System Inhibiting. ... 22-1 22.1.2 Packing. ... 22-2 22.1.3 Storage. ... 22-3

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1 FUNDAMENTALS

1.1 WORK, POWER & ENERGY

Work, power, and energy are all interrelated. Work is the amount of movement a given force causes; energy is the ability to do work, and power is the rate of doing work.

1.1.1 WORK

In its technical sense, work is the product of force and distance, and work is done only when a force causes movement. We can see this by the formula:

Work = Force x Distance

We normally measure distance in feet or inches, and force in pounds or ounces. This allows us to measure work in foot-pounds or inch-ounces.

Example:

To find the amount of work done when a 500 pound load is lifted for a distance of 6 feet, we can use the formula:

Work = Force x Distance = 500 X 6

= 3,000 foot-pounds

1.1.2 POWER

The rate of doing work is called power, and it is defined as the work done in unit time. As a formula, this would be:

power = work done time taken

Power is expressed in several different units, such as the watt, ergs per second, and foot-pounds per second. The most common unit of power in general use in the United States is the horsepower. One horsepower (hp) is equal to 550 ft-lb’s or 33000 ft-1b/min. In the metric system the unit of power is the watt (W) or the kilowatt (kW). One hp is equal to 746 watts; and 1 kW = 1.34 hp.

Example:

To compute the power necessary to raise an elevator containing 10 persons a distance of 100 ft in 5 s (assuming the loaded elevator weighs 2500 lb), proceed as follows:

Power = work done = 2500 x 100 = 50,000 ft-lb’s/sec Time taken 5

Since 1hp = 550 ft-lb’s/sec then required hp = 50,000 550

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1.1.3 ENERGY

The term energy may be defined as the capacity for doing work. There are two forms of energy: potential energy and kinetic energy.

1.1.3.1 POTENTIAL ENERGY

Potential energy is the stored energy possessed by a system, because of the relative positions of the components of that system. If work done raises an object to a certain height, energy will be stored in that object in the form of the gravitational force. This energy, waiting to be released is called potential energy. The amount of potential energy a system possesses is equal to the work done on the system previously.

Potential energy can be found in forms other than weights and height. Electrically charged components contain potential (electrical) energy because of their position within an electric field. An explosive substance has chemical potential energy that is released in the form of light, heat and kinetic energy, when detonated.

Example :

A weight of 50 pounds is raised 5 feet. Using the formula: Potential Energy = Force x Distance

= 50 x 5 = 250 ft-lb’s.

Note: That energy is expressed in the same units as those used for work and in all cases energy is the product of force x distance.

1.1.3.2 KINETIC ENERGY

Kinetic energy is the energy possessed by an object, resulting from the motion of that object. The magnitude of that energy depends on both the mass and speed of the object. This is demonstrated by the simple equation:

Energy =½mv2 _{or w v}2

2g

where m = mass, v = velocity (in feet or metres per second), w = weight, g = gravity (32 ft/sec2_{or 9.81m/sec}2_).

All forms of energy convert into other forms by appropriate processes. In this process of transformation, either form of energy can be lost or gained but the total energy must remain the same.

Example:

A weight of 50lbs dropped from a height of 5 ft has kinetic energy of KE = 50 x 25

2 x 32 = 19.53 ft-lb’s

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1.2 FORCE AND MOTION

1.2.1 FORCE

Force may be defined as a push or a pull upon an object. In the English system the pound (1b) is used to express the value of a force. For example, we say that a force of 30 lb is acting upon a hydraulic piston.

A unit of force in the metric system is the newton (N). The newton is the force required to accelerate a mass of 1 kilogram (kg) 1 meter per second per second (m/s2_).

The dyne (dyn) is also employed in the metric system as a unit of force. One dyne is the force required to accelerate a mass of 1g 1 centimetre per second per second (cm/s2_{). One newton is equal to 100,000 dynes (0.225 Ib).}

1.2.2 VELOCITY

It is common to find people confusing the terms velocity and speed when describing how fast an object is moving. The difference is that speed is a scalar quantity, whilst the term velocity refers to both speed and direction of an object. The full definition of velocity is that it is the rate at which its position changes, over time, and the direction of the change.

The simple diagram below shows how an aircraft, which flies the irregular path from 'A' to 'B' in an hour, (a speed of 350 mph), has an actual velocity of 200 mph in an East-Northeast direction.

A

B

C

N

200 Ml (322 Km)

350 Ml (563 Km)

Path of Aircraft

Diagram Showing Difference Between Velocity and Speed Figure 1.1.

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1.2.3 ACCELERATION

This term describes the rate at which velocity changes. If an object increases in speed, it has positive acceleration; if it decreases in speed, it has negative acceleration. A reference to Newton's Second law of Motion will explain the principles of acceleration. Acceleration can be in a straight line, which is referred to a linear acceleration and it can apply to rotating objects whose speed of rotation is increasing, (or decreasing), when it is called angular acceleration.

