MODULE 15. GAS TURBINE ENGINE

European Aviation Safety Agency (EASA) Part-66

JAR 66 CATEGORY B1

uk

engineering

MODULE 15 GAS TURBINE ENGINES

CONTENTS 1

FUNDAMENTALS ........................................................................ 1-1 1.1

1.2

1.3 1.4 1.5 1.6

1.7

1.8

1.9 16

WORK, POWER & ENERGY ...................................................... 1-1 1.1.1 Work ............................................................................. 1-1 1.1.2 Power............................................................................ 1-1 1.1.3 Energy .......................................................................... 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 PRINCIPLES OF JET PROPULSION ......................................... 1-4 1.3.1 Thrust Calculation. ........................................................ 1-4 GAS TURBINES ......................................................................... 1-6 THE BRAYTON CYCLE ............................................................. 1-7 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 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 CONTINUITY EQUATION AND BERNOULLI’S THEOREM ....... 1-13 1.8.1 Incompressible fluid. ..................................................... 1-13 1.8.2 Gas Laws ...................................................................... 1-15 SUBSONIC AIRFLOW THROUGH DIVERGENT AND CONVERGENT DUCTS

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 2.2 2.3

2.4

METHOD OF CALCULATING THE THRUST FORCES ............. 2-1 CALCULATING THE THRUST OF THE ENGINE ....................... 2-2 2.2.1 Comparison between thrust and horse-power ............... 2-6 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 PROPULSIVE EFFICIENCY ....................................................... 2-14

Issue 2 – April 2003

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engineering 2.5 2.6 2.7 2.8

3

GAS TURBINE ENGINES

FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP 2-15 SPECIFIC FUEL CONSUMPTION ............................................. 2-16 2.6.1 SPECIFIC FUEL CONSUMPTION – DEFINITION ........ 2-16 FLAT RATING ............................................................................ 2-16 PERFORMANCE RATINGS ....................................................... 2-16

INLET ............................................................................................ 3-1 3.1 3.2 3.3

3.4 3.5

4

MODULE 15

INTRODUCTION ........................................................................ 3-1 RAM COMPRESSION ................................................................ 3-1 3.2.1 Importance of Ram Compression .................................. 3-1 TYPES OF AIR INTAKES ........................................................... 3-2 3.3.1 PITOT INTAKES ........................................................... 3-2 3.3.2 DIVIDED ENTRANCE DUCT ........................................ 3-3 IDEAL INTAKE CONDITIONS .................................................... 3-4 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

COMPRESSORS .......................................................................... 4-1 4.1 4.2 4.3 4.4

4.5 4.6 25

COMPRESSORS GENERAL...................................................... 4-1 CENTRIFUGAL FLOW ............................................................... 4-1 4.2.1 Operation ...................................................................... 4-3 THE AXIAL FLOW COMPRESSOR............................................ 4-5 Operation .................................................................................... 4-6 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 418 AIR FLOW CONTROL SYSTEM – OPERATION ........................ 4-20 AEROFOIL THEORY AND THE AXIAL FLOW COMPRESSOR (CONTINUED) 4-

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 4.12 COMPARING THE FEATURES OF CENTRIFUGAL AND AXIAL FLOW COMPRESSORS ................................................................................... 4-30


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5

5.5 5.6 5.7 5.8

5.9 5.10

ENGINES

Centrifugal .................................................................... 4-30 Axial Flow ..................................................................... 4-30

INTRODUCTION ........................................................................ 5-1 COMBUSTION PROCESS ......................................................... 5-1 FUEL SUPPLY ........................................................................... 5-3 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 COMBUSTION CHAMBER PERFORMANCE ............................ 5-10 5.5.1 Combustion intensity ..................................................... 5-10 COMBUSTION EFFICIENCY ..................................................... 5-11 COMBUSTION STABILITY ......................................................... 5-11 POLLUTION CONTROL ............................................................. 5-12 5.8.1 Introduction ................................................................... 5-12 5.8.2 Sources of Pollution ...................................................... 5-12 EMISSIONS................................................................................ 5-12 MATERIALS ............................................................................... 5-14

TURBINE SECTION ..................................................................... 6-1 6.1 6.2 6.3

6.4 6.5

6.6

7

GAS TURBINE

COMBUSTION SECTION ............................................................. 5-1 5.1 5.2 5.3 5.4

6

MODULE 15

INTRODUCTION ........................................................................ 6-1 ENERGY TRANSFER FROM GAS FLOW TO TURBINE ........... 6-5 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 COMPRESSOR-TURBINE MATCHING ..................................... 6-11 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 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

EXHAUST ..................................................................................... 7-1 7.1 INTRODUCTION ........................................................................ 7-1 EXHAUST GAS FLOW .......................................................................... 7-3 7.3 CONSTRUCTION AND MATERIALS ......................................... 7-7


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8

GAS TURBINE ENGINES

NOISE REDUCTION .................................................................. 7-8 7.4.1 Sources of Engine Noise ............................................... 7-8 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

BEARINGS, SEALS AND GEARBOXES ..................................... 8-1 8.1

8.2

8.3

9

MODULE 15

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 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 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

LUBRICANTS AND FUEL ............................................................ 9-1 9.1 9.2 9.3

9.4

GAS TURBINE FUEL PROPERTIES AND SPECIFICATION ..... 9-1 FRACTIONAL DISTILLATION .................................................... 9-1 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 9.3.12 Safety precautions ........................................................ 9-8 GAS TURBINE OIL PROPERTIES AND SPECIFICATIONS ...... 9-9


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9.5 9.6


9.4.1 Viscosity........................................................................ 9-9 9.4.2 Hydro-Dynamics or Fluid Film Lubrication ..................... 9-9 9.4.3 Boundary Lubrication .................................................... 9-10 LUBRICATING OILS .................................................................. 9-10 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 10.2 10.3

10.4

10.5

10.6 10.7

INTRODUCTION ........................................................................ 10-1 BEARINGS ................................................................................. 10-1 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 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 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 OIL SEALS ................................................................................. 10-21 SERVICING ................................................................................ 10-21

11 ENGINE FUEL CONTROL SYSTEMS ......................................... 11-1 11.1 11.2 11.3

11.4 11.5 11.6 11.7

INTRODUCTION ........................................................................ 11-1 PURPOSE OF THE ENGINE FUEL SYSTEM ............................ 11-1 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 FACTORS GOVERNING FUEL REQUIREMENTS .................... 11-5 REQUIREMENTS OF THE ENGINE FUEL SYSTEM ................. 11-5 ENGINE FUEL SYSTEM COMPONENTS .................................. 11-5 FUEL PUMPS ............................................................................. 11-5 11.7.1 Fuel Pump Requirements.............................................. 11-5


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engineering


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 12.2

12.3 12.4 12.5

12.6

12.7 12.8

INTRODUCTION ........................................................................ 12-1 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 SEALING .................................................................................... 12-3 COOLING. .................................................................................. 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 EXTERNAL COOLING ............................................................... 12-13 12.6.1 External skin of aero-engine. ......................................... 12-13 12.6.2 Cooling of Accessories .................................................. 12-14 HP AIR FOR AIRCRAFT SERVICES.......................................... 12-15 External Air Tappings ................................................................. 12-15 ANTI-ICING SYSTEMS .............................................................. 12-18

13 STARTING AND IGNITION SYSTEMS ........................................ 13-1 13.1

BASIC PRINCIPLES OF GAS TURBINE ENGINE STARTING SYSTEMS 13.1.1 Purpose ........................................................................ 13-1 13.1.2 Essential Starting Requirements ................................... 13-1 STARTER MOTORS.............................................................................. 13-2 13.2.1 Electrical Starter Motor .................................................. 13-3


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13.3 13.4


13.2.2 Electric Starter/Generator.............................................. 13-3 13.2.3 Safety Interlocks ........................................................... 13-4 13.2.4 Air Turbo Starters .......................................................... 13-5 A300 STARTING SYSTEM ......................................................... 13-8 13.3.1 GE 6-50 Starting Procedure .......................................... 13-8 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 14.2 14.3

14.4 14.5 14.6

14.7 14.8

INTRODUCTION. ....................................................................... 14-1 ENGINE SPEED INDICATORS. ................................................. 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 EXHAUST GAS TEMPERATURE .............................................. 14-13 14.4.1 Thermocouples ............................................................. 14-13 FUEL FLOW METERING ........................................................... 14-17 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 VIBRATION ................................................................................ 14-24 WARNING LIGHTS .................................................................... 14-24

15 THRUST AUGMENTATION ......................................................... 15-1 15.1 15.2

15.3

INTRODUCTION ........................................................................ 15-1 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 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 16.1 16.2

INTRODUCTION ........................................................................ 16-1 TYPES OF TURBO-PROP ENGINES ........................................ 16-1 16.2.1 Coupled Power Turbine ................................................ 16-1


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16.4 16.5

16.6 16.7 16.8

16.9


16.2.2 Free Power Turbine ...................................................... 16-2 16.2.3 Compounded Engine .................................................... 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 TURBO-PROP PERFORMANCE ............................................... 16-7 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 ENGINE AND PROPELLER CONTROLS................................... 16-9 CONTROL OUTSIDE NORMAL FLIGHT RANGE ...................... 16-9 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 OVERSPEED SAFETY DEVICES .............................................. 16-16

17 TURBOSHAFT ENGINES ............................................................ 17-1 17.1 17.2 17.3 17.4 17.5

INTRODUCTION. ....................................................................... 17-1 FUEL CONTROL SYSTEM ........................................................ 17-4 ARRANGEMENTS ..................................................................... 17-6 DRIVE SYSTEMS....................................................................... 17-10 COUPLINGS .............................................................................. 17-13

18 AUXILLIARY POWER UNITS ...................................................... 18-1 18.1 18.2

18.3 18.4

18.5 18.6

18.7

18.8 18.9

INTRODUCTION ........................................................................ 18-1 GENERAL ARRANGEMENTS AND CONFIGURATION............. 18-3 18.2.1 Inlet Duct Arrangement ................................................. 18-7 18.2.2 Exhaust Duct Arrangement ........................................... 18-9 THE APU ENGINE ..................................................................... 18-10 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 APU OIL SYSTEM ...................................................................... 18-23 APU BLEED AIR SYSTEMS ....................................................... 18-25 18.6.1 direct from engine compressor ...................................... 18-25 18.6.2 SEPARATE LOAD COMPRESSOR .............................. 18-27 BAY COOLING ........................................................................... 18-28 18.7.1 Ram Air Cooling ............................................................ 18-28 18.7.2 Fan Air Cooling ............................................................. 18-28 APU POWERPLANT INSTALLATION. ....................................... 18-32 APU STARTING SEQUENCE .................................................... 18-34


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19 POWERPLANT INSTALLATION ................................................. 19-1 19.1

19.2

19.3

19.4

19.5 19.6 19.7

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 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 ENGINE DRAINS. ...................................................................... 19-18 19.3.1 Controlled Drains .......................................................... 19-18 19.3.2 Uncontrolled Drains....................................................... 19-20 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 ENGINE BUILD UNIT ................................................................. 19-29 19.5.1 Turbofan Engine ........................................................... 19-29 FIRE PREVENTION – BAYS OR ZONES................................... 19-38 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 20.2

20.3 20.4

20.5

FIRE DETECTORS .................................................................... 20-1 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 FIRE AND LOOP FAULT INDICATION (E.C.A.M.) ..................... 20-8 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 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 20.5.5 Extinguishant ................................................................ 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 21.2 21.3 21.4 21.5 21.6 21.7

PROCEDURES FOR STARTING AND GROUND RUNNING..... 21-1 STARTING ................................................................................. 21-3 UNSATISFACTORY STARTS .................................................... 21-7 ENGINE STOPPING. ................................................................. 21-8 ENGINE FIRES .......................................................................... 21-9 INTERPRETATION OF ENGINE POWER OUTPUTS AND PARAMETERS. 21-10 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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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 Time taken

= 2500 x 100 5

= 50,000 ft-lb’s/sec

Since 1hp = 550 ft-lb’s/sec then required hp = 50,000 550 = 90.9 hp (67.81 kw assuming no friction losses) Issue 2 – April 2003

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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 v2 2g where m = mass, v = velocity (in feet or metres per second), w = weight, g = gravity (32 ft/sec2 or 9.81m/sec2). 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.

Path of Aircraft

B

350 Ml (563 Km)

200 Ml (322 Km) N

C

A 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 g = the gravitational constant (taken as 32 ft/sec/sec or 9.81 m/sec2) The ACCELERATION imparted to the air is the difference between its inlet and outlet velocity. Issue 2 – April 2003

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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 =

W V2 - V1 g

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 =

W V2 - V1 g

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 Working Cycle. Figure 1.3.

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 Brayton Cycle. Figure 1.4.


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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 CHANGES ARE 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:

Open Ended Duct to Illustrate Continuity Equation Figure 1.6. Now volume flow = Area of duct x distance travelled (L) Time (sec) Issue 2 – April 2003

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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.

Duct System Figure 1.7.

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.


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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. ½mv2). 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.

Duct with Flow Control Valve Figure 1.8.

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.


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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.

Illustration of Pitot and Static Pressures Figure 1.9.

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.).

Duct of Varying Cross Sectional Area Figure 1.10.

The effects of a steady flow of incompressible fluid flows through a duct of varying cross sectional area shows: Issue 2 – April 2003

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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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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.

Diffuser Section Figure 1.12.

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


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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.

Divergent Duct. Figure 1.13. 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.

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.

A Con-Di Nozzle Figure 1.14.

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.


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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).

Changes in Temperature, Pressure and Velocity and the Brayton Cycle Figure 1.16. Issue 2 – April 2003

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

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

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


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.

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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 bypass 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.

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


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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. FORWARD GAS LOAD 57836 lbs

REARWARD GAS LOAD 46678 lbs

TOTAL THRUST 11158 lbs

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


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

Wv J g

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)

=

Pressure (P)

=

94 lb. per sq. in. (gauge)

Velocity

=

406 ft. per sec.

Mass flow (W) =

153 lb. per sec.

(vj)

182 sq. in.

The thrust = ( A  P) 

Wv j g

= (182  94) 

0

153  406 0 32

= 19,049lb. of thrust in a forward direction.


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Total Thrust of the Compressor. Figure 2.2.


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International Standard Atmosphere Figure 2.3. Issue 2 – April 2003

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Choked Nozzle Considering the formula for thrust under “choked” nozzle conditions: Thrust = ( P  P0 )A +

Wv J g

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  30

30 PO

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

= 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) Where P0 T0 30 273 + 15 273 + T0

= s.h.p. (observed) 

30 273  15  PO 273  TO

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

atmospheric temperature in deg. C (observed) I.S.A. standard sea level pressure (in Hg) I.S.A. standard sea level temperature in deg. K Atmospheric 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. 

jet thrust lb. 2.6

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 horsepower 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

FV 550 ft. per sec

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


5,000  600  8,000 375

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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 100 8,000   14,545 t.h.p. would be: 55 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: ( P  P0 )A +

Wv J g

The momentum drag is the drag due to the momentum of the air passing into the WV engine relative to the aircraft velocity, expressed as where: g 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.