1.3 PRINCIPLES OF JET PROPULSION

Newton’s Laws of Motion. To understand the basic principles of jet propulsion it is necessary to understand the practical application of Sir Isaac Newton's Laws of Motion. There are three laws.

1. The First Law States. A mass will remain stationary until acted upon by a force. If the mass is already moving at a constant speed in a straight line, it will. continue to move at that constant speed in a straight line until acted upon by a force.

2. The Second Law States. When a force acts on a mass the mass will accelerate in the direction in which the force acts.

3. The Third Law States. To every action there is an equal and opposite reaction. The function of any propeller or gas turbine engine is to produce THRUST, (or a propulsion force), by accelerating a mass of air or gas rearwards. If we apply Newton's Laws of Motion to aircraft propulsion it can be said that:-

 a FORCE must be applied in order to accelerate the mass of air or gas: first law,

the acceleration of the mass is proportional to the force applied: second law,

 there must be an equal and opposite reaction, in our case this is THRUST, a forward acting force: third law.

1.3.1 THRUST CALCULATION.

The amount of thrust produced depends upon two things:- the MASS of air which is moved rearwards in a given time,

 the ACCELERATION imparted to the air.

It can be expressed as:- Thrust = Mass x Acceleration The MASS is defined as “the quantity of matter in a body". It is expressed as W

g

Where:- W = the weight of the body (in lb’s or newtons) and

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If we let: -

V2 = the air velocity at exit (in ft/sec/sec or 9.81m/sec2)

and

V1 = the air velocity at inlet (in ft/sec/sec or 9.81m/sec2)

It may be expressed as V2 – V1

Taking these expressions for Mass and Acceleration, the thrust produced by an engine or propeller can be calculated from the following formula:-

THRUST =



V2-V1



g W

Example 1.

The airflow through a propeller is 256 lbs/sec, Inlet velocity 0 ft/sec, outlet velocity 700 ft/sec.

Thrust developed will be: THRUST =



V2-V1



g W THRUST = 256 x (700 – 0) 32 = 5600 lbs Example 2.

The mass airflow through a gas turbine engine is 128lbs/sec, inlet velocity is 0 ft/sec, outlet velocity is 1400 ft/sec. Using the formula :

THRUST = 128 x (1400 – 0) 32

= 5600lbs

By comparing both examples, you can see that the gas turbine produced the same thrust as the propeller by giving a greater acceleration to a smaller mass. It can be said that a propeller accelerates a large mass slowly whilst the gas turbine produces the same thrust by giving a greater acceleration to a smaller mass. Note that in both of the examples the inlet velocity was zero ft/sec. The aircraft was stationary so the thrust produced is referred to as STATIC THRUST.

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1.4 GAS TURBINES

A gas turbine engine is essentially a heat engine using a mass of air as a working fluid to provide thrust. To achieve this, the mass of air passing through the engine has to be accelerated, which means that the velocity, (or kinetic energy), of the air is increased. To obtain this increase, the pressure energy is first of all increased, followed by the addition of heat energy, before final conversion back to kinetic

energy in the form of a high velocity jet efflux.

The simplest form of gas turbine engine is the turbojet engine, which has three major parts; the compressor, the combustion section and the turbine. A shaft connects the compressor and the turbine to form a single, rotating unit. These engines produce thrust in the manner described in the Brayton Cycle.

The simplest turbojet engine is the unit shown below with a single centrifugal(Double Entry)compressor and a single stage turbine. This type of engine can still be found in certain special installations but generally, they have been superseded by engines with axial compressors and multiple stage turbines. The advantages and disadvantages of the two types of compressor will be discussed in depth later in this module

Simple Centrifugal Gas Turbine (Derwent) Figure 1.2.

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1.5 THE BRAYTON CYCLE

The working cycle of the gas turbine engine is similar to that of the four-stroke piston engine. There is induction, compression, ignition and exhaust in both cases, although the process is continuous in a gas turbine. Also, the combustion in a piston engine occurs at a constant volume, whilst in a gas turbine engine it occurs at a constant pressure.

The cycle, upon which the gas turbine engine functions, in its simplest form, is the Brayton cycle, which is represented by the pressure/volume diagram, shown in figure 1.4.

The Working Cycle. Figure 1.3.

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• The air entering the engine is compressed.

• Heat is added to the air by burning fuel at a constant pressure, thereby considerably increasing the volume of the resulting gas.

• The gases resulting from combustion expand through the turbine, which converts some of the energy in the expanding gases into mechanical energy to drive the compressor.

• The remainder of the expanding gases are propelled through the turbine and jet pipe back to the atmosphere where they provide the propulsive jet.

There are three main stages in the engine working cycle during which the changes discussed occur:

• During compression. Work is done on the air. This increases the pressure and temperature and decreases the volume of air.

• During combustion. Fuel is added to the air and then burnt. This increases the temperature and volume of the gas, whilst the pressure remains almost constant (the latter being arranged by design in a gas turbine engine).

• During expansion. Energy is taken from the gas stream to drive the compressor via the turbine; this decreases the temperature and pressure, whilst the volume increases. The rapidly expanding gases are propelled through the turbine and jet pipe to give a final momentum that is much greater than the initial momentum; it is this change in momentum which produces the propulsive jet.

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1.6 CHANGES IN TEMPERATURE, PRESSURE AND VELOCITY .

1.6.1 TEMPERATURE AND PRESSURE

The changes in temperature and pressure of the gases through a gas turbine engine are illustrated in Figure 1.5 The efficiency with which these changes are made will determine to what extent the desired relations between pressure, temperature and velocity are obtained. The more efficient the compressor, the higher is the pressure generated for a given work input - i.e. for a given temperature rise of the gas. Conversely, the more efficiently the turbine uses the expanding gas, the greater is the output of work for a given temperature drop in gas.

Gas Flow Through an Engine Figure 1.5

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1.6.2 VELOCITY AND PRESSURE

During the passage of the air (gas) through the engine, aerodynamic and energy requirements demand changes in its velocity and pressure. For example, during compression a rise in the pressure of the air is required with no increase in its velocity. After the air has been heated and its internal energy increased by combustion, an increase in the velocity of the gases is necessary to cause the turbine to rotate. Also at the propelling nozzle, a high velocity is required, for it is the change in momentum of the air that provides the thrust on the aircraft. Local decelerations of gas flow are also required - for example, in the combustion chambers to provide a low velocity zone for the flame.

1.6.3 HOW THE CHANGESARE OBTAINED.

The various changes in temperature, pressure and velocity are effected by means of the ducts through which the air (gas) passes on its way through the engine. When a conversion from kinetic energy to pressure energy is required, the ducts are divergent in shape. Conversely, when it is required to convert the energy stored in the combustion gases to velocity, a convergent nozzle is used. The design of the passages and nozzles is of great importance, for upon their good design depends the efficiency with which the energy changes are effected. Any interference with the smooth flow of gases creates a loss in efficiency and could result in component failure because of vibration caused by eddies or turbulence of the gas flow.

1.7 DUCTS AND NOZZLES

1.7.1 CONTINUITY EQUATION.

If we consider the machine to be an open-ended duct (Fig 1.6.), we find that the mass flow per second will depend on the density of the fluid and the volume flowing per sec:

Now volume flow = Area of duct x distance travelled (L)

Open Ended Duct to Illustrate Continuity Equation Figure 1.6.

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But the distance travelled per second = Velocity. Therefore, Mass flow = density x area x velocity.

This is known as the ‘continuity equation’ and it is true for any steady flow system regardless of changes in the cross-sectional area of the duct.

1.7.2 INCOMPRESSIBLE FLUID FLOW.

Now consider an incompressible fluid as it flows through the duct system shown in the fig. 1.7. We know that the mass flow is of a constant value and, naturally, as the fluid enters the larger cross sectional area it will take up the new shape and the initial volume will now occupy less length in the duct. Therefore, in a given time, less distance is travelled and the velocity is reduced.

Thus we conclude that if the mass flow is to remain constant, as it must, an increase in duct area must be accompanied by a reduction in flow velocity, and a decrease in duct area must bring about an increase in velocity; we can express this action as – velocity varies inversely with changes in duct area.

1.7.3 BERNOULLI’S THEOREM

This theorem can be related to the relationship between pressure and velocity existing in the air flowing through a duct, such as a jet engine. The theorem states that the total energy per unit mass is constant for a fluid moving inside a duct and that total energy consists mainly of pressure energy and kinetic energy:

 Pressure energy.

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

Duct System Figure 1.7.

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 Kinetic energy.

This kind of energy is more often called ‘dynamic pressure’ and this term is used to define the extra pressure created by the movement of the medium. Dynamic pressure is proportional to ½ mass x velocity 2_{(ie. ½mv}2_).

When the medium (gas or fluid) is moving, the total energy = static pressure + dynamic pressure.

Consider a duct which is filled with an incompressible fluid and pressurised from one end by an external force (Fig 1.8.). The other end of the duct is sealed by a valve, which can be opened or closed, and a pressure gauge is fitted into the wall of the duct to indicate the static pressure (PS). With the valve closed, static pressure and total energy are the same. However, when the valve is opened to allow a fluid flow, the circumstances changes and, although the total energy must remain the same, it now consists of static pressure + dynamic pressure. As the velocity V increases, so dynamic pressure increases and the static pressure is reduced.