WVJ   Momentum Thrust  wv  WV g Momentum Drag   Gross Thrust  ( P  Po ) A  J  g g   Pr essure Thrust  ( P  PO ) A  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, W Vj  V  simplifying, can be expressed as: P  Po  A  g All pressures are total pressures except P which is static pressure at the propelling nozzle W VJ P PO A V G

= = = = = = =

Mass of air passing through engine (lb. Per sec.) Jet velocity at propelling nozzle (ft. per sec) Static pressure across propelling nozzle (lb. Per sq. in) Atmospheric pressure (lb. Per sq. in) Propelling nozzle area (sq. in) Aircraft speed (ft. per sec.) Gravitational constant 32.2


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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

W (v j  V ) 2 2g

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.

Efficiency Plots of Differing Types of Engine to Airspeed Figure 2.9.


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2.5 FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP Primary engine design considerations, particularly for commercial transport duty, are those of low specific fuel consumption and weight. Considerable improvement has been achieved by use of the by-pass principle and by advanced mechanical and aerodynamic features and the use of improved materials. With the trend towards higher by-pass ratios, in the range of 15:1, the triple-spool and contrarotating rear fan engines allow the pressure and by-pass ratios to be achieved with short rotors, using fewer compressor stages, resulting in a lighter and more compact engine. S.f.c. is directly related to the thermal and propulsive efficiencies; that is, the overall efficiency of the engine. Theoretically, high thermal efficiency requires high pressures which in practice also means high turbine entry temperatures. In a pure turbo-jet engine this high temperature would result in a high jet velocity and consequently lower the propulsive efficiency. However, by using the by-pass principle, high thermal and propulsive efficiencies can be effectively combined by by-passing a proportion of the L.P. compressor or fan delivery air to lower the mean jet temperature and velocity. With advanced technology engines of high bypass and overall pressure ratios, a further pronounced improvement in s.f.c. is obtained. The turbines of pure jet engines are heavy because they deal with the total airflow, whereas the turbines of by-pass engines deal only with part of the flow; thus the H.P. compressor, combustion chambers and turbines, can be scaled down. The increased power per lb. of air at the turbines, to take advantage of their full capacity, is obtained by the increase in pressure ratio and turbine entry temperature. It is clear that the by-pass engine is lighter, because not only has the diameter of the high pressure rotating assemblies been reduced, but the engine is shorter for a given power output. With a low by-pass ratio engine, the weight reduction compared with a pure jet engine is in the order of 20 per cent for the same air mass flow. With a high by-pass ratio engine of the triple-spool configuration, a further significant improvement in specific weight is obtained. This is derived mainly from advanced mechanical and aerodynamic design, which in addition to permitting a significant reduction in the total number of parts, enables rotating assemblies to be more effectively matched and to work closer to optimum conditions, thus minimising the number of compressor and turbine stages for a given duty. The use of higher strength lightweight materials is also a contributory factor. For a given mass flow, less thrust is produced by the by-pass engine due to the lower exit velocity. Thus, to obtain the same thrust, the by-pass engine must be scaled to pass a larger total mass airflow than the pure turbo-jet engine. The weight of the engine, however, is still less because of the reduced size of the H.P. section of the engine. Therefore, in addition to the reduced specific fuel consumption, an improvement in the power-to-weight ratio is obtained.


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2.6 SPECIFIC FUEL CONSUMPTION When comparing engine performance, one of the most important considerations is how efficiently the power is produced. The amount of fuel consumed to produce a given horsepower lbs. thrust is known as “specific fuel consumption” or SFC. A typical aircraft fuel system measures the volume of fuel consumed. This is displayed in pounds per hour or PPH. To calculate fuel flow, specific fuel consumption found on the customer data sheet, is multiplied by the horsepower lbs. thrust produced. 2.6.1 SPECIFIC FUEL CONSUMPTION – DEFINITION

SFC = SPECIFIC FUEL CONSUMPTION is defined as the lbs of fuel used per HP/lbs of thrust per hour 2.7 FLAT RATING “Flat rating” is used by aircraft manufacturers when they select an engine that has a capability greater than the requirements of the aircraft. They then limit the power output of the engine. There are three distinct benefits derived from flat rating. One is the engine will have the ability to make take-off power at lower turbine temperatures over a wide range of outside air temperatures and pressure altitudes. Performance at altitude will be greatly enhanced. These two benefits result in the third benefit, longer engine life. A fourth benefit available on some engines is, a reserve of power which can be used to boost performance in an emergency ie. Loss of an engine during take - off. 2.8 PERFORMANCE RATINGS In the chart, performance ratings are compared on –1 through –12 engines. Notice the modifiers on the –1, -5, -6, -8 and –10 engines. These temperatures represent the effects of flat rating engines. Each engine will make take-off power below their turbine temperature limits to the ambient temperatures indicated. Engines that are not flat rated, such as the –3 or –11, would be unable to make take-off power below their turbine temperature limits when operating in conditions above 59F outside air temperatures.


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INLET

3.1 INTRODUCTION An air intake should deliver air to the engine compressor with a minimum loss of energy and at a uniform pressure under all engine operating conditions. The inlet duct is built in the shape of a subsonic divergent diffuser, so that the kinetic energy of the rapidly moving air can be converted into a ram pressure rise within the duct. This condition is referred to as “Ram Recovery”. 3.2 RAM COMPRESSION The degree of Ram Compression depends upon the following:i.

Frictional losses at those surfaces ahead of the intake entry which are “wetted” by the intake airflow.

ii.

Frictional losses at the intake duct walls.

iii.

Turbulence losses due to accessories or structural members located in the intake.

iv.

Aircraft speed.

v.

In a turbo-prop, drag and turbulence losses due to the prop blades and spinner.

Ram compression causes a re-distribution in the forms of energy existing in the air-stream. As the air in the intake is slowed up in endeavouring to pass into and through the compressor element against the air of increasing pressure and density which exists therein so the kinetic energy of the air in the intake decreases. This is accompanied by a corresponding increase in its pressure and internal energies and consequently compression of the air-stream is achieved within the intake, thus converting the unfavourable intake lip conditions into the compressor inlet requirements. Although ram compression improves the performance of the engine, it must be realised that during the process there is a drag force on the engine and hence the aircraft. This drag must be accepted since it is a penalty inherent in a ram compression process. (The added thrust more than makes up for this drag). 3.2.1 IMPORTANCE OF RAM COMPRESSION

At subsonic flight speeds, the ram pressure ratio is apparently quite small, say 1.33: 1 at 0.8M. Nevertheless, since the pressure rise due to ram compression is multiplied by the pressure ratio of the compressor, the ram pressure rise becomes significant even at subsonic speeds. Furthermore, the greater the forward speed of the aircraft becomes, the more significant is the ram compression; e.g. at 1.5M the ram pressure ratio may be about 3.5 : 1, and at 2.5M about 8 : 1.


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3.3 TYPES OF AIR INTAKES 3.3.1 PITOT INTAKES

This intake is suitable for subsonic or low supersonic speeds. Examples, 707, 747, A300B, Tristar, etc. The intake is usually short and is very efficient because the duct inlet is located directly ahead of the engine compressor. As the duct length increases, the risk of small airflow disturbances and pressure drop is increased. This inlet makes maximum use of ram effect until sonic speed is approached when efficiency falls due to shock wave formation at the intake lip. Pitot inlets can however suffer from inlet turbulences at high angles of attack and/or at low speeds.

Pitot Type Intakes.

Figure 3.1.


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The pitot type intake can be used for engines that are mounted in pods or in the wings although the latter sometimes requires a departure from the circular cross section due to the wing thickness.

Wing Leading Edge Intakes Figure 3.2 3.3.2 DIVIDED ENTRANCE DUCT

On a single engine aircraft with fuselage mounted engines, either a wing root inlet or a side scoop inlet may be used. The wing root inlet presents a problem to designers in the forming of the curvature necessary to deliver the air to the engine compressor. The side scoop inlet is placed as far forward of the compressor as possible to approach the straight line effect of the single inlet. Both types suffer faults, in a yaw or turn, a loss of ram pressure occurs on one side of the intake and separated, turbulent boundary layer air is fed to the engine compressor.

Divided Intakes. Figure 3.3.


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3.3.3 SUPERSONIC INTAKES At supersonic speeds, the pitot type of air intake is unsuitable due to the severity of shock waves which form and progressively reduce the intake efficiency as speed increases. To overcome this problem the compression intake was designed.

Supersonic Intakes. Figure 3.4. This type of intake produces a series of mild shock waves without reducing the intake efficiency, as the aircraft speed increases, so also does the intake compression ratio. At high mach numbers it becomes necessary to have an air intake which has a variable thrust area and spill doors to control the column of air.

3.4 IDEAL INTAKE CONDITIONS For air to flow smoothly through a compressor, its velocity should be about 0.5 mach at the compressor inlet; this includes aircraft flying faster than the speed of sound. Hence intakes are designed to decelerate the free stream airflow to this condition over the range of aircraft speeds. Intakes should also convert the kinetic energy into pressure energy without undue shock or energy loss. This means that the ideal compressor inlet pressure should be the same as the total head pressure at the intake lip. (Total head pressure = stagnation pressure, ie. static and dynamic pressure).


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Intake Efficiency The magnitude of the losses occurring in an intake during ram compression are measured by means of the intake efficiency. Typical optimum efficiencies of some common types of intake, at subsonic speeds assuming straight-through flow, are: a

Turbo-jet engine

Pitot

99 to 96%

Wing root 95 to 87% b

Turbo-prop engine

Side

89 to 80%

Annular

82 to 74% (DART)

In cases where the direction of flow of the air is reversed within the intake, these values are reduced by about 10%. 3.5 INTAKE ANTI-ICING Operations of present day aircraft necessitates flying in all weather conditions plus the fact that high velocity air induced into the intakes means a provision must be made for ice protection. There are three systems of thermal anti-icing; hot air, hot oil or electrical There is, however, one disadvantage and that is the loss of engine power. This loss must be corrected for on ground runs and power checks. 3.5.1 ENGINE HOT AIR ANTI-ICING

The hot air system provides surface heating of the engine and/or power plant where ice is likely to form. The affected parts are the engine intake, the intake guide vanes, the nose cone, the leading edge of the nose cowl and, sometimes, the front stage of the compressor stator blades. The protection of rotor blades is rarely necessary, because any ice accretions are dispersed by centrifugal action. The hot air for the anti-icing system is usually taken from the latter stages of the HP compressor and externally ducted, through pressure regulation valves, to the parts requiring protection. When the nose cowl requires protection, hot air exhausting from the air intake manifold may be collected and ducted to the nose cowl. Exhaust outlets are provided to allow the air to pass into the compressor intake or vent to atmosphere, thus maintaining a flow of air through the system.


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Hot Air Anti-Icing. Figure 3.5.


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3.5.2 ENGINE ELECTRICAL ANTI-ICING

There are two methods of electrical anti-icing: 1. Spray mat 2. Heater mats. 3.5.2.1

Spray Mat

The spray mat is so called because the conductor element is sprayed onto the base insulator to protect the spray mat from damage. An outer coating is sprayed on, sometimes called “Stone Guard” or “Erocoat”.

Spraymat Construction. Figure 3.6.


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Heater Mats

Heater Mat Construction. Figure 3.7.

Heater mats differ in design and construction according to their purpose and environment. The latest mats have elements which are made from a range of alloys woven in continuous filament glass yarn. Other elements are made from nickel chrome foil. The insulating material is usually polytetrafluoroethylene (PTFE) and the electrical control may be continuous or intermittent. 3.5.3 OIL ANTI-ICE

Oil anti-ice supplements the other two systems (hot air/electrical) and will also assist in cooling the oil system.


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Hot Oil Anti-Ice Figure 3.8.


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Intentionally Blank


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COMPRESSORS

4.1 COMPRESSORS GENERAL Compressors impart energy to the air stream raising its pressure and temperature. They are designed to operate efficiently over as wide a range of operating conditions as possible. The two basic types of compressor are: a

Centrifugal flow

b

Axial flow

4.2 CENTRIFUGAL FLOW The figure below illustrates different types of centrifugal compressors.

Types of Centrifugal Impeller. Figure 4.1.


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A Double Entry Centrifugal Compressor Figure 4.2.


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4.2.1 OPERATION

The centrifugal impeller is rotated at high speed by the turbine and centrifugal action causes the air between the impeller vanes to accelerate radially outwards until it is thrown off at the tip into the diffuser. The radial movement of the air across the impeller, from eye to tip, causes a drop in air pressure at the eye and the faster the impeller is turning, the lower the pressure at the eye becomes. The low pressure existing at the eye of the revolving impeller induces a continuous flow of air through the engine intake and into the eye of the impeller. The air, in turn, is accelerated across the impeller and passed into the diffuser. The kinetic energy in the air is then converted to pressure energy ready to enter the combustion chamber. The action of the diffuser is illustrated in figure 4.3.


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VANELESS SPACE

Centrifugal Compressor Function. Figure 4.3.


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The final volume and mass airflow delivered by the centrifugal compressor is dependent on: a

Pressure ratio

b

Operating RPM

c

Diameter of the impeller

NOTE:

This is assuming a constant air density at the inlet of the compressor.

4.2.1.1

Pressure Ratio

The ratio of the inlet pressure to outlet pressure of the compressor is called pressure ratio. The higher the pressure of the air the more efficiently the thrust will be produced with a corresponding improvement to the fuel economy of the engine. The maximum pressure ratio normally obtainable from a single stage centrifugal compressor is approximately 5:1 and from a two stage, approximately 8:1.Design of the more modern centrifugal compressors sees them approaching pressure ratios of 15:1. 4.2.1.2

Diameter of Impeller

A large impeller will deliver a greater mass of air than a small impeller, however a large diameter compressor leads to an increase in the frontal area of the engine causing excess drag forces on the aircraft. 4.3 THE AXIAL FLOW COMPRESSOR The axial flow compressor is by far the most popular type of compressor and, although it is more difficult to manufacture, it is a more efficient compressor. Handling a larger mass of air for any given diameter, it produces more power; and because the compression ratio is high – at least 9:1 and, it can be very much higher – it is a more economical engine. The airflow through the engine is parallel with the axis, hence the name ‘axial flow compressor’. The compressor consists of a single or multi-rotor assembly that carries blades of aerofoil section; it is mounted in a casing, which also houses the stator blades. The axial flow compressor increases the pressure of the air gradually (by approximately 1.2:1 per stage) over a number of ‘stages’, each stage comprising of a row of ‘rotor blades’, followed by a row of ‘stator blades’. Both the rotor and stator blades are of aerofoil section and form divergent passageways between adjacent blades of the same row. Figure 4.4 refers.


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Axial Flow Compressor Function. Figure 4.4. 4.3.1 OPERATION

The compressor rotor spool is driven by the turbine. The rotor blades accelerate the air rearwards, inducing a continuous flow of air into the inlet of the combustion chamber. The airflow emerges from the rotor stage with an increase in velocity, due to the rotating action of the blades, and with a rise in pressure and temperature caused by flowing through the divergent passage formed by the rotor.


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The airflow then passes through the divergent passages formed by the stator blades which convert some of the kinetic energy into pressure energy and directs the airflow onto the next set of rotors at the correct angle. The airflow emerges from each stage at approximately the same velocity as it entered, but with an increase (approximately 1.2:1) in pressure and, an increase in temperature. See graph below.