1.7.4 TOTAL ENERGY.

Total energy can be measured as a ram pressure and is usually called the ‘total head’ or pitot pressure (PT). It is measured by placing a ram tube in the fluid flow. The ram tube must be parallel to the flow with its open end facing the flow. A gauge connected into such a tube always records the total head (pitot) pressure regardless of the rate of flow, refer to Fig 1.9.

Duct with Flow Control Valve Figure 1.8.

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In a situation where there is a no fluid flow, the static pressure (PS) gauge, and the total head pressure (PT) gauge will show the same value, but when there is a fluid flow, the total pressure reading remains the same although the static pressure drops.

1.8 CONTINUITY EQUATION AND BERNOULLI’S THEOREM

1.8.1 INCOMPRESSIBLE FLUID.

The combined effect of the continuity equation and Bernoulli’s theorem produces the effects shown, when a steady flow of incompressible fluid flows through a duct of varying cross sectional area (Fig 1.10.).

The effects of a steady flow of incompressible fluid flows through a duct of varying

Illustration of Pitot and Static Pressures Figure 1.9.

Duct of Varying Cross Sectional Area Figure 1.10.

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

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

1.8.1.1 Compressibility Fluid (Atmosphere).

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

The air entering the duct at section A (Fig 1.11), consists of air at pressure (P1) and velocity (V1); then as the air enters the increased area of the duct at B it will spread out to fill the increased area and this will cause the air flow to slow down (continuity equation) and give a change in velocity to V2. The static pressure of the air will increase (Bernoulli’s theorem) to become P2 in the wider section of the duct and, because air is compressible, the air density will increase as it is compresses by the rise in pressure in section B of the duct.

Airflow Through a Duct Section Figure 1.11.

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1.8.1.2 Diffuser action.

The flare, which increases the area of the duct, is known as a diffuser (Fig 1.12.)and its shape determines the rate of compression and the amount by which the air is compressed. For best results, the airflow must remain smooth and, because of this, a most important design feature is the angle of divergence. When air is compressed by this process it is called subsonic diffusion and it is a principle that is used extensively in jet engine design.

1.8.2 GAS LAWS

In addition to the preceding information, the following gas laws are closely related to the function of a gas turbine engine:

 Boyle’s Law. This law is related to temperature and pressure of a gas. It

states that if the temperature T remains constant, the volume V of a given mass varies inversely as the pressure P exerted upon it (ie. PV = Constant).

 Charles’ Law. This law states that the volume V of a given mass of gas

increases by 1/273 of its volume at 0°C for a rise of 1°C when the pressure P of the gas is kept constant. These laws are now combined in what is called the ideal gas law. It gives the relationship:

PV = RT where: P = pressure V = volume R = a constant T = absolute temperature in K Diffuser Section Figure 1.12.

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1.9 SUBSONIC AIRFLOW THROUGH DIVERGENT AND CONVERGENT DUCTS

1.9.1 DIVERGENT DUCT

A divergent duct widens out as the airflow progresses through it. At subsonic speeds the effect of this kind of duct is to decrease the velocity and increase the pressure and temperature of the air passing through it.

1.9.2 CONVERGENT DUCT

A convergent duct is such that the space inside reduces as the airflow progresses through it. At subsonic speeds the effect of this kind of duct is to increase the velocity and decreases the pressure and temperature of the air passing through it.

Divergent Duct. Figure 1.13.

Convergent Duct. Figure 1.14.

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1.10 SONIC AIRFLOW THOUGH DIVERGENT AND CONVERGENT DUCTS

When a flow of fluid (i.e. gas) flows at sonic speed through a convergent duct a shock wave forms at the exit area of the duct - The exit area is said to be choked. The shock wave forms a restriction to the fluid and pressure will increase,

temperature will increase and velocity will decrease.

When a gas flow reaches sonic velocity in a convergent duct the nozzle will choke and the pressure will increase. To prevent a pressure rise that would eventually prevent a 'fluid' flow and completely choke the duct a divergent section is added making the duct convergent/divergent (Con/DI). The pressure of gas released into the divergent section of the nozzle causes the velocity of the 'fluid' to increase, pressure to decrease, and therefore temperature to decrease. Gas pressure acts on the walls of the divergent section, this pressure gives additional thrust that is known as pressure thrust.

Airflow Through a Con-Di Nozzle or Venturi. Figure 1.15.

A Con-Di Nozzle Figure 1.14.

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1.11 THE WORKING CYCLE ON A PRESSURE VOLUME DIAGRAM

Air is drawn from the atmosphere (Ambient Air) into the compressor. The

compressor raises the pressure of the air (A to B) on diagram. If the pressure of the air is increased the volume is decreased. The air passes to the combustion system and heat is added by burning fuel with a proportion of the air. From the diagram (B to C) it is seen that combustion takes place at constant pressure so the gas turbine working cycle is known as the constant pressure cycle. In the

combustion system the air expands rearwards and the volume of the gas increases and the gas kinetic energy increases. The gas flow passes to the turbine section to drive the turbine (s), energy is extracted and the pressure decreases. The gas passes via an exhaust unit to the propelling nozzle which forms a convergent duct. The velocity of the gas increases. The reaction to the high velocity jet produces thrust (C to D on diagram).

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1.12 ENGINE CONFIGURATIONS.

There are two main types of gas turbine engines:

 Reaction engines, which derive their thrust by jet reaction

 Power engines, which provide a mechanical output to drive another device.

1.12.1 REACTION ENGINES

These can be divided into several categories.

a. Turbojet engines. The turbojet was the first type of jet engine developed. In this engine all the air passes through the core engine (i.e. the compressor, combustor and turbine). The engine may be single shaft as in the Avon engine, or twin shafted as in the Olympus 593 fitted to Concorde.

These engines are noisy and are not the most fuel efficient for normal use, however for high altitude high speed flight they are in a class of their own.

Turbo jet Engines. Figure 1.17.

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b. Low and Medium By-pass or turbofan engines. These engines will have two or three shafts. The Low Pressure (LP) shaft drives a larger diameter compressor. Some of the air produced by-passes the core engine (hence the name) and is used to provide thrust. The core airflow provides power for the compressors and thrust. These engine are quieter than turbojets and more fuel efficient. The Spey and Tay engines fall into this category.

The pass ratio is determined by the ratio of the air in flowing through the by-pass to the air by-passing through the core of the engine. Low by-by-pass less than 2:1, medium by-pass 2:1 to 4:1, high by pass greater than 5:1.

c. High by-pass turbofan engines. These engines have very large fans driven by a relatively small core engine. Often the fan is geared to run at a lower speed than the LP turbine, which gives the turbine mechanical advantage and also allows it to run at higher speed where it is more efficient. The ALF 502, RB211 and the Trent engines are all high by-pass

High by-pass engines are very fuel efficient, powerful and quiet. These engines have a very large diameter which does give drag problems, and are not suitable for high speed flight as the blade tips will suffer compressibility problems as they approach the speed of sound.

Low By-pass Twin Spool Engine (Spey) Figure 1.17.

A Three Spool High By-pass Engine (RB211) Figure 1.18.

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1.12.2 POWER ENGINES

Power producing engines come in two main forms Turboprop and turboshaft. a. Turboprop Engines. Turboprop engines extract most of the energy from the

gas stream and convert it into rotational energy to drive a propeller. The engines are either single or twin shaft and may be direct drive where the LP or main shaft drive the propeller through a gearbox, or they may have a separate power turbine to drive the propeller. Turboprop engines differ from high by-pass turbofans in that the propeller does not have an intake to slow and prepare the air before passing through it. The propeller therefore has to meet the demands of airspeed etc. Examples of turboprops are the Dart, PW125 and Tyne engines.

Turboprop Engines Figure 1.19.

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b. Turboshaft Engines. These engines are used in helicopters. They share many of the attributes of turboprop engines, but are usually smaller. They do not have propeller control systems built into the engine and usually do not have many accessories attached such as generators etc. as these are driven by the main rotor gearbox. Modern turboshaft and turbo prop engines run at constant speed which tends to prolong the life of the engine and also means that they are more efficient as the engine can run at its optimum speed all the time.

There are other types of engine such as ram jets, pulse jets, turbo-ram jet and turbo - rockets, but none of these are used commercially if at all.

Turboshaft Engine with Free power Turbine. (Gem) Figure 1.20.

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2 ENGINE PERFORMANCE

2.1 METHOD OF CALCULATING THE THRUST FORCES

The thrust forces or gas loads can be calculated for the engine, or for any flow section of the engine, provided that the areas, pressures, velocities and mass flow are known for both the inlet and outlet of the particular flow section.

The distribution of thrust forces shown in Fig 2.1. can be calculated by considering each component in turn and applying some simple calculations. The thrust produced by the engine is mainly the product of the mass of air passing through the engine and the velocity increase imparted to it (ie. Newtons Second Law of Motion), however the pressure difference between the inlet to and the outlet from the particular flow section will have an effect on the overall thrust of the engine and must be included in the calculation.

To calculate the resultant thrust for a particular flow section it is necessary to calculate the total thrust at both inlet and outlet, the resultant thrust being the difference between the two values obtained.

Thrust Distribution of a Typical Single Spool Axial Flow Engine. Figure 2.1.

TOTAL THRUST 11158 lbs

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Calculation of the thrust is achieved using the following formula: Thrust = g Wv P A  J ) (

Where A = Area of flow section in sq. in. P = Pressure in lb. per sq. in. W = Mass flow in lb. per sec.