Combined Graph of Airflow Through an Axial Compressor. Figure 4.5. To present the airflow onto the first stage rotor blades at a suitable angle, some engines have inlet guide vanes in the air intake casing. The last row of stator blades is normally of wider chord than the preceding ones and serve to straighten the airflow before it enters the combustion system. In order to maintain the overall axial velocity more or less constant, the passageway between the stator casing and the compressor rotor forms a convergent duct in the direction of airflow, with long blades at the low pressure end and progressively shorter ones towards the high pressure end. (Figure 4.6 refers)


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Axial Compressor Layouts. Figure 4.6.


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The stator vanes are secured into the compressor casing or into stator vane retaining rings, which are themselves secured to the casing.

Axial Compressor Configuration Details. Figure 4.7. The stator vanes are positively locked in such a manner that they will not rotate around the casing. NOTE: Some stator vanes are variable to give variable airflow control, but these will be looked at when airflow control is studied.

Compressor Blade Attachment Methods Figure 4.8 Issue 2 – April 2003

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Compressor Blade Attachment Figure 4.9

The engine rotor assembly forms a hollow “drum” and is supported in ball and roller bearings and coupled to a turbine shaft. The rotor discs make up the drum and the rotor blades are attached as shown in the figure. On some smaller engines it becomes difficult to design a practical fixing, this is overcome by designing and producing blades integral with the disc and is called a “BLISK”.


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1

Extension Shaft Drive Stub

2

1st Stage Disk

3

Balance Weight

4

1st Stage Rotor Blades

5

Shroud Rings

6

7th Stage Rotor Blades

7

Air Inlet to Rotor Drum

8

1st Stage Blade Locking Strips

9

Front Main Bearing Housing Axial Compressor Rotor Details. Figure 4.10.


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Axial Compressor Stator Details Figure 4.11


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The mass and final volume of the airflow delivered by the compressor is dependent on: a. Pressure Ratio. Dependent on the number of stages employed. Axial flow compressors can achieve a much higher value than centrifugal. b. Diameter. For a similar mass flow capability, the axial flow compressor can be made smaller in diameter than the centrifugal type. c. Operating RPM. As with the centrifugal type, the RPM and hence the mass flow, is controlled by varying the amount of fuel delivered to the combustion system, but because of the way that the pressure rise takes place, the maximum pressure ratio in an axial flow compressor is achieved at a lower RPM, than in a centrifugal compressor. 4.4 COMPRESSOR STALL AND SURGE ‘Surge’ can occur in both centrifugal and axial flow compressors and is the reversal of the airflow in the compressor. It is a very undesirable condition, which can rapidly cause serious damage to the engine. In an axial flow compressor, ‘surge’ is nearly always preceded by stalling of some of the compressor blades. An aerofoil is said to be in a stalled condition when the airflow over its surface has broken down and no lift is being produced. If a row of compressor blades stall, then they will not be able to pass the airflow rearwards to the next stage and the airflow to the combustion chamber will ultimately stop. The lack of rearward airflow will allow the air in the combustion chamber to flow forward into the compressor until it reaches the row of stalled blades. Then a violent backwards and forwards oscillation of the airflow is likely to occur, which can rapidly cause extensive damage to the compressor blades and also overheating of the combustion and turbine assemblies. Stalling of the compressor blades can occur for various reasons and to appreciate how the condition comes about, a review of aerofoil theory and its application to the compressor is required. 4.4.1 AIRFLOW CONTROL SYSTEM PRINCIPLES 4.4.1.1

Compressor Stall and Surge

For any given engine there is only one set of conditions, mass flow, pressure ratio and rpm, at which all the compressor components are operating at their optimum effect. Compressors are designed to be most efficient in the higher rpm range of operation. The point at which the compressor reaches its maximum efficiency is known as the DESIGN POINT. Under design conditions the compressor produces Volume 2 a given compression ratio (ie. ) and the axial velocity (average velocity) Volume 1 of the gas remains approximately constant from the front to the rear of the compressor.


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The Angle of Attack of the airflow to the compressor aerofoil blades will be at its optimum. This is the design condition and the compressor is operating at its optimum performance. Although compression ratio varies with rpm it is not proportional to rpm. This fact emerges due to the fixed blade angles, which can only be correct at the design point. To illustrate this fact, refer to the diagram showing rpm and compression ratio. Consider a compressor running at 8,000 rpm and its compression ratio is 10:1. Let us say that the volume of air entering the compressor is 100cm3. The volume of the air passing through the fixed outlet annulus of the compressor will be 10cm3.

COMPRESSION RATIO

10:1

4:1

4000

8000

RPM Graph of Compression Ratio to RPM. Figure 4.12. Compressor R.P.M = 8,000

Compressor R.P.M. = 4,000

Compression Ratio = 10:1

Compression Ration = 4:1

Volume of gas (V1) = 100cm3



Volume of gas (V2) = 12.5cm3


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Now consider the same compressor operating at 4,000 rpm, the volume of air entering the compressor will be halved, eg. 50cm 3 there will also be a reduction in compression ratio to 4:1. Therefore the volume of air passing through the compressor fixed outlet annulus will be 12.5cm3. The following conditions will occur: a. Axial velocity will increase as it moves towards the rear stages relative to the front Low pressure stages. b…Since all stages are rotating at the same speed, there will be a NEGATIVE angle of attack at the rear high pressure stages and a POSITIVE angle of attack at the front low pressure stages.

Front

Rear Effect of Velocity on Blade Angle. Figure 4.13.


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Due to the increased velocity at the rear of the compressor, the outlet of the compressor will choke as the airflow reaches sonic velocity. At this point there will be a dramatic reduction in axial velocity resulting in the front compressor blades stalling. The end result will be compressor surge. To overcome the problem, a bleed valve is normally fitted in an intermediate stage of the compressor to bleed off the excess volume of air. This relieves the rear stages of the excess air causing choking while inducing an increased axial airflow through the early stages of the compressor, thus establishing conditions which are not conducive of stall and surge. Unfortunately this bleed valve does not completely cure the problem of stall as far as the first rotor stages are concerned and stall is still likely to occur. The blades stall when the angle of attack increases to too large a value. To overcome this problem, inlet guide vanes are used to pre-swirl the air onto the rotor blades. The effect of pre-swirling the air alters the angle of attack from a large value to the correct angle of attack. See figure 4.14.

Effect of Variable Guide Vane on Compressor Stage Figure 4.14 Issue 2 – April 2003

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4.4.2 COMPRESSOR CHARACTERISTICS

PRESSURE RATIO - Increasing

When a compressor is designed it is essential to establish the points at which it is likely to surge. Tests are carried out to determine the relationship between pressure ratio and mass flow at speeds covering the whole working range of the compressor. The results are recorded on a series of curves known as surge lines. To obtain the curves, the compressor is run at a constant speed, the mass airflow is gradually decreased and during this test the pressure ratio is carefully monitored. As the mass airflow reduces, there is an increase in pressure ratio. Eventually the compressor airflow becomes turbulent and the compressor surges. When this occurs, there is a rapid drop in pressure. The tests are carried out at various speeds until the whole working range of the compressor has been covered. During the test the points at which turbulence occurred at the various speeds are plotted. The points are then connected by drawing a line, this line is the surge line of the particular compressor being tested. During normal operation the engine is never allowed to operate beyond the surge line. A safety margin is established and the fuel and airflow control systems are adjusted so the engine will run within the safe limits. Figure 4.15 refers.

SAFETY MARGIN

UNSTABLE AREA

SURGE LINE WORKING LINE

80% 60%

70%

100 % 90% CONSTANT RPM LINES

AIRFLOW - Increasing Engine working line and surge margin. Figure 4.15.


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4.4.3 EFFECT OF TEMPERATURE ON THE OPERATING POINT OF THE AIRFLOW CONTROL SYSTEM

A change in temperature will affect mass airflow, compressor pressure ratio fuel flow and engine performance. The effect of a reduced temperature on the compressor at a fixed rpm being that the performance is comparable with that at a higher rpm at STANDARD TEMPERATURE. Consider an engine running at 10,000 rpm, the temperature of the day is 2ºC. If this is corrected for standard conditions (ISA 15ºC) the corrected rpm will be 10,235 see below. Observed rpm

=

Corrected rpm

=

10,000 rpm N



ISAinK Where 

=

 corrected rpm

T ambient in K 273  2 = ISA in K 273  15

10,000

=

275 288

Corrected rpm

=

10,000 0.977

=

10,235

From the above it is clear that temperature has an effect on the compressors mass flow rate. This is compounded further by the effect that temperature has a direct effect on the speed of sound and hence when the compressor chokes. It must be understood that if the engine is running at a fixed rpm and the temperature of the air is altered, the actual rpm of the compressor will be unaffected. However, the temperature change will affect the mach number of mass airflow and it is the speed of the compressor relative to the speed of the airflow (ie. Mach. Number) which is the critical factor. A decrease in temperature will raise the mach. Number. The mach. Number is the: SPEED OF THE OBJECT LOCAL SPEED OF SOUND


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The speed of the object is the compressor blade, if as previously stated, the mach. Number is raised with a decrease in temperature, the ‘fixed’ blade speed relative to the speed of the air, will be increased. To cater for this situation the operating point at which the variable inlet guide vanes move will have to be altered for varying air temperatures. To achieve this the actuator or ram of an airflow control system is temperature compensated. On a ‘cold’ day, the variable inlet guide vanes will operate earlier than on a ‘warm’ day.

Variation of Mach Number with Temperature. Figure 4.16. At a temperature of +60F

Local speed of sound is Mach 0.9 , no need for the VIGV’s as the compressor out let is not choked.

At a temperature of –40°F

Local speed of sound is Mach 1.0, the compressor outlet is choked, the first stages may stall, VIGV’s must start to open.


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4.5 AIR FLOW CONTROL SYSTEM – OPERATION The stages of the compressor are matched to give the highest efficiency in the speed range maximum rev/min. To extend the range of smooth operation over lower engine speeds, variable-incidence intake guide vanes and/or an air bleed valve are fitted. In the lower speed range the bleed valve opens to allow some of the air to escape from the rear stages of the compressor, thus restricting the mass air flow through the later stages and preventing an unstable flow pattern. When the bleed valve is open, the guide vanes if fitted are partially closed; at higher engine speeds, when the bleed valve is closed, the guide vanes if fitted move progressively towards the open position. The vanes are operated by a hydraulic ram which incorporates its own control mechanism and which receives a signal of engine speed in terms of hydraulic pressure from the engine speed governor in the fuel pump.

Combined Bleed Valve and Variable Guide Vane Operating System. Figure 4.17. Issue 2 – April 2003

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Intake Guide Vane Ram Setting Curve. Figure 4.18. Issue 2 – April 2003

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Air Bleed Valve Figure 4.19.

Intake Guide Vane Ram Setting Curve. Variable Guide Vane Hydraulic Figure 4.18. Actuator Figure 4.20. Issue 2 – April 2003

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To further improve airflow control, some engines will adopt a system of Variable Stator Vanes (VSV’s) as well as Variable Inlet Guide Vanes (VIGV’s) figure 4.21.

Variable IGV and Stator Vanes. Figure 4.21.


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Inlet Guide Vane and Variable stator Blade Linkwork. Figure 4.22.


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4.6 AEROFOIL THEORY AND THE AXIAL FLOW COMPRESSOR (CONTINUED) The blades of the axial flow compressor are aerofoils and as such behave in a similar way to aircraft mainplanes and propeller blades. The airflow across their surfaces produces lift and the amount of lift produced by an aerofoil depends on: a Its shape, area and smoothness of its surface. b the speed of airflow over the aerofoil. c

the angle at which the aerofoil meets the air.

Once manufactured, their area and shape will remain the same unless they are damaged in any way. Assuming the blades are in good condition, the variables will be the speed of the airflow and the angle at which the blades meet the air (angle of attack). 4.6.1 SPEED OF AIRFLOW OVER BLADES

This will vary with the rpm of the compressor rotor. The faster the rotor turns, then the faster the air flows over the blades. This will result in an increase in the axial velocity of the airflow through the compressor. 4.6.2 ANGLE OF ATTACK

This will vary with the combination of the rotational velocity of the blades and the axial velocity of the airflow. In the normal course of events, the angle of attack (VA) becomes progressively smaller as the compressor moves from a low rpm to a high rpm.(VT)

VT VA

VT

VT

VA

VT VA VA

Low R.P.M

R.P.M Increasing

High angle of attack

Angle of attack decreasing

High R.P.M Low angle of attack

Change of Angle of Attack Due to Increase in RPM. Figure 4.23.


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4.6.3 SOME IMPORTANT POINTS ABOUT ANGLE OF ATTACK

Airflow Over an Aerofoil Figure 4.24. An aerofoil can only produce lift between certain limits of angle of attack. 0 approx. 15.

At very large angles of attack the airflow breaks down and the aerofoil stalls.

At High Angles of Attack the Blade Will Stall. Figure 4.25 The greater the angle of attack (up to the stalling angle), the greater the lift and, also, the greater the drag. This means that a greater effort will be required to move the aerofoil through the air.

Lift/drag Vectors for Different Angles of Attack. Figure 4.26. All aerofoils have an ‘optimum’ angle of attack at which they produce most lift for the least drag. (‘Lift/drag ratio’) [2-4°].


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4.7 APPLICATION TO THE AXIAL FLOW COMPRESSOR In order for the compressor to deliver a high mass airflow for a minimum effort required to drive it, it is important that all the compressor blades are operating close to their optimum angle of attack at the designed optimum rpm of the engine. This is achieved by setting the blades onto the rotor assembly at a large enough angle so as to make allowance for the automatic reduction in angle of attack that will occur with increase in rpm. 4.7.1 COMPRESSOR RPM

An axial flow compressor is designed to operate at maximum speeds in the region of 8000-10,000 rpm, depending on size. At this rpm the engine will be producing a large amount of thrust and in order to vary the thrust it is necessary to vary the compressor rpm. When the compressor is operating at speeds below its designed rpm range, the axial velocity of the airflow through the compressor will decrease which will cause an increase in the angle of attack of the compressor blades. At low rpm, such as idling, the reduced axial velocity of the airflow may cause the angle of attack of some of the blades to increase beyond their stalling angle. A slight amount of LP blade stalling during ‘off design’ conditions is to be expected and only becomes a problem if a complete row of blades stall. 4.7.2 COMMON CAUSES OF COMPRESSOR STALL

Compressor stall normally occurs at low rpm and can be induced by: a disturbance of smooth airflow due to damaged or dirty blades. b disturbance of smooth airflow caused by damaged aircraft air intake. c

high combustion chamber pressure caused by over-fuelling during engine acceleration.

4.7.3 STAGGER ANGLE AND END BEND

The rotor blades are of airfoil section and usually designed to give a pressure gradient along their length to ensure that the air maintains a reasonably uniform axial velocity. The higher pressure towards the tip balances out the centrifugal action of the rotor on the airstream. To obtain these conditions, it is necessary to 'twist' the blade from root to tip to give the correct angle of incidence at each point. Air flowing through a compressor creates two boundary layers of slow to stagnant air on the inner and outer walls. In order to compensate for the slow air in the boundary layer a localised increase in blade camber both at the blade tip and root has been introduced. The blade extremities appear as if formed by bending over each corner, hence the term 'end-bend' Figure 4.27. 4.7.4 RECENT INNOVATIONS

The latest engines incorporate blades that have been designed and profiled using 3-D design techniques. This produces blades, which are curved in 3 dimensions, which are more aerodynamically efficient. Figure 4.28. Issue 2 – April 2003

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Stagger Angle and End Bend Figure 4.27.

3-D Blades Figure 4. 28.


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4.8 AIRFLOW CONTROL The higher the pressure ratio required from a compressor, the greater the number of compressor stages needed. The more stages there are, the more difficult becomes the problem of matching all the blades in both size and angle of attachment to make the compressor operate satisfactorily over a wide range of rpm. In order to maintain the airflow stability and reduce the tendency of high pressure ratio compressors to stall under certain conditions of aircraft flight and engine handling, methods of airflow control have already been discussed. 4.9 AIR BLEED VALVES (SUMMARY) The air bleed valve is operated automatically in response to signals of compressor rpm. It is in the open position below a certain critical rpm and bleeds air away from the centre stages of the compressor, ducting it overboard to atmosphere. This has the effect of increasing the axial velocity of the airflow through the early stages of the compressor, thereby reducing the angle of attack of the blades in that area. This prevents the early stages of the compressor from passing more air to the rear stages than can be accommodated in the space available. Above the critical rpm range the bleed valve is closed and all the air available from the compressor passes to the combustion system. 4.10 VARIABLE INTAKE GUIDE VANES (SUMMARY) All intake guide vanes give a certain amount of swirl to the incoming airflow. The swirl is in the direction of rotation of the compressor and the amount of swirl determines the angle of attack of the first stage rotor blades. The greater the degree of swirl imported by the IGV’s then the smaller the resultant angle of attack of the first stage rotor blades. Variable IGV’s present the air onto the first stage rotor blades with a maximum swirl angle during operation in the critical low rpm range and progressively reduce the degree of swirl in response to signals of compressor rpm. When operating at high rpm the airflow enters the compressor more or less axially. 4.11 MULTI-SPOOL COMPRESSORS (SUMMARY) Pressure ratios in excess of approximately 9:1 are best achieved by splitting the compressor into two independent sections as shown in the figure 4.29.


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Twin Spool Engine Figure 4.29. The total number of stages of compression is divided between two spools, each spool being driven at a different speed by separate turbines. This eases the problems of compressor blade matching and results in a very powerful, efficient and flexible engine. 4.12 COMPARING THE FEATURES OF CENTRIFUGAL AND AXIAL FLOW COMPRESSORS 4.12.1 CENTRIFUGAL

Merits.  Simplicity, cheaper, lighter, less prone to damage by FOD.  Not critical to surge and stall.  Will tolerate icing conditions. Associated Problems  Max pressure ratios 4:1 or 5:1. (on early types)  Capacity limited by tip speed.  Larger diameter of engine which leads to more drag.  Severe directional changes of gas flow which leads to friction.  High specific fuel consumption. 4.12.2 AXIAL FLOW

Merits  High Pressure Ratio.  Low specific fuel consumption.  More capacity for development.  Greater axial thrust. Associated Problems  Complex and expensive to produce.  Critical to stall/surge.


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COMBUSTION SECTION

5.1 INTRODUCTION The combustion chamber has the difficult task of burning large quantities of fuel, supplied through the fuel burners, with extensive volumes of air, supplied by the compressor, and releasing the heat in such a manner that the air is expanded and accelerated to give a smooth stream of uniformly heated gas at all conditions required by the turbine. This task must be accomplished with the minimum loss in pressure and with the maximum heat release for the limited space available. The amount of fuel added to the air will depend upon the maximum temperature rise required and, as this is limited by the materials from which the turbine blades and nozzles are made, the rise must be in the range of 700 to 1,200 deg.C. Because the air is already heated by the work done during compression, the temperature rise required at the combustion chamber may be between 500 and 800 deg.C. Since the gas temperature required at the turbine varies with engine speed, and in the case of the turbo-prop engine upon the power required, the combustion chamber must also be capable of maintaining stable and efficient combustion over a wide range of engine operating conditions. Efficient combustion has become more and more important because of the rapid increase in commercial aircraft traffic and the consequent increase in atmospheric pollution, which is seen by the general public as exhaust smoke. 5.2 COMBUSTION PROCESS Air from the engine compressor enters the combustion chamber at a velocity up to 500 feet per second, but because at this velocity the air speed is far too high for combustion, the first thing that the chamber must do is to diffuse it, i.e. decelerate it and raise its static pressure. Because the speed of burning kerosene at normal mixture ratios is only a few feet per second, any fuel lit even in the diffused air stream, which now has a velocity of about 80 feet per second, would be blown away. A region of low axial velocity has therefore to be created in the chamber, so that the flame will remain alight throughout the range of engine operating conditions. In normal operation, the overall air/fuel ratio of a combustion chamber can vary between 45:1 and 130:1. Kerosene, however, will only burn efficiently at, or close to, a ratio of 15:1, so the fuel must be burned with only part of the air entering the chamber, in what is called a primary combustion zone. This is achieved by means of a flame tube (combustion liner) that has various devices for metering the airflow distribution along the chamber.


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Typical Combustion Chamber Figure 5.1. Approximately 20 per cent of the air mass flow is taken in by the snout or entry section. Immediately downstream of the snout are swirl vanes and a perforated flare, through which air passes into the primary combustion zone. The swirling air induces a flow upstream of the centre of the flame tube and promotes the desired recirculation. The air not picked up by the snout flows into the annular space between the flame tube and the air casing. Through the wall of the flame tube body, adjacent to the combustion zone, are a selected number of holes through which a further 20 per cent of the main flow of air passes into the primary zone. The air from the swirl vanes and that from the primary air holes interacts and creates a region of low velocity recirculation. This takes the form of a toroidal vortex similar to a smoke ring, and has the effect of stabilising and anchoring the flame. The recirculating gases hasten the burning of freshly injected fuel droplets by rapidly bringing them to ignition temperature.


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It is arranged that the conical fuel spray from the burner intersects the recirculation vortex at its centre. This action, together with the general turbulence in the primary zone, greatly assists in breaking up the fuel and mixing it with the incoming air. The temperature of the combustion gases released by the combustion zone is about 1,800 to 2,000 deg.C., which is far too hot for entry to the nozzle guide vanes of the turbine. The air not used for combustion, which amounts to about 60 per cent of the total airflow, is therefore introduced progressively into the flame tube. Approximately half of this is used to lower the gas temperature before it enters the turbine and the other half is used for cooling the walls of the flame tube. Combustion should be completed before the dilution air enters the flame tube, otherwise the incoming air will cool the flame and incomplete combustion will result. An electric spark from an igniter plug initiates combustion and the flame is then self-sustaining. The design of a combustion chamber and the method of adding the fuel may vary considerably, but the airflow distribution used to effect and maintain combustion is always very similar to that described.

5.3 FUEL SUPPLY

Apportioning the Airflow Figure 5.2

So far little has been said of the way in which the fuel is supplied to the air stream. In general, however, two distinct principles are in use, one based on the injection of a finely atomised spray into a recirculating air stream, and the other based on the pre-vaporisation of the fuel before it enters the combustion zone. Although the injection of fuel by atomiser jets is the most common method, some engines use the fuel vaporising principle. In this instance, the flame tube is of the same general shape as for atomisation, but has no swirl vanes or flare. The primary airflow passes through holes in a baffle plate that supports a fuel feed tube.


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A Vaporising Combustion Chamber. Figure 5.3. The fuel is sprayed from the feed tube into vaporising tubes that are positioned inside the flame tube. These tubes bend through 180 degrees and, as they are heated by combustion, the fuel vaporises before passing forwards into the flame tube. The primary airflow passes down the vaporising tubes with the fuel and also through large (secondary) nozzles, which provide 'fans' of air to sweep the flame rearwards. Cooling and dilution air is metered into the flame tube in a manner similar to the atomiser flame tube. Vaporisers require starter spray nozzles to set the light up process in motion. 5.4 TYPES OF COMBUSTION CHAMBER There are three main types of combustion chamber at present in use for gas turbine engines. These are the multiple chamber, the tubo-annular chamber and the annular chamber. 5.4.1 MULTIPLE COMBUSTION CHAMBER

This type of combustion chamber is used on centrifugal compressor engines and the earlier types of axial flow compressor engines. It is a direct development of the early type of Whittle combustion chamber. The major difference is that the Whittle chamber had a reverse flow as this created a considerable pressure loss, the straight through multiple chamber was developed by Joseph Lucas Limited. The chambers are disposed around the engine and compressor delivery air is directed by ducts to pass into the individual chambers. Each chamber has an inner flame tube around which there is an air casing. The air passes through the flame tube snout and also between the tube and the outer casing as already described. The separate flame tubes are all interconnected. This allows each tube to operate at the same pressure and also allows combustion to propagate around the flame tubes during engine starting.


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Multiple Combustion Chambers. Figure 5.4.


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5.4.2 TUBO-ANNULAR COMBUSTION CHAMBER (ALSO KNOWN AS CAN-ANNULAR OR CANNULAR.)

The tubo-annular combustion chamber is a combination of the multiple and annular types. A number of flame tubes are fitted inside a common air casing. The airflow is similar to that already described and this arrangement embodies the ease of overhaul and testing of the multiple system with the compactness of the annular system.

Turbo-Annular Combustion System Figure 5.5.


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5.4.3 ANNULAR COMBUSTION CHAMBER

This type of combustion chamber consists of a single flame tube, completely annular in form, which is contained in an inner and outer casing. The airflow through the flame tube is similar to that previously described, the chamber being open at the front to the compressor and at the rear to the turbine nozzles. The main advantage of the annular chamber is that, for the same power output, the length of the chamber is only 75 per cent of that of a tubo-annular system for an engine of the same diameter, resulting in considerable saving of weight and production cost. Another advantage is that because interconnectors are not required, the propagation of combustion is improved. In comparison with a tubo-annular combustion system, the wall area of a comparable annular chamber is much less; consequently, the amount of cooling air required to prevent the burning of the flame tube wall is less, by approximately 15 per cent. This reduction in cooling air raises the combustion efficiency, to virtually eliminate unburnt fuel, and oxidises the carbon monoxide to non-toxic carbon dioxide, thus reducing air pollution. The introduction of the air spray type burner to this type of combustion chamber also greatly improves the preparation of fuel for combustion by aerating the overrich pockets of fuel vapour close to the burner; this results in a large reduction in initial carbon formation. A high by-pass ratio engine will also reduce air pollution, since for a given thrust the engine burns less fuel.

An Air Spray Fuel Nozzle. Figure 5.6.


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A Spray Nozzle. Figure 5.6.

Annular Combustion Chamber. Figure 5.7.


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5.4.4 REVERSE FLOW COMBUSTION CHAMBER

Reverse flow combustion chambers are used where the engine length is critical or where the thrust of the engine is not being produced by the exhaust of the primary air. They are often found on APU’s, turboprop and turbo-shaft engines or their derivatives such as the ALF 502 and LF507 engines used in the BAE 146 and RJ aircraft. By wrapping the combustion chamber around other components such as turbines the length of the engine can be significantly reduced. Losses in thrust do occur due to the changes in airflow and direction of pressure forces. This is not important in the types of engine where they are used as the majority of the thrust is derived by other sources. They are often found on engines with compound compressors, which have a centrifugal stages as the last stage of compression.

Reverse Flow Combustion Chamber. Figure 5.8.


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5.5 COMBUSTION CHAMBER PERFORMANCE A combustion chamber must be capable of allowing fuel to burn efficiently over a wide range of operating conditions without incurring a large pressure loss. In addition, if flame extinction occurs, then it must be possible to relight. In performing these functions, the flame tube and burner atomiser components must be mechanically reliable. Because the gas turbine engine operates on a constant pressure cycle, any loss of pressure during the process of combustion must be kept to a minimum. In providing adequate turbulence and mixing, a total pressure loss varying from about 5 to 10 per cent of the air pressure at entry to the chamber is incurred. 5.5.1 COMBUSTION INTENSITY

The heat released by a combustion chamber or any other heat generating unit is dependent on the volume of the combustion area. Thus, to obtain the required high power output, a comparatively small and compact gas turbine combustion chamber must release heat at exceptionally high rates. For example, a Rolls-Royce Spey engine will consume in its ten flame tubes 7,500 lb. of fuel per hour. The fuel has a calorific value of approximately 18,550 British Thermal Units per lb., therefore each flame tube releases nearly 232,000 British Thermal Units per minute. Expressed in another way, this is an expenditure of potential heat at a rate equivalent to approximately 54,690 horsepower for the whole engine.

Graph of Combustion Efficiency to Overall Air/Fuel Ratio. Figure 5.9.


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5.6 COMBUSTION EFFICIENCY The combustion efficiency of most gas turbine engines at sea-level take-off conditions is 100 per cent which reduces to 98 per cent at altitude cruise conditions. The values vary as shown in because of the reducing air pressure, temperature and fuel/air ratio. 5.7 COMBUSTION STABILITY Combustion stability means smooth burning and the ability of the flame to remain alight over a wide operating range. For any particular type of combustion chamber there is both a rich and a weak limit to the air/fuel ratio, beyond which the flame is extinguished. An extinction is most likely to occur in flight during a glide or dive with the engine idling, when there is a high airflow and only a small fuel flow, i.e. a very weak mixture strength. The range of air/fuel ratio between the rich and weak limits is reduced with an increase of air velocity, and if the air mass flow is increased beyond a certain value, flame extinction occurs. A typical stability loop is illustrated. The operating range defined by the stability loop must obviously cover the required air/fuel ratios and mass flow of the combustion chamber. The ignition process has weak and rich limits similar to those shown for stability. The ignition loop, however, lies within the stability loop, since it is more difficult to establish combustion under ‘cold' conditions than to maintain normal burning.

Combustion Stability Limits Figure 5.10.


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5.8 POLLUTION CONTROL 5.8.1 INTRODUCTION

Pollution of the atmosphere by gas turbine engines falls into two categories; visible (ie. smoke) and invisible constituents (eg. oxides or nitrogen, unburnt hydrocarbons, oxides of sulphur and carbon monoxide). The combination of the traditional types of HP burner (eg. Duplex) with increasing compression ratios has led to visible smoke during take-off and climb. The very strong public opinion against pollution of the atmosphere has forced engine manufacturers to develop methods of reducing smoke and other emissions. 5.8.2 SOURCES OF POLLUTION

Pollution occurs from incomplete combustion. When engines with high compression ratios (ie. above 15:1) are fitted with the traditional type of atomising burner, the high temperature, pressure and low turbulence within the combustion chamber prohibits adequate atomisation of the fuel when the engine is operating at low altitude, thus causing the formation of carbon particles. This can be reduced to an acceptable level by improving the airflow inside the combustion chamber and by introducing burners that are not so susceptible to changes in pressure 5.9 EMISSIONS The unwanted pollutants which are found in the exhaust gases are created within the combustion chamber. There are four main pollutants which are legislatively controlled; unburnt hydrocarbons (unburnt fuel), smoke (carbon particles), carbon monoxide and oxides of nitrogen. The principal conditions which affect the formation of pollutants are pressure, temperature and time. In the fuel rich regions of the primary zone, the hydrocarbons are converted into carbon monoxide and smoke. Fresh dilution air can be used to oxidise the carbon monoxide and smoke into non-toxic carbon dioxide within the dilution zone. Unburnt hydrocarbons can also be reduced in this zone by continuing the combustion process to ensure complete combustion. Oxides of nitrogen are formed under the same conditions as those required for the suppression of the other pollutants. Therefore it is desirable to cool the flame as quickly as possible and to reduce the time available for combustion. This conflict of conditions requires a compromise to be made, but continuing improvements in combustor design and performance has led to a substantially 'cleaner' combustion process.