VJ = Velocity of flow in feet per sec.

g = Gravitational constant 32.2 ft. per sec. per sec.

2.2 CALCULATING THE THRUST OF THE ENGINE

When applying the above method to calculate the individual thrust loads on the various components it is assumed that the engine is static. The effect of aircraft forward speed on the engine thrust will be dealt with later. In the following calculations ‘g’ is taken to be 32 for convenience.

Compressor casing

To obtain the thrust on the compressor casing, it is necessary to calculate the conditions at the inlet to the compressor and the conditions at the outlet from the compressor. Since the pressure and the velocity at the inlet to the compressor are zero, it is only necessary to consider the force at the outlet from the compressor. Therefore, given that the compressor –

OUTLET Area (A) = 182 sq. in.

Pressure (P) = 94 lb. per sq. in. (gauge) Velocity (vj) = 406 ft. per sec.

Mass flow (W) = 153 lb. per sec. The thrust = (  ) 0 g Wv P A j = 0 32 406 153 ) 94 182 (    

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Total Thrust of the Compressor. Figure 2.2.

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Choked Nozzle

Considering the formula for thrust under “choked” nozzle conditions: Thrust = (PP0)A + g Wv_J Where: P = Pressure P = Ambient Pressure A = Area W = Mass Flow V = Velocity

It can be seen that the thrust can be further affected by a change in the mass flow rate of air through the engine and by a change in jet velocity. An increase in mass airflow may be obtained by using water injection to cool the air and increases in jet velocity by using after-burning.

Changes in ambient pressure and temperature considerably influence the thrust of the engine. This is because of the way they affect the air density and hence the mass of air entering the engine for a given engine rotational speed.

Thrust Correction - Turbojet

To enable the performance of similar engines to be compared when operating under different climatic conditions, or at different altitudes, correction factors must be applied to the calculations to return the observed values to those which would be found under I.S.A. conditions. For example, the thrust correction for a turbo-jet engine is:

Thrust (lb) (corrected) = thrust (lb) (observed) x

O

P

30

 Where P0 = atmospheric pressure in inches of mercury (in Hg)

(observed)

30 = I.S.A. standard sea level pressure (in Hg)

Shaft Horsepower Correction - Turboprop

The observed performance of the turbo-propeller engine is also corrected to I.S.A. conditions, but due to the rating being in s.h.p. and not in pounds of thrust the factors are different. For example, the correction for s.h.p. is:

S.h.p. (corrected) = s.h.p. (observed) O O T P     273 15 273 30

Where P0 = atmospheric pressure (in Hg) (observed)

T0 = atmospheric temperature in deg. C (observed)

30 = I.S.A. standard sea level pressure (in Hg) 273 + 15 = I.S.A. standard sea level temperature in deg. K

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Equivalent Shaft Horsepower (EHP)

In practice there is always a certain amount of jet thrust in the total output of the turbo-propeller engine and this must be added to the s.h.p. The correction for jet thrust is the same as that specified earlier.

To distinguish between these two aspects of the power output, it is usual to refer to them as s.h.p. and thrust horse-power (t.h.p.). The total equivalent horse-power is denoted by t.e.h.p. (sometimes e.h.p.) and is the s.h.p. plus the s.h.p. equivalent to the net jet thrust. For estimation purposes it is taken that, under sea-level static conditions, one s.h.p. is equivalent to approximately 2.6 lb. of jet thrust. Therefore: t.e.h.p. = s.h.p. 6 . 2 . lb thrust jet 

The ratio of jet thrust to shaft power is influenced by many factors. For instance, the higher the aircraft operating speed the larger may be the required proportion of total output in the form of jet thrust. Alternatively, an extra turbine stage may be required if more than a certain proportion of the total power is to be provided at the shaft. In general, turbo-propeller aircraft provide one pound of thrust for every 3.5 h.p. to 5 h.p.

2.2.1 COMPARISON BETWEEN THRUST AND HORSE-POWER

Because the turbo-jet engine is rated in thrust and the turbo-propeller engine in s.h.p., no direct comparison between the two can be made without a power conversion factor. However, since the turbo-propeller engine receives its thrust mainly from the propeller, a comparison can be made by converting the horse-power developed by the engine to thrust or the thrust developed by the turbo-jet engine to t.h.p.; that is, by converting work to force or force to work. For this purpose, it is necessary to take into account the speed of the aircraft.

t.h.p. is expressed as sec . 550ft per FV Where F = lb of thrust

V = aircraft speed (ft. per sec)

Since one horse-power is equal to 550ft.lb. per sec. and 550 ft. per sec. is equivalent to 375 miles per hour, it can be seen from the above formula that one lb. of thrust equals one t.h.p. at 375 m.p.h. It is also common to quote the speed in knots (nautical miles per hour); one knot is equal to 1.1515 m.p.h. or one pound of thrust is equal to one t.h.p. at 325 knots.