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Olympus 593 Smoke Results Figure 5.11.

Pilot fuel

Main fuel

Dump diffuser

Main stage Exhaust gases to turbine

Compressor air

Pilot stage

BMW Rolls Royce are testing an axially staged combustion chamber for the BR715 engine, they claim it will cut the NOx by 50% without increasing CO, UHC and smoke emissions. Figure 5.12. x


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5.10 MATERIALS The containing walls and internal parts of the combustion chamber must be capable of resisting the very high gas temperature in the primary zone. In practice, this is achieved by using the best heat resisting materials available, the use of high heat resistant coatings and by cooling the inner wall of the flame tube as an insulation from the flame. The combustion chamber must also withstand corrosion due to the products of the combustion, creep failure due to temperature gradients and fatigue due to vibrational stresses.

Methods of Cooling the Flame Tube. Figure 5.13.


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TURBINE SECTION

6.1 INTRODUCTION The turbine has the task of providing the power to drive the compressor and accessories and, in the case of engines which do not make use solely of a jet for propulsion, of providing shaft power for a propeller or rotor. It does this by extracting energy from the hot gases released from the combustion system and expanding them to a lower pressure and temperature. High stresses are involved in this process, and for efficient operation, the turbine blade tips may rotate at speeds over 1,500 feet per second. The continuous flow of gas to which the turbine is exposed may have an entry temperature between 850 and 1,700 deg.C. and may reach a velocity of over 2,500 feet per second in parts of the turbine.

A Triple Stage Turbine with a Single Shaft. Figure 6.1.


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To produce the driving torque, the turbine may consist of Several stages each employing one row of stationary nozzle guide vanes and one row of moving blades. The number of stages depends upon the relationship between the power required from the gas flow, the rotational speed at which it must be produced and the diameter of turbine permitted. The number of shafts, and therefore turbines, varies with the type of engine., high compression ratio engines usually have two shafts, driving high and low pressure compressors. On high by pass ratio fan engines that feature an intermediate pressure system, another turbine may be interposed between the high and low pressure turbines, thus forming triple-spool system. On some engines, driving torque is derived from a free-power turbine. This method allows the turbine to run at its optimum speed because it is mechanically independent of other turbine and compressor shafts.

A Multi Stage Turbine driving Two Shafts. Figure 6.2. Issue 2 – April 2003

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The mean blade speed of a turbine has considerable effect on the maximum efficiency possible for a given stage output. For a given output the gas velocities, deflections, and hence losses, are reduced in proportion to the square of higher mean blade speeds. Stress in the turbine disc increases as the square of the speed, therefore to maintain the same stress level at higher speed the sectional thickness, hence the weight, must be increased disproportionately. For this reason, the final design is a compromise between efficiency and weight. Engines operating at higher turbine inlet temperatures are thermally more efficient and have an improved power to weight ratio. By-pass engines have a better propulsive efficiency and thus can have a smaller turbine for a given thrust.

A Multi Stage Turbine Driving Three Shafts. Figure 6.3.


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A Free Power Turbine. Figure 6.4. The design of the nozzle guide vane and turbine blade passages is based broadly on


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aerodynamic considerations, and to obtain optimum efficiency, compatible with compressor and combustion design, the nozzle guide vanes and turbine blades are of a basic aerofoil shape. There are three types of turbine; impulse, reaction and a combination of the two known as impulse-reaction. In the impulse type the total pressure drop across each stage occurs in the fixed nozzle guide vanes which, because of their convergent shape, increase the gas velocity whilst reducing the pressure. The gas is directed onto the turbine blades which experience an impulse force caused by the impact of the gas on the blades. In the reaction type the fixed nozzle guide vanes are designed to alter the gas flow direction without changing the pressure. The converging blade passages experience a reaction force resulting from the expansion and acceleration of the gas. Normally gas turbine engines do not use pure impulse or pure reaction turbine blades but the impulse-reaction combination. The proportion of each principle incorporated in the design of a turbine is largely dependent on the type of engine in which the turbine is to operate, but in general it is about 50 per cent impulse and 50 per cent reaction. Impulse-type turbines are used for cartridge and air starters.

Comparison between a Pure Impulse Turbine and an Impulse Reaction Turbine. Figure 6.5. 6.2 ENERGY TRANSFER FROM GAS FLOW TO TURBINE It will be seen that the turbine depends for its operation on the transfer of energy between the combustion gases and the turbine. This transfer is never 100 per cent because of thermodynamic and mechanical losses.


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When the gas is expanded by the combustion process, it forces its way into the discharge nozzles of the turbine where, because of their convergent shape, it is accelerated to about the speed of sound which, at the gas temperature, is about 2,500 feet per second. At the same time the gas flow is given a 'spin' or 'whirl' in the direction of rotation of the turbine blades by the nozzle guide vanes. On impact with the blades and during the subsequent reaction through the blades, energy is absorbed, causing the turbine to rotate at high speed and so provide the power for driving the turbine shaft and compressor.


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The torque or turning power applied to the turbine is governed by the rate of gas flow and the energy change of the gas between the inlet and the outlet of the turbine blades. The design of the turbine is such that the whirl will be removed from the gas stream so that the flow at exit from the turbine will be substantially 'straightened out' to give an axial flow into the exhaust system (Part 6). Excessive residual whirl reduces the efficiency of the exhaust system and also tends to produce jet pipe vibration which has a detrimental effect on the exhaust cone supports and struts. It will be seen that the nozzle guide vanes and blades of the turbine are 'twisted', the blades having a stagger angle that is greater at the tip than at the root. The reason for the twist is to make the gas flow from the combustion system do equal work at all positions along the length of the blade and to ensure that the flow enters the exhaust system with a uniform axial velocity. This results in certain changes in velocity, pressure and temperature occurring through the turbine.

Twisted Contour of Blades Figure 6.6.

The 'degree of reaction' varies from root to tip, being least at the root and highest at the tip, with the mean section having the chosen value of about 50 per cent.

Gas Flow Pattern Through a Nozzle and Turbine. Figure 6.7.


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The losses which prevent the turbine from being 100 per cent efficient are due to a number of reasons. A typical uncooled three-stage turbine would suffer a 3.5 per cent loss because of aerodynamic losses in the turbine blades. A further 4.5 per cent loss would be incurred by aerodynamic losses in the nozzle guide vanes, gas leakage over the turbine blade tips and exhaust system losses; these losses are of approximately equal proportions. The total losses result in an overall efficiency of approximately 92 per cent. 6.3 CONSTRUCTION The basic components of the turbine are the combustion discharge nozzles, the nozzle guide vanes, the turbine discs and the turbine blades. The rotating assembly is carried on bearings mounted in the turbine casing and the turbine shaft may be common to the compressor shaft or connected to it by a self-aligning coupling. 6.3.1 NOZZLE GUIDE VANES

The nozzle guide vanes are of an aerofoil shape with the passage between adjacent vanes forming a convergent duct. The vanes are located in the turbine casing in a manner that allows for expansion. The nozzle guide vanes are usually of hollow form and may be cooled by passing compressor delivery air through them to reduce the effects of high thermal stresses and gas loads.

Typical Nozzle Guide Vane Construction. Figure 6.8.


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6.3.2 TURBINE DISCS

Turbine discs are usually manufactured from a machined forging with an integral shaft or with a flange onto which the shaft may be bolted. The disc also has, around its perimeter, provision for the attachment of the turbine blades. To limit the effect of heat conduction from the turbine blades to the disc a flow of cooling air is passed across both sides of each disc. 6.3.3 TURBINE BLADES

The turbine blades are of an aerofoil shape, designed to provide passages between adjacent blades that give a steady acceleration of the flow up to the 'throat', where the area is smallest and the velocity reaches that required at exit to produce the required degree of reaction.

Methods of Turbine Blade Attachment. Figure 6.9.

The actual area of each blade cross-section is fixed by the permitted stress in the material used and by the size of any holes which may be required for cooling purposes. High efficiency demands thin trailing edges to the sections, but a compromise has to be made so as to prevent the blades cracking due to the temperature changes during engine operation. The method of attaching the blades to the turbine disc is of considerable importance, since the stress in the disc around the fixing or in the blade root has an important bearing on the limiting rim speed. The blades on the early Whittle engine were attached by the de Laval bulb root fixing, but this design was soon superseded by the 'fir-tree' fixing that is now used in the majority of gas turbine engines. This type of fixing involves very accurate machining to ensure that the loading is shared by all the serration’s. The blade is free in the serration’s when the turbine is stationary and is stiffened in the root by centrifugal loading when the turbine is rotating. Various methods of blade attachment are shown; however, the B.M.W. hollow blade and the de Laval bulb root types are not now generally used on gas turbine engines.


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A gap exists between the blade tips and casing, which varies in size due to the different rates of expansion and contraction. To reduce the loss of efficiency through gas leakage across the blade tips, a shroud is often fitted. This is made up by a small segment at the tip of each blade which forms a peripheral ring around the blade tips. An abradable lining in the casing may also be used to reduce gas leakage. Active Clearance Control (A.C.C.) is a more effective method of maintaining minimum tip clearance throughout the flight cycle. Air from the compressor is used to cool the turbine casing and when used with shroudless turbine blades, enables higher temperatures and speeds to be used.

Active Tip Clearance Control. Figure 6.10.


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6.3.4 DUAL ALLOY DISCS

Very high stresses are imposed on the blade root fixing of high work rate turbines, which make conventional methods of blade attachment impractical. A dual alloy disc, or 'blisk' as shown in fig. 6.11., has a ring of cast turbine blades bonded to the disc. This type of turbine is suitable for small high Power helicopter engines. Section Through a Dual Alloy Disc. Figure 6.11.

6.4 COMPRESSOR-TURBINE MATCHING The flow characteristics of the turbine must be very carefully matched with those of the compressor to obtain the maximum efficiency and performance of the engine. If, for example, the nozzle guide vanes allowed too low a maximum flow, then a back pressure would build up causing the compressor to surge; too high a flow would cause the compressor to choke. In either condition a loss of efficiency would very rapidly occur. 6.5 MATERIALS Among the obstacles in the way of using higher turbine entry temperatures have always been the effects of these temperatures on the nozzle guide vanes and turbine blades. The high speed of rotation which imparts tensile stress to the turbine disc and blades is also a limiting factor. 6.5.1 NOZZLE GUIDE VANES

Due to their static condition, the nozzle guide vanes do not endure the same rotational stresses as the turbine blades. Therefore, heat resistance is the property most required. Nickel alloys are used, although cooling is required to prevent melting. Ceramic coatings can enhance the heat resisting properties and, for the same set of conditions, reduce the amount of cooling air required, thus improving engine efficiency. 6.5.2 TURBINE DISCS

A turbine disc has to rotate at high speed in a relatively cool environment and is subjected to large rotational stresses. The limiting factor which affects the useful disc life is its resistance to fatigue cracking.


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In the past, turbine discs have been made in ferritic and austenitic steels but nickel based alloys are currently used. Increasing the alloying elements in nickel extend the life limits of a disc by increasing fatigue resistance. Alternatively, expensive powder metallurgy discs, which offer an additional 10% in strength, allow faster rotational speeds to be achieved. 6.5.3 TURBINE BLADES

A brief mention of some of the points to be considered in connection with turbine blade design will give an idea of the importance of the correct choice of blade material. The blades, while glowing red-hot, must be strong-enough to carry the centrifugal loads due to rotation at high speed. A small turbine blade weighing only two ounces may exert a load of over two tons at top speed and it must withstand the high bending loads applied by the gas to produce the many thousands of turbine horsepower necessary to drive the compressor. Turbine blades must also be resistant to fatigue and thermal shock, so that they will not fail under the influence of high frequency fluctuations in the gas conditions, and they must also be resistant to corrosion and oxidisation. In spite of all these demands, the blades must be made in a material that can be accurately formed and machined by current manufacturing methodsFrom the foregoing, it follows that for a particular blade material and an acceptable safe life there is an associated maximum permissible turbine entry temperature and a corresponding maximum engine power. It is not surprising, therefore, that metallurgists and designers are constantly searching for better turbine blade materials and improved methods of blade cooling. Over a period of operational time the turbine blades slowly grow in length. This phenomenon is known as 'creep' and there is a finite useful life limit before failure occurs.

Effect of Heat on Creep at Fixed Load. Figure 6.12.


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Effect of Load on Creep at Constant Temperature. Figure 6.13. The early materials used were high temperature steel forgings, but these were rapidly replaced by cast nickel base alloys which give better creep and fatigue properties. Close examination of a conventional turbine blade reveals a myriad of crystals that lie in all directions (equi-axed). Improved service life can be obtained by aligning the crystals to form columns along the blade length, produced by a method known as 'Directional Solidification'. A further advance of this technique is to make the blade out of a single crystal. Each method extends the useful creep life of the blade and in the case of the single crystal blade, the operating temperature can be substantially increased. A non-metal based turbine blade can be manufactured from reinforced ceramics. Their initial production application is likely to be for small high speed turbines which have very high turbine entry temperatures.


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Various turbine Blade Crystal Structures. Figure 6.14.


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Comparison of Turbine Blade Life Properties. Figure 6.15.


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6.6 DYNAMIC BALANCING PRINCIPLES 6.6.1 INTRODUCTION

We must all be familiar with the effects of unbalance in one form or another, but perhaps the most common effect is that arising from wheel unbalance in motor cars. At resonance conditions it causes wobble or bounce, the effects of which are transmitted to the driver through the steering column. This effect may be so violent as to make the car unsafe or at least uncomfortable to ride in, and the continual vibratory movements set up, even outside the resonance range will increase the rate of wear on the various linkages and add to driver and passenger fatigue. In order to increase passenger comfort, reduce wear and noise levels and also to increase the life of the engine between overhauls, design effort is put into the various aspects of minimising vibration in aero-engines. Design features are also included to permit correction of unbalance forces. Efforts are made to design engine bearing housings and carcasses with suitable stiffness to avoid resonance in the engine running range. In addition, precise balancing instructions are issued to control the rotating forces on the bearings which could:a) be transmitted to other parts of the engine or airframe structure. b) lead to engine failure in extreme cases. The loads on the bearings are of three main forms. These are: a) thrust loads due to the engine doing work. b) journal loads due to the dead weight of engine parts. c) unbalance loads.


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6.6.2 CENTRIFUGAL FORCE

Centrifugal Forces. Figure 6.16.. Centrifugal force acts on every particle which makes up the mass of the rotating element impelling each particle outwards and away from the axis, about which it is rotating, in a radial direction. If the mass of the rotating element is EVENLY DISTRIBUTED about the axis of rotation, the part is BALANCED and rotates WITHOUT VIBRATION. However, if there is a greater mass on one side of the rotor than the other, the centrifugal force acting on this heavy side exceeds the centrifugal force on the light side and pulls the entire assembly in the direction of the heavy side.

Eccentric Mass. Figure 6.17.


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ENGINES

The rotor has a heavy mass M on one side. The centrifugal force exerted by M causes the entire rotor to be pulled in the direction of force F. 6.6.3 CAUSES OF UNBALANCE

Unbalance may be caused by a variety of factors occurring singly or in combination with others. These factors include:a)

Eccentricity

Eccentricity exists when the geometric centreline of a part or assembly does not coincide with its axis of rotation. This may be as a result of locating features (eg. spigot location, bolt holes, splines, serration’s, couplings), being eccentric to the bearing location.

Eccentricity. Figure 6.18. b)

Variation in Wall Thickness

Variation in Wall Thickness. Figure 6.19. Variation in wall thickness may be as a result of eccentricity between an inner and outer diameter of a cylindrical type feature, or it may be as a result of a difference in thickness between a radial section of a disk type feature and the section diametrically opposite.


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c) Blade Distribution Unbalance can be caused by an unequal or unsymmetrical arrangement of a set of blades, either by reference to their mass moments or their dead weights depending on the size of the blades. This can be as a result of faulty weighting, inaccurate or illegible recording or assembly errors. d) Unsymmetrical Features These may be due to manufacturing processes, such as blow holes in castings or design features such as offset holes, locating dogs, slots, keyways, etc.

Unsymmetrical Features Figure 6.20. e) Distortion This can be caused by stress relieving, eg. after welding, or by unequal thermal growth during running. f) Fits and Clearances Clearance between mating parts allows relative movement of the parts and a consequent shift of the axis of rotation during running (or even during balancing). Joints incompletely assembled, eg. chamfers fouling radii, abutment faces not pulled together, may cause a ‘bent’ rotor or an unsuitable joint, which may cause a shift during running. It is important to prevent separate locating, or fixing, features from influencing each other eg. bolt holes, spigot locations, serration’s, etc. must be geometrically controlled to prevent ‘fighting’ between more than one feature. See also the section on tooling, adapters, drives, dummy rotors, etc.


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g) Swash

Swash. Figure 6.21. Swash is caused by out of squareness of abutment faces relative to the bearing diameter, abutment faces not being parallel across the component, eg. spacers, adjusting washers, disks, etc. It is important that the bolted joints are tightened in sequence and in increments according to the torquing instructions. h) Miscellaneous Foreign bodies inside assemblies, oil accumulation, carbon deposits, usually found when check balancing after running. 6.6.4 OBJECTIVE OF BALANCING

The objective of balancing is to determine how the unbalanced mass of the rotor must be compensated for in order to keep the bearings free of centrifugal force loading. 6.6.5 DEFINITION OF UNBALANCE

Unbalance can be defined as that condition which exists in a rotor when vibratory force or motion is imparted to its bearings as a result of centrifugal forces. Unbalance will, in general, be distributed throughout the rotor but can be reduced to:a)

static unbalance

b)

couple unbalance

c)

dynamic unbalance


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Static Unbalance In a gas turbine engine, static unbalance is primarily associated with thin discs such as turbine wheels or single compressor discs. It can be corrected by adding mass to the light side of the rotor. This can be achieved by a single weight DIAMETRICALLY OPPOSITE to the out of balance or by adding a number of smaller distributed weights having the same effect as a single weight. (This distribution can be determined by vectors).

Static Balance. Figure 6.22. Unbalance in a Long Rotor If a rotor is checked for static balance using knife edges it is possible to correct an out of balance condition to one end of the rotor by a correction weight at the other end of the rotor. Although in static balance, the rotor may now suffer from other kinds of unbalance. These are couple and dynamic unbalance.


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Couple Unbalance This arises when two EQUAL unbalance masses are positioned at opposite ends of a rotor and spaced at 180 from each other. If placed on knife-edges, the rotor would be statically balanced. However, when the rotor is rotated, the out of balance masses will cause a centrifugal force to act at each end and hence each end will vibrate independently as shown in figure 6.23.

Couple Unbalance. Figure 6.23.

Dynamic Unbalance This occurs when the unbalanced masses may be either unequal in size or positioned at some angle other than 180 to each other, or even both of these conditions. These unbalanced forces now cause the rotor to vibrate.


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6.6.6 FAN BALANCING

Before we look at fan balancing we must first look at vibration analysis techniques adopted on modern gas turbines and the reason for doing it. One of the requirements of an on-condition maintenance policy is that defects can be detected sufficiently early to permit rectification before secondary damage occurs. The analysis of engine vibration signatures is becoming an increasingly important tool for detecting early failure in mechanical components. A vibration monitoring system begins with a sensor, which may be a velocity transducer or a peizo electric accelerometer. They both convert the mechanical vibration of the machine into an electrical signal proportional to the vibrations produced and together with the associated electrical circuitry feed signals to either cockpit mounted gauges warning systems or a separate vibration analyser. Velocity Transducer This device operates on the principle of a permanent magnet to move within a coil, inducing voltage. Because of the moving parts with all the inherent disadvantages of wear, friction, etc. they have been superseded by the peizo electric principle. Peizo Electric Accelerometer In this device, vibrating forces are transmitted to a peizo electric disc the resultant deformation of the disc produces an electrical charge. Accelerometers have a greater frequency range than velocity transducers and their lack of moving parts makes them a much more stable and reliable means of collecting the basic vibration signal. Many different specifications for accelerometers and transducers are available and some of the considerations which govern their choice are:(1)

DYNAMIC RANGE. The amplitude range over which the device is required to perform.

(2)

SENSITIVITY. The severity of the vibration liable to be encountered.

(3)

FREQUENCY RESPONSE. The full operating frequency range required.

(4)

TEMPERATURE RANGE. The upper and lower temperature extremities to which the device will be subjected and also any heat soak conditions.


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Peizo Electric Transducer. Figure 6.24.


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Figure 6.24. shows a schematic diagram of a typical peizo electric accelerometer. The top nut is torque loaded to give the correct starting datum on the peizo crystal. When subjected to a force (caused by engine vibration) the piezo electric crystal produces an electric charge on its opposite faces. The output is fed to a charge amplifier, which produces the voltage required for the cockpit indicator or frequency analyser. Most modern transducers employ a synthetic piezo electric such as lead zirconate in preference to natural quartz crystal because of the higher sensitivity for the same force. In many cases, however, the choice of transducer will be dictated by the operating temperature. The maximum allowable temperature for transducers is typically 260C so they have to be sited on fan casings or in the by-pass ducting. Transducers may be fitted in more than one plane or more than one location. The analyser can then be used to select a ‘broadband’ or overall vibration measurement, which will give a quick guide to the condition of the engine. Vibration monitoring varies greatly from aircraft to aircraft. The operator’s requirements and the technology of the aircraft will dictate the equipment fitted. Large commercial aircraft will have fitted a flight deck indication of the vibration levels of engine spools, N1, N2, N3. Their main function is to warn the crew of a malfunction, ie. shed blade. The sensitivity of the vibration sensors may not be good enough for detailed condition monitoring or fan balancing. Extra vibration sensors are fitted to enable these functions to be carried. There are some modern aircraft, which will carry as a permanent fixture, eg. equipment that can carry out all vibration analysis requirement.


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EXHAUST

7.1 INTRODUCTION Aero gas turbine engines have an exhaust system which passes the turbine discharge gases to atmosphere at a velocity, and in the required direction, to provide the resultant thrust. The velocity and pressure of the exhaust gases create the thrust in the turbo-jet engine, but in the turbo-propeller engine only a small amount of thrust is contributed by the exhaust gases, because most of the energy has been absorbed by the turbine for driving the propeller. The design of the exhaust system therefore, exerts a considerable influence on the performance of the engine. The areas of the jet pipe and propelling or outlet nozzle affect the turbine entry temperature, the mass airflow and the velocity and pressure of the exhaust jet. The temperature of the gas entering the exhaust system is between 550 and 850 deg.C. according to the type of engine and with the use of afterburning can be 1,500 deg.C. or higher. Therefore, it is necessary to use materials and a form of construction that will resist distortion and cracking, and prevent heat conduction to the aircraft structure.

A Basic Exhaust System. Figure 7.1. A basic exhaust system is shown in fig. 7.1. The use of a thrust reverser, noise suppressor and a two position propelling nozzle entails a more complicated system as shown in fig. 7.2. The low by-pass engine may also include a mixer unit to encourage a thorough mixing of the hot and cold gas streams.


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An Exhaust System with a Thrust Reverser and Variable area propelling nozzle. Figure 7.2.


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7.2 EXHAUST GAS FLOW Gas from the engine turbine enters the exhaust system at velocities from 750 to 1,200 feet per second but, because velocities of this order produce high friction losses, the speed of flow is decreased by diffusion. This is accomplished by having an increasing passage area between the exhaust cone and the outer wall as shown in fig. 7.3. The cone also prevents the exhaust gases from flowing across the rear face of the turbine disc. It is usual to hold the velocity at the exhaust unit outlet to a Mach number of about 0.5, i.e. approximately 950 feet per second. Additional losses occur due to the residual whirl velocity in the gas stream from the turbine. To reduce these losses, the turbine rear struts in the exhaust unit are designed to straighten out the flow before the gases pass into the jet pipe.

Exhaust Cone Detail Figure 7.3. The exhaust gases pass to atmosphere through the propelling nozzle, which is a convergent duct, thus increasing the gas velocity. In a turbo-jet engine, the exit velocity of the exhaust gases is subsonic at low thrust conditions only. During most operating conditions, the exit velocity reaches the speed of sound in relation to the exhaust gas temperature and the propelling nozzle is then said to be 'choked'; that is, no further increase in velocity can be obtained unless the temperature is increased. As the upstream total pressure is increased above the value at which the propelling nozzle becomes ‘choked', the static pressure of the gases at the exit increases above atmospheric pressure. This pressure difference across the propelling nozzle gives what is known as 'pressure thrust' and is effective over the nozzle exit area. This is additional thrust to that obtained due to the momentum change of the gas stream.


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MODULE 15 GAS TURBINE ENGINES With the convergent type of nozzle a wastage of energy occurs, since the gases leaving the exit do not expand rapidly enough to immediately achieve outside air pressure. Depending on the aircraft flight plan, some high pressure ratio engines can with advantage use a convergent-divergent nozzle to recover some of the wasted energy This nozzle utilises the pressure energy to obtain a further increase in gas velocity and, consequently, an increase in thrust.

Gas Flow Through a Convergent Divergent Nozzle Figure 7.4.

From the illustration (fig. 7.4), it will be seen that the convergent section exit now becomes the throat, with the exit proper now being at the end of the flared divergent section. When the gas enters the convergent section of the nozzle, the gas velocity increases with a corresponding fall in static pressure. The gas velocity at the throat corresponds to the local sonic velocity. As the gas leaves the restriction of the throat and flows into the divergent section, it progressively increases in velocity towards the exit. The reaction to this further increase in momentum is a pressure force acting on the inner wall of the nozzle. A component of this force acting parallel to the longitudinal axis of the nozzle produces the further increase in thrust. The propelling nozzle size is extremely important and must be designed to obtain the correct balance of pressure, temperature and thrust. With a small nozzle these values increase, but there is a possibility of the engine surging, whereas with a large nozzle the values obtained are too low. A fixed area propelling nozzle is only efficient over a narrow range of engine operating conditions. To increase this range, a variable area nozzle may be used (Fig. 7.2.). This type of nozzle is usually automatically controlled and is designed to maintain the correct balance of pressure and temperature at all operating conditions. In practice, this system is seldom used as the performance gain is offset by the increase in weight. However, with afterburning a fully variable area nozzle is necessary. The by-pass engine has two gas streams to eject to atmosphere, the cool by-pass airflow and the hot turbine discharge gases.


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In a low by-pass ratio engine, the two flows are combined by a mixer unit (fig. 7.5.) which allows the by-pass air to flow into the turbine exhaust gas flow in a manner that ensures thorough mixing of the two streams.

Low By-pass Mixer Figure 7.5.


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In high by-pass ratio engines, the two streams are usually exhausted separately. The hot and cold nozzles are co-axial and the area of each nozzle is designed to obtain maximum efficiency. However, an improvement can be made by combining the two gas flows within a common, or integrated, nozzle assembly. This partially mixes the gas flows prior to its ejection to atmosphere. An example of both types of high by-pass exhaust system is shown in fig. 7.6.

High By-pass Engine Exhaust Systems. Figure 7.6.


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7.3 CONSTRUCTION AND MATERIALS The exhaust system must be capable of withstanding the high gas temperatures and is therefore manufactured from nickel or titanium. It is also necessary to prevent any heat being transferred to the surrounding aircraft structure. This is achieved by passing ventilating air around the jet pipe, or by lagging the section of the exhaust system with an insulating blanket. Each blanket has an inner layer of fibrous insulating material contained by an outer skin of thin stainless steel, which is dimpled to increase its strength. in addition, acoustically absorbent materials are sometimes applied to the exhaust system to reduce engine noise. When the gas temperature is very high (for example, when afterburning is employed), the complete jet pipe is usually of double-wall construction with an annular space between the two walls. The hot gases leaving the propelling nozzle induce, by ejector action, a flow of air through the annular space of the engine nacelle. This flow of air cools the inner wall of the jet pipe and acts as an insulating blanket by reducing the transfer of heat from the inner to the outer wall. The cone and streamline fairings in the exhaust unit are subjected to the pressure of the exhaust gases; therefore, to prevent any distortion, vent holes are provided to obtain a pressure balance. The mixer unit used in low by-pass ratio engines consists of a number of chutes through which the by-pass air flows into the exhaust gases. A bonded honeycomb structure is used for the integrated nozzle assembly of high by-pass ratio engines to give lightweight strength to this large component. Due to the wide variations of temperature to which the exhaust system is subjected, it must be mounted and have its sections joined together in such a manner as to allow for expansion and contraction without distortion or damage.

An Insulation Blanket Figure 7.7


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7.4 NOISE REDUCTION The problem of engine noise has always been associated with aircraft. Increases in engine power have given rise to increases in noise and the indications are that the increasing power trend will continue even more rapidly in future. High noise levels are responsible for psychological and physiological damage to humans and can also cause structural damage to aircraft; this has led to limits being set on maximum noise levels of aircraft by airport authorities and it appears that these limitations will be even more severe in future. The unit that is commonly used for measuring the noise annoyance level is the perceived noise decibel (PNdB). A PNdB is a measure of noise annoyance that take into account the pitch as well as the pressure (decibel) of a sound.

Comparative Noise Levels of Various Engine Types. Figure 7.8. The figure compares the noise level bands of various jet engine types (a busy restaurant will be 75-80 PNdB). 7.4.1 SOURCES OF ENGINE NOISE

To understand the problem of engine noise suppression, it is necessary to have a working knowledge of the noise sources and their relative importance. The noise from the jet engine mainly originates from three sources: a)

Exhaust jet

b)

Turbine

c)

Compressor and/or front fan.


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Exhaust Jet Jet noise is an externally generated source, which radiates in a rearward direction. It is caused by the mixing process of the high-speed exhaust gases with the surrounding air. In the mixing regions, a severe gradient of velocity exists normal to the jet and due to the viscosity of the air, this gradient produces vortices and shear forces which, in turn, produce quadrupole noise sources.

Noise Production in Sub & Super Sonic Air Flows. Figure 7.9 The noise produced by such a source will be proportional to p2Vje8, where p is the air density and Vje is the jet efflux velocity.