Thus if a turbo-jet engine produces 5,000 lb. of net thrust at an aircraft speed of 600 m.p.h. the t.h.p. would be

8 ,

000

375

600

000 ,

5 

_

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However, if the same thrust was being produced by a turbo-propeller engine with a propeller efficiency of 55 percent at the same flight speed of 600 m.p.h., then the t.h.p. would be: 14,545 55 100 000 , 8  

Thus at 600 m.p.h. one lb. of thrust is the equivalent of about 3 t.h.p.

2.3 ENGINE THRUST IN FLIGHT

Since reference will be made to gross thrust, momentum drag and net thrust, it will be helpful to define these terms:

Gross or total thrust is the product of the mass of air passing through the engine and the jet velocity at the propelling nozzle, expressed as:

0

(PP )A +

g Wv_J

The momentum drag is the drag due to the momentum of the air passing into the engine relative to the aircraft velocity, expressed as

g WV

where: W = Mass flow in lb. per sec.

V = Velocity of aircraft in feet per sec.

G = Gravitational constant 32.2 ft. per sec. per sec.

The net thrust or resultant force acting on the aircraft in flight is the difference between the gross thrust and the momentum drag. From the definitions and formulae stated earlier under flight conditions, the net thrust of the engine, simplifying, can be expressed as:









g V V W A P P o  j

All pressures are total pressures except P which is static pressure at the propelling nozzle

W = Mass of air passing through engine (lb. Per sec.) VJ = Jet velocity at propelling nozzle (ft. per sec)

P = Static pressure across propelling nozzle (lb. Per sq. in) PO = Atmospheric pressure (lb. Per sq. in)

A = Propelling nozzle area (sq. in) V = Aircraft speed (ft. per sec.) G = Gravitational constant 32.2                 A P P Thrust essure g WV Thrust Momentum g wv A P P Thrust Gross g WV Drag Momentum O J J o ) ( Pr ) (

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The Balance of Forces and Expression for Thrust and Momentum Drag. Figure 2.4.

Graph of Thrust Against Forward Speed. Figure 2.5.

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2.3.1 EFFECT OF FORWARD SPEED

Since reference will be made to ‘ram ratio’ and Mach number, these terms are defined as follows:

Ram ratio is the ratio of the total air pressure at the engine compressor entry to

the static air pressure at the air intake entry.

Mach number is an additional means of measuring speed and is defined as the

ratio of the speed of a body to the local speed of sound. Mach 1.0 therefore represents a speed equal to the local speed of sound.

From the thrust equation, it is apparent that if the jet velocity remains constant, independent of aircraft speed, then as the aircraft speed increases the thrust would decrease in direct proportion. However, due to the ‘ram ratio’ effect from the aircraft forward speed, extra air is taken into the engine 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 forward speed so that the resultant loss of net thrust is partially recovered as the aircraft speed increases. A typical curve illustrating this point is shown in the figure 2.5. Obviously, the ‘ram ratio’ effect, or the return obtained in terms of pressure rise at entry to the compressor in exchange for the unavoidable intake drag, is of considerable importance to the turbo-jet engine, especially at high speeds. Above speeds of Mach 1.0, as a result of the formation of shock waves at the air intake, this rate of pressure rise will rapidly decrease unless a suitably designed air intake is provided; an efficient air intake is necessary to obtain maximum benefit from the ram ratio effect.

As aircraft speeds increase into the supersonic region, the ram air temperature rises rapidly consistent with the basic gas laws. This temperature rise affects the compressor delivery air temperature proportionally and, in consequence, to maintain the required thrust, the engine must be subjected to higher turbine entry temperatures. Since the maximum permissible turbine entry temperature is determined by the temperature limitations of the turbine assembly, the choice of turbine materials and the design of blades and stators to permit cooling are very important.

With an increase in forward speed, the increased mass airflow due to the ‘ram ratio’ effect must be matched by the fuel flow and the result is an increase in fuel consumption. Because the net thrust tends to decrease with forward speed, the end result is an increase in specific fuel consumption (s.f.c.), as shown by the curves for a typical turbo-jet engine in the figure 2.6.

At high forward speeds at low altitudes, the ‘ram ratio’ effect causes very high stresses on the engine and, to prevent over-stressing, the fuel flow is automatically reduced to limit the engine speed and airflow.

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Effects of speed on Thrust and Fuel Consumption. Figure 2.6.

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2.3.2 EFFECT OF AFTERBURNING ON ENGINE THRUST

At take-off conditions, the momentum drag of the airflow through the engine is negligible, so that the gross thrust can be considered to be equal to the net thrust. If after-burning is selected, an increase in take-off thrust in the order of 30 percent is possible with the pure jet engine and considerably more with the by-pass engine. This augmentation of basic thrust, is of greater advantage for certain specific operating requirements.

Under flight conditions, however, this advantage is even greater, since the momentum drag is the same with or without after-burning and, due to the ram effect, better utilisation is made of every pound of air flowing through the engine.

2.3.3 EFFECT OF ALTITUDE

With increasing altitude the ambient air pressure and temperature are reduced. This affects the engine in two inter-related ways:-

The fall of pressure reduces the air density and hence the mass airflow into the engine for a given engine speed. This causes the thrust or s.h.p. to fall. The fuel control system adjusts the fuel pump output to match the reduced mass airflow, so maintaining a constant engine speed.

The fall in air temperature increases the density of the air, so that the mass of air entering the compressor for a given engine speed is greater. This causes the mass airflow to reduce at a lower rate and so compensates to some extent for the loss of thrust due to the fall in atmospheric pressure. At altitudes above 36,089 feet and up to 65,617 feet, however, the temperature remains constant, and the thrust or s.h.p. is affected by pressure only.

Graphs showing the typical effect of altitude on thrust and fuel consumption are illustrated in Figure 2.7.

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Effects of Altitude on Thrust and Fuel Consumption. Figure 2.7.

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2.3.4 EFFECT OF TEMPERATURE

On a cold day the density of the air increases so that the mass of air entering the compressor for a given engine speed is greater, hence the thrust or s.h.p. is higher. The denser air does, however, increase the power required to drive the compressor or compressors; thus the engine will require more fuel to maintain the same engine speed or will run at a reduced engine speed if no increase in fuel is available.

On a hot day the density of the air decreases, thus reducing the mass of air entering the compressor and, consequently, the thrust of the engine for a given r.p.m. Because less power will be required to drive the compressor, the fuel control system reduces the fuel flow to maintain a constant engine rotational speed or turbine entry temperature, as appropriate; however, because of the decrease in air density, the thrust will be lower. At a temperature of 45C, depending on the type of engine, a thrust loss of up to 20 percent may be experienced. This means that some sort of thrust augmentation, such as water injection, may be required.

The fuel control system, controls the fuel flow so that the maximum fuel supply is held practically constant at low air temperature conditions, whereupon the engine speed falls but, because of the increased mass airflow as a result of the increase in air density, the thrust remains the same. For example, the combined acceleration and speed control (CASC) fuel system schedules fuel flow to maintain a constant engine r.p.m., hence thrust increases as air temperature decreases until, at a predetermined compressor delivery pressure, the fuel flow is automatically controlled to maintain a constant compressor delivery pressure and, therefore, thrust, Figure 2.8. illustrates this for a twin-spool engine where the controlled engine r.p.m. is high pressure compressor speed and the compressor delivery pressure is expressed as P3. It will also be apparent from this graph that

the low pressure compressor speed is always less than its limiting maximum and that the difference in the two speeds is reduced by a decrease in ambient air temperature. To prevent the L.P. compressor overspeeding, fuel flow is also controlled by an L.P. governor which, in this case, takes a passive role.

The Effect of Air Temperature on a Typical Twin Spool Engine Figure 2.8.

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2.4 PROPULSIVE EFFICIENCY

Performance of the jet engine is not only concerned with the thrust produced, but also with the efficient conversion of the heat energy of the fuel into kinetic energy, as represented by the jet velocity, and the best use of this velocity to propel the aircraft forward, ie. the efficiency of the propulsive system.

The efficiency of conversion of fuel energy to kinetic energy is termed thermal or internal efficiency and, like all heat engines, is controlled by the cycle pressure ratio and combustion temperature. Unfortunately this temperature is limited by the thermal and mechanical stresses that can be tolerated by the turbine. The development of new materials and techniques to minimise these limitations is continually being pursued.

The efficiency of conversion of kinetic energy to propulsive work is termed the propulsive or external efficiency and this is affected by the amount of kinetic energy wasted by the propelling mechanism. Waste energy dissipated in the jet wake, which represents a loss, can be expressed as

g V v W _j 2 ) (  2

where (vJ - V) is the waste velocity.

It is therefore apparent that at the aircraft lower speed range the pure jet stream wastes considerably more energy than a propeller system and consequently is less efficient over this range. However, this factor changes as aircraft speed increases, because although the jet stream continues to issue at a high velocity from the engine, its velocity relative to the surrounding atmosphere is reduced and, in consequence, the waste energy loss is reduced.

MODULE 15. GAS TURBINE ENGINE

European Aviation Safety Agency (EASA)

Part-66

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CONTENTS

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1

FUNDAMENTALS

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A

B

C

200 Ml (322 Km)

350 Ml (563 Km)

Path of Aircraft

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





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