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Typical Quadrupole Noise Sources. Figure 7.10. Turbine Noise Noise from the turbine is made up from two sources: a. White Noise. “White”, random or background noise is caused by the reaction of each blade to the passage of air over its surface. There will always be noise from eddy shedding in the blade wake reacting back on the blade and causing random fluctuations over the blade surface (this source of noise may be likened to that produced by opening the quarter-light window on a car). Random noise will also be caused by turbulence in the air stream, which is sensed by the blade as a change in incidence with corresponding lift fluctuations and hence noise. b. Discrete Noise. Discrete noise is produced by the regular passage of rotating blades through the wakes from the preceding stationary vanes. If the space between vanes and blades is small, there is a cyclic interaction between pressure field. This can be overcome to some extent by design, ie. increasing the space. An additional source of discrete tones is caused by the rotating stage sensing changes of incidence and hence lift pressure, passing through the wakes of the upstream vanes. Compressor and Fan Noise Compressor noise whilst significant, was relatively small compared with the exhaust noise generated by turbojet and low by-pass engines. However as fans have got larger and by-pass ratios have increased the noise generated by the fan and compressor may well exceed that produced by the exhaust.


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Resultant Engine Noise Noise from an engine is the combination of noises produced by the compressor, the turbine and the nozzle. With the low by-pass engine, the exhaust noise level drops as the velocity of the exhaust gases is reduced and the turbine noise level drops as LP turbine mass flows and velocities are relatively reduced; but LP compressor noise becomes significant over a wider range of thrust. As the by-pass ratio is increased, the exhaust jet and turbine noise levels continue to drop and the LP compressor (fan) noise level continues to rise. This trend continues until the exhaust jet noise level is less than the turbine noise level and the fan noise reaches a level comparable with exhaust jet of a pure jet engine. There will be no such increase in the fan noise if a single-stage fan without IGV’s is aerodynamically suitable; instead, a significant decrease to a level comparable to the turbine noise will occur, as illustrated in the figure 7.12. This is because the more powerful elements of discrete tone and background noise are obviated.

Comparison of Noise Sources of Low and High By-pass Engines. Figure 7.12. Noise Suppression It has been seen that the first step towards noise suppression is at the design stage of the rotating and static parts of the engine. Thereafter, further reduction in the noise level emanating from a particular engine may be achieved by the incorporation of special materials and innovations during its construction. These additional methods of noise suppression are briefly described as: a)

Absorption by acoustic linings.

b)

Turbine, compressor and fan noise alleviated by control of nozzle area and shape.


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c)

Reduction of exhaust jet noise by mixing.

d)

Fan duct cowling design

Acoustic Linings One method of suppressing the noise from the fan stage of a high by-pass ratio engine is to incorporate a noise absorbent liner around the inside wall of the by-pass duct. The lining comprises a porous face-sheet, which acts as a resistor to the motion of the sound waves and is placed in a position such that it senses the maximum particle displacement in the progression of the wave. The depth of the cavity between absorber and solid backing is the tuning device, which suppresses the appropriate part of the noise spectrum. The figure shows two types of noise absorbent line; the figure shows the location of a liner to suppress fan noise from a high by-pass ratio engine and also the use of a liner to suppress the noise from the engine core. The disadvantage of using liners for reducing noise are the addition of weight and the increase in specific fuel consumption caused by increasing the friction of the duct walls.


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Noise Absorbing Materials and Location. Figure 7.13


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Nozzle Area and Shape Control In a high by-pass ratio engine with a single-stage fan without inlet guide vanes, the predominant sources governing the overall noise level are the fan and turbine. If the fan speed can be reduced without loss of thrust, then the engine noise level would be reduced. At conditions below maximum thrust, the multi-spool engine enables this to be accomplished by using a variable area nozzle to mechanically reduce the area of the hot stream final nozzle. This causes the speed of the LP turbine and its associated compressor spool to be reduced, producing a corresponding reduction in fan and turbine noise levels. However, the velocity of the hot stream will increase, producing a corresponding rise in exhaust jet noise. If the final nozzle area is reduced until the noise level of the fan, turbine and exhaust are of the same order, the optimum mean noise level for the engine will have been achieved. This normally occurs when the area of the hot stream final nozzle is reduced by approx. 50%. At the optimum nozzle area, the noise radiated towards the ground can be further reduced by a change in the geometrical shape of the nozzle.

Variable Area Nozzle Figure 7.14.


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Exhaust Jet Mixing

Figure 7.12. shows that the noise from the exhaust jet is the main contributor to the total noise generated by a low by-pass ratio turbo-fan. For a turbo-jet the noise from the exhaust is an even greater contributor to the whole. Fortunately it is comparatively easy to reduce the noise by increasing the mixture rate of the exhaust gases with the atmosphere. This can be achieved by increasing the contact area of the atmosphere with the gas stream by incorporating a corrugated or lobe-type suppresser in the propelling nozzle. The addition of a corrugate nozzle gives the effect shown in figure 7.16. In the corrugated nozzle, atmospheric air flows down the outside corrugations and into the exhaust jet to promote rapid mixing. In the lobe-type nozzle, the exhaust gases are divided to flow through the lobes and a small central nozzle. This forms a number of separate exhaust jets which rapidly mix with the air entrained by the suppresser lobes. Deep corrugations or lobes give a greater noise reduction, but the penalties incurred limit the size of the suppressers, eg. to achieve the required nozzle area, the overall diameter of the suppresser may have to be so large that excessive drag results. A nozzle may be designed to give a large reduction in noise level, but this could incur a considerable weight penalty due to the additional strengthening required. A compromise that gives a noticeable reduction in noise level with the minimum sacrifice of engine thrust or increase in weight is, therefore, the designer’s aim.

Type of Noise Suppressor. Figure 7.15.


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Improved Mixing by Corrugated Nozzle. Figure 7.16.

7.4.1.2

Recent Developments in Fan noise suppression

Rolls-Royce and GE are presently developing modified Trent and CF6 engines, respectively, which aim to reduce noise by incorporating chevron/saw tooth profiles to trailing edges of the fan and exhaust ducts. The manufacturers are also implementing extended areas of acoustic nacelle lining. In the case of the Trent proof of-concept study, the acoustic liner area is increased by 30 per cent to 95 sq ft.


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The Rolls-Royce programme, in conjunction with Boeing, is already commencing test flights of a modified Trent 800-powered B777-200ER as part of an overall effort to comply with ICAO Stage IV and QC2 noise levels - a pressing requirement necessary for future operations out of London's Heathrow airport. To this end, Rolls-Royce's new technology may be applicable to B747s either on a retrofit or new-build basis, and the team expects jet noise reductions of at least 3 EPNdB at ground level. Moreover, the modified fan case is expected to confer fan-noise reductions of 1.2 EPNdB and 7 EPNdB from inside the cabin - particularly regarding the frequencies which cause a "fan-buzz" signature. GE meanwhile, is also targeting future Airbus and Boeing aircraft operations with its modified CF6 engine. This has been statically tested in the autumn of 2002, with modified ducts and a new nozzle centrebody, for applicability to existing A300/A310s. According to GE, a peak jet noise reduction of 3.5dB is anticipated, while perceived reductions are in the order of ldB. GE intends to implement the new configuration into all its new-build CF6 engines from 2003. Like Rolls-Royce, GE is also targeting its big-fan modifications in conjunction with Boeing to facilitate ICAO Stage IV/QC2 compliant B747 operations in the near future.

These serrated ducts will improve flow mixing and reduce noise on the Trent 800. Figure 7.17.


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7.5 THRUST REVERSAL 7.5.1 INTRODUCTION

Thrust reversal is a means of reducing the landing run of an aircraft without excessive use of wheel brakes or the use of braking parachutes. On a propeller driven aircraft (piston and turbo prop), reverse thrust can be obtained by reversing the pitch of the propellers. On a pure turbo-jet this is not possible and the only simple and effective way of slowing the aircraft down quickly is to reverse the power as a deceleration force. This method is much safer than wheel brakes when landing on ice or snow covered runways. It can on some aircraft also be used to reduce speed in flight thus allowing a rapid rate of descent without an air brake system. The difference in landing distances between the same aircraft without reverse thrust and using reverse, is shown.

Effect of Thrust Reverse on Landing Run Figure 7.18. 7.5.2 REQUIREMENT FOR THRUST REVERSAL

To obtain reverse thrust, the jet efflux must be given a forward component of velocity. The mechanism to achieve this should fulfil the following requirements: a. A reasonable amount of thrust (50% of take-off thrust would be adequate) should be available in the reverse direction. b. The reverser should not affect the normal working of the engine and there should be no appreciable loss of thrust or increase in specific fuel consumption (SFC). c. When in use, the reverser should not cause debris or excessive amounts of hot air to enter the intake. d. The discharged hot gases should not impinge on parts of the aircraft (eg. nacelles, tyres, landing flaps, cabin windows, etc.). Impingement of the turbulent gas stream may cause damage by vibration as well as by heating. e. Fire hazards must be avoided. Hydraulic and lubricating systems should not be fitted near the jet pipe. f.

Weight, complexity and cost must be kept to a minimum.


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g. The reverser must not operate until required to do so. It is necessary to ensure that: 1. Accidental selection of reverse thrust is impossible. 2. No single failure in the operating system selects reverse thrust. 3. The thrust changing elements are biased away from the reverse thrust

position. 7.5.3 LAYOUT AND OPERATION OF TYPICAL THRUST REVERSING SYSTEMS

Clamshell door system The clamshell door system is a pneumatically operated system, as shown in detail in fig. 7.19. Normal engine operation is not affected by the system, because the ducts through which the exhaust gases are deflected remain closed by the doors until reverse thrust is selected by the pilot. On the selection of reverse thrust, the doors rotate to uncover the ducts and close the normal gas stream exit. Cascade vanes then direct the gas stream in a forward direction so that the jet thrust opposes the aircraft motion. The clamshell doors are operated by pneumatic rams through levers that give the maximum load to the doors in the forward thrust position; this ensures effective sealing at the door edges, so preventing gas leakage. The door bearings and operating linkage operate without lubrication at temperatures of up to 600 deg.C.

Clamshell Doors. Figure 7.19.

Bucket target system The bucket target system is hydraulically actuated and uses bucket-type doors to reverse the hot gas stream. The thrust reverser doors are actuated by means of a conventional pushrod system. A single hydraulic powered actuator is connected to a drive idler, actuating the doors through a pair of pushrods (one for each door). The reverser doors are kept in through the drive idler. The hydraulic actuator incorporates a mechanical lock in the stowed (actuator extended) position. Issue 2 – April 2003

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In the forward thrust mode (stowed) the thrust reverser doors form the convergentdivergent final nozzle for the engine.

Bucket Type Thrust Reverser. Figure 7.20.

Cold stream reverser system The cold stream reverser system can be actuated by an air motor, the output of which is converted to mechanical movement by a series of flexible drives, gearboxes and screwjacks, or by a system incorporating hydraulic rams. When the engine is operating in forward thrust, the cold stream final nozzle is 'open' because the cascade vanes are internally covered by the blocker doors (flaps) and externally by the movable (translating) cowl; the latter item also serves to reduce drag. On selection of reverse thrust, the actuation system moves the translating cowl rearwards and at the same time folds the blocker doors to blank off the cold stream final nozzle, thus diverting the airflow through the cascade vanes.


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engineering

ENGINES

Cold Stream Reverser. Figure 7.21.

7.5.3.1

Combination Reversers

Some engines are equipped with both cold and hot stream reversers, these have the some benefits of both types as well as some of the disadvantages.

Hot and Cold Stream Reverser. Figure 7.22.


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7.5.4 SAFETY FEATURES

Reverse thrust systems will have some of the following safety features incorporated: a. Reverse thrust cannot be selected until the engine throttle is brought back to idle. b. A mechanical lock prevents doors moving from the forward thrust position until reverse thrust is selected. c. Acceleration in forward thrust can only be obtained when the reverse thrust lever is de-selected and the doors are in the open position. d. Acceleration in reverse thrust can only be obtained when the reverse thrust lever is selected and the doors are in the closed position. e. The aircraft has to be on the ground or very close to it before reverse thrust selection is allowed (this does not apply to aircraft that use reverse thrust as an airbrake in flight). On the cold stream reverser/hot stream spoiler system, a mechanical interlock prevents reverse thrust being selected except when the throttle lever is at the idle position. After selection, acceleration of the engine to give reverse thrust is prevented until the translating cowl has moved rearwards. When the cowl has moved into position, a mechanical feedback from the cowl screw-jack unlocks the throttle control. 7.5.5 CFM 56 THRUST REVERSER FOR BOEING 737-300

The 737-300 is equipped with electrically controlled, hydraulically powered, fan only thrust reversers. The thrust reversers are interchangeable between the two engines except for the cascade basket assemblies and the strikers which deflect the Krueger flaps when the fan cowl translates aft.

Boeing 737-300 Thrust Reverser in Stowed and Deployed Positions. Figure 7.23


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Operation of the Blocker Doors. Figure 7. 24.

Reverser Control Valve Module. Figure 7.25. Reverser actuation is controlled by a control valve module, located on the forward bulkhead of each air-conditioning bay. This module contains two control valves (isolation and direction) and a manually operated (pinnable) maintenance shut-off valve. The control valves are operated by solenoids which are actuated by the thrust lever switches.


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The thrust reverser hydraulic system is only pressurised when thrust reverser actuation is required, or when required to resist motion from the stow commanded position.

Location of T/R Actuators and Synchronisation System. Figure 7.26. Application of hydraulic power to the reversers by operation of the reverse thrust levers is prevented unless the aeroplane is within 10 feet of the ground (radio altimeter 1 or 2), or is on the ground (right-hand main gear oleo compressed). Pulling an engine fire handle prevents the isolation valve from opening, or closes it if it is already open. A high idle is maintained for 4 seconds after activation of the weight on wheels switch in order to improve engine spool-up time in reverse. Each thrust reverser is powered by a separate hydraulic system, with a standby system available as an alternate source with a reduced deployment rate. An automatic restow system activates an actuator stow force anytime the reverser is sensed to be out of the stowed position during forward thrust operation.


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Thrust Reverser Schematic. Figure 7.27.


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A throttle interlock system restricts application of engine thrust when the reverser is not in its commanded position and automatically reduces engine thrust if uncommanded reverser translation occurs. Amber lights on the centre panel identify when the reversers are in the unlocked position. A "fault light" for each reverser is located in the Engine Module on the aft overhead panel. When this fault light is illuminated, the Master Caution is triggered after 12 seconds to indicate that a subsequent failure in the reverser system may cause uncommanded reverser motion.


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A

B

Thrust Reverser Controls. Figure 7.28. Issue 2 – April 2003

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BEARINGS, SEALS AND GEARBOXES

8.1 BEARINGS 8.1.1 INTRODUCTION

A bearing is any surface that supports or is supported by another surface. Bearings are designed to produce a minimum of friction and a maximum of wear resistance. Bearings must reduce the friction of moving parts and also take thrust loads or a combination of thrust and radial loads. Those which are designed primarily for thrust loads are called thrust bearings. The ball bearings are used to provide the thrust bearing as they can take both thrust and radial loads, and roller bearings are used to support the shaft whilst allowing axial movement. They are sometimes called expansion bearings. 8.1.2 BALL BEARINGS

A ball bearing consists of an inner race, an outer race and one or more sets of balls, and a ball retainer or cage. The purpose of the retainer or cage is to prevent the balls touching one another. Ball bearings are used for radial and thrust loads; a ball bearing specially designed for thrust loads would have very deep grooves in the races or be of the angular bearing type, these must always be fitted the correct way round! 8.1.3 ROLLER BEARINGS

These bearings are manufactured in various shapes and sizes and will withstand greater radial loads than ball bearings because of greater contact area. They allow axial movement of the shaft, this is very useful in a gas turbine due to expansion of the engine due to the heat it produces. 8.1.4 OTHER TYPES OF BEARINGS

It is rare to find taper roller or needle bearings used in gas turbine engines, however some APU’s use plain bearings to support the turbine end of the main shaft.


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Plain Roller Bearing Figure 8.1.

Examples of Bearing Types. Figure 8.2.


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8.2 BEARING CHAMBER OR SUMP One or more bearings are contained within a bearing chamber or sump. The chamber is sealed to prevent oil escaping into the engine and to prevent excessive air getting into the oil. 8.2.1 LUBRICATION

The bearing chamber will have an oil feed which is sprayed on to the bearing to lubricate and cool it. On some engines, to minimise the effect of the dynamic loads transmitted from the rotating assemblies to the bearing housings, a ‘squeeze film' type of bearing is used. They have a small clearance between the outer race of the bearing and housing with the clearance being filled with pressurised oil. The oil film dampens the radial motion of the rotating assembly and the dynamic loads transmitted to the bearing housing thus reducing the vibration level of the engine and the possibility of damage by fatigue. The oil will return to the oil system from the bottom of the bearing chamber, either by gravity or by suction from a scavenge pump. 8.2.2 SEALING Bearing chambers are usually sealed using air. The internal cooling air within the engine provides the air. Typical seals used are labyrinth, screw back and carbon types. . All of these seals need a differential pressure between inside and outside the bearing housing . Where pressure is available it is used, if the differential is too low, it can be boosted by suction from a scavenge pump. Carbon seals require the oil to be in contact with them to provide cooling for the seal. To prevent excess pressure building up within the bearing chamber, it is usually vented. This vent on some engines is taken to the oil tank to ensure that the whole system is working against the same pressure, or it goes to the oil pressure regulator to ensure that there is a constant pressure drop across the spray jets in the bearing housings. 8.2.2.1

Labyrinth Seals

Labyrinth seals are constructed of metal non-rotating lands, which are secured to various parts of the engine case and a series of cylindrical rotating knife-edge steps that mate with the lands. With this type of seal, there are no contacting parts. A precise clearance is designed into the seals to control the pressure, as the compressor air passes over the cascade of knife-edges, the pressure is reduced.


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Labyrinth Seal Figure 8.3. The labyrinth seal may be used in conjunction with an abradable coating on the stationary member as shown in the figure 8.3. 8.2.3 THREAD SEALS

Thread seals or screw back seals work in the same way as labyrinth seals, with a screw thread instead of the rings of a labyrinth seal. This means that any oil leakage towards the air will be driven back by the thread. This type of seal is used with other types of seal to reduce migration of oil to those seals.

Thread Seal Figure 8.4.


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8.2.4 CARBON SEAL

Another method of air sealing is achieved by using a carbon seal arrangement. They are used on the rotating assembly of a gas turbine and protection of engine drive components in accessory gearboxes. Carbon seals are manufactured of a mixture of carbon and graphite powder, bonded together with a viscous substance, such as coal tar. The carbon seal is fixed and held against the rotating seal by springs. Both the rotating seal and the carbon seals are machine ground and precision lapped to a micro finish.

Carbon Seal. Figure 8.5. 8.2.5 SPRING RING SEAL

This type of seal would normally be used around a main bearing assembly within the engine. It may be used in conjunction with a labyrinth or screw back type of seal.

Ring Seal

The ring seal is similar to a large stepped piston ring; it is located on a rotating shaft. When the shaft is stationary, the seal clamps tightly to the shaft. As the shaft rotates, the spring ring can expand slightly, under centrifugal force, when it then forms an effective seal with the adjacent stationary housing.

Figure 8.6.


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8.2.6 HYDRAULIC SEAL

This type of seal may also be found protecting the bearings on the main rotating assembly of an engine. It is fitted between the rotating shafts on a twin or triple spool engine. A hydraulic seal would be used in conjunction with another type of seal, as shown in figure 8.7.

Hydraulic Seal Figure 8.7. The seal consists of a circular baffle ring mounted on a rotating shaft; the rim of this ring sits in the centre of a circular depression in an outer rotating shaft. Oil from the bearing will fill this depression and be held there by centrifugal force. This oil reservoir will form a liquid seal with the rim of the rotating baffle ring. Any tendency for the oil to leak across this seal will be counteracted by air leakage across a backup seal.


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8.3 ACCESSORY DRIVE GEARBOXES 8.3.1 INTRODUCTION

Gearboxes provide the power for aircraft hydraulic, pneumatic and electrical systems in addition to providing various pumps and control systems for efficient engine operation. The high level of dependence upon these units requires an extremely reliable drive system. The drive for the gearbox is typically taken from a rotating engine shaft usually the HP shaft, via an internal gearbox, to an external gearbox that provides a mount for the accessories and distributes the appropriate geared drive to each accessory. A starter may also be fitted to provide an input torque to the engine. An accessory drive system on a high by-pass engine takes between 400 and 500 horsepower from the engine. 8.3.2 INTERNAL GEARBOX

The location of the internal gearbox within the core of an engine is dictated by the difficulties of bringing a driveshaft radially outwards and the space available within the engine core.

Mechanical Arrangements of Accessory Drive Gearboxes. Figure 8.8.


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Thermal fatigue and a reduction in engine performance, due to the radial driveshaft disturbing the gasfiow, create greater problems within the turbine area than the compressor area. For any given engine, which incorporates an axial-flow compressor, the turbine area is smaller than that containing the compressor and therefore makes it physically easier to mount the gearbox within the compressor section. Centrifugal compressor engines can have limited available space, so the internal gearbox may be located within a static nose cone or, in the case of a turbopropeller engine, behind the propeller reduction gear as shown in fig.8.8. On multi-shaft engines, the choice of which compressor shaft is used to drive the internal gearbox is primarily dependent upon the ease of engine starting. This is achieved by rotating the compressor shaft, usually via an input torque from the external gearbox. In practice the high pressure system is invariably rotated in order to generate an airflow through the engine and the high pressure compressor shaft is therefore coupled to the internal gearbox.

Types of Internal Gearbox Figure 8.9. Issue 2 – April 2003

To minimise unwanted movement between the compressor shaft bevel gear and radial driveshaft bevel gear, caused by axial movement of the compressor shaft, the drive is taken by one of three basic methods (fig. 8.9.). The least number of components is used when the compressor shaft bevel gear is mounted as close to the compressor shaft location bearing as possible, but a small amount of movement has to be accommodated within the meshing of the bevel gears. Alternatively, the compressor shaft bevel gear may be mounted on a stub shaft that has its own location bearing. The stub shaft is splined onto the compressor shaft that allows axial movement without affecting the bevel gear mesh. A more complex system utilises an idler gear that meshes with the compressor shaft via straight spur gears, accommodating the axial movement, and drives the radial driveshaft via a bevel gear arrangement. The latter method was widely employed on early engines to overcome gear engagement difficulties at high speed.

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To spread the load of driving accessory units, some engines take a second drive from the slower rotating low pressure shaft to a second external gearbox (fig.8.8.). This also has the advantage of locating the accessory units in two groups, thus overcoming the possibility of limited external space on the engine. When this method is used, an attempt is made to group the accessory units specific to the engine onto the high pressure system, since that is the first shaft to rotate, and the aircraft accessory units are driven by the low pressure system. A typical internal gearbox showing how both drives are taken is shown in fig.8.10. This method may also be used to drive speed sensors and governors for the low pressure shaft.

An Internal Gearbox With an LP and HP Output. Figure 8.10.


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8.3.3 RADIAL DRIVESHAFT

The purpose of a radial driveshaft is to transmit the drive from the internal gearbox to an accessory unit or the external gearbox. It also serves to transmit the high torque from the starter to rotate the high pressure system for engine starting purposes. The driveshaft may be direct drive or via an intermediate gearbox. To minimise the effect of the driveshaft passing through the compressor duct and disrupting the airflow, it is housed within the compressor support structure. On bypass engines, the driveshaft is either housed in the outlet guide vanes or in a hollow streamlined radial fairing across the low pressure compressor duct. To reduce airflow disruption it is desirable to have the smallest driveshaft diameter as possible. The smaller the diameter, the faster the shaft must rotate to provide the same power. However, this raises the internal stress and gives greater dynamic problems, which result in vibration. A long radial driveshaft usually requires a roller bearing situated halfway along its length to give smooth running. This allows a rotational speed of approximately 25,000 r.p.m. to be achieved with a shaft diameter of less than 1.5 inch without encountering serious vibration problems. 8.3.4 DIRECT DRIVE

In some early engines, a radial driveshaft was used to drive each, or in some instances a pair, of accessory units. Although this allowed each accessory unit to be located in any desirable location around the engine and decreased the power transmitted through individual gears, it necessitated a large internal gearbox. Additionally, numerous radial driveshafts had to be incorporated within the design. This led to an excessive amount of time required for disassembly and assembly of the engine for maintenance purposes. In some instances the direct drive method may be used in conjunction with the external gearbox system when it is impractical to take a drive from a particular area of the engine to the external gearbox. For example, figure8.8. shows a turbopropeller engine which requires accessories specific to the propeller reduction drive, but has the external gearbox located away from this area to receive the drive from the compressor shaft. 8.3.5 GEAR TRAIN DRIVE

When space permits, the drive may be taken to the external gearbox via a gear train (fig.8.8). This involves the use of spur gears, sometimes incorporating a centrifugal breather. However, it is rare to find this type of drive system in current use. 8.3.6 INTERMEDIATE GEARBOX

Intermediate gearboxes are employed when it is not possible to directly align the radial driveshaft with the external gearbox. To overcome this problem an intermediate gearbox is mounted on the high pressure compressor case and redirects the drive, through bevel gears, to the external gearbox. An example of this layout is shown in fig.8.8. Issue 2 – April 2003

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8.3.7 EXTERNAL GEARBOX

The external gearbox contains the drives for the accessories, the drive from the starter and provides a mounting face for each accessory unit. Provision is also made for hand turning the engine, via the gearbox, for maintenance purposes. Fig.8.11. shows the accessory units that are typically found on an external gearbox.

An External Gearbox. Figure 8.11.


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The overall layout of an external gearbox is dictated by a number of factors. To reduce drag it is important to present a low frontal area to the airflow. Therefore the gearbox is 'wrapped' around the engine. For maintenance purposes the gearbox is generally located on the underside of the engine to allow ground crew to gain access. However, helicopter installation design usually requires the gearbox to be located on the top of the engine for ease of access. The starter/driven gearshaft (fig.8.11.) roughly divides the external gearbox into two sections. One section provides the drive for the accessories which require low power whilst the other drives the high power accessories. This allows the small and large gears to be grouped together independently and is an efficient method of distributing the drive for the minimum weight. If any accessory unit fails, and is prevented from rotating, it could cause further failure in the external gearbox by shearing the teeth of the gear train. To prevent secondary failure occurring a weak section is machined into the driveshafts, known as a ‘shear-neck', which is designed to fail and thus protect the other drives. This feature is not included for primary engine accessory units, such as the oil pumps, because these units are vital to the running of the engine and any failure would necessitate immediate shutdown of the engine. Since the starter provides the highest torque that the drive system encounters, it is the basis of design. The starter is usually positioned to give the shortest drive line to the engine core. This eliminates the necessity of strengthening the entire gear train, which would increase the gearbox weight. However, when an auxiliary gearbox is fitted the starter is moved along the gear train to allow the heavily loaded auxiliary gearbox drive to pass through the external gearbox. This requires the spur gears between the starter and starter/driven gearshaft to have a larger face width to carry the load applied by the starter (fig.8.12.). When drive is taken from two compressor shafts, two separate gearboxes are required. These are mounted either side of the compressor case and are generally known as the 'low speed' and 'high speed' external gearboxes. 8.3.8 AUXILIARY GEARBOX

An auxiliary gearbox is a convenient method of providing additional accessory drives when the configuration of an engine and airframe does not allow enough space to mount all of the accessory units on a single external gearbox. A drive is taken from the external gearbox (fig.8.12.) to power the auxiliary gearbox, which distributes the appropriate gear ratio drive to the accessories in the same manner as the external gearbox.


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An External Gearbox with an Auxiliary Gearbox Drive. Figure 8.12.


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8.3.9 CONSTRUCTION AND MATERIALS Gears The spur gears of the external or auxiliary gearbox gear train (figs.8.11. and 8.12.) are mounted between bearings supported by the front and rear casings which are bolted together. They transmit the drive to each accessory unit, which is normally between 5000 and 6000 r.p.m. for the accessory units and approximately 20,000 r.p.m. for the centrifugal breather. All gear meshes are designed with 'hunting tooth' ratios which ensure that each tooth of a gear does not engage between the same set of opposing teeth on each revolution. This spreads any wear evenly across all teeth. Spiral (helical) bevel gears are used for the connection of shafts whose axes are at an angle to one another but in the same plane. The majority of gears within a gear train are of the straight spur gear type, those with the widest face carry the greatest loads. For smoother running, helical gears are used but the resultant end thrust caused by this gear tooth pattern must be catered for within the mounting of the gear. Gearbox sealing Sealing of the accessory drive system is primarily concerned with preventing oil loss. The internal gearbox has labyrinth seals where the static casing mates with the rotating compressor shaft. For some of the accessories mounted on the external gearbox, an air blown pressurised labyrinth seal is employed. This prevents oil from the gearbox entering the accessory unit and also prevents contamination of the gearbox, and hence engine, in the event of an accessory failure. The use of an air blown seal results in a gearbox pressure of about 3 lbs. per sq. in. above atmospheric pressure. To supplement a labyrinth seal, an 'oil thrower ring' may be used. This involves the leakage oil running down the driving shaft and being flung outwards by a flange on the rotating shaft. The oil is then collected and returned to the gearbox. Materials To reduce weight, the lightest materials possible are used. The internal gearbox casing is cast from aluminium but the low environmental temperatures that an external gearbox is subjected to allows the use of magnesium castings which are lighter still, The gears are manufactured from non-corrosion resistant steels for strength and toughness. They are case hardened to give a very hard wear resistant skin and feature accurately ground teeth for smooth gear meshing.


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LUBRICANTS AND FUEL

9.1 GAS TURBINE FUEL PROPERTIES AND SPECIFICATION Introduction In the earliest days of the gas turbine engine, kerosene was regarded as the most suitable fuel. It commended itself on the grounds of availability, cost, calorific value, burning characteristics and low fire hazard. Other types of petroleum fuels are not suitable for use in gas turbines because of the wide range of temperature and pressure over which combustion must occur and the necessity of keeping the weight and volume to a minimum. General Requirements A gas turbine fuel should have the following qualities: a)

Ease of flow under all operating conditions.

b)

Quick starting of the engine.

c)

Complete combustion under all conditions.

d)

A high calorific value.

e)

Non-corrosive.

f)

The by-products of combustion should have no harmful effect on the flame tubes, turbine blades, etc.

g)

Minimum fire hazards.

h)

Provide lubrication of the moving parts of the fuel system.

i)

The by-products of combustion should have minimal harmful effect on the environment

9.2 FRACTIONAL DISTILLATION This process is carried out in a fractionating column, which has a series of trays as shown in the figure. The effect of the superheated steam on the heated crude petroleum is to cause the lighter fractions to rise up the column. When rising, the vapour cools and a certain amount condenses on each tray until the tray is full of liquid to the overflow. Thus, each tray is a little cooler than the one below it, and therefore, lighter and lighter fractions will be present on each tray, as the vapours pass up the column. The temperature is controlled at the bottom of the column by the temperature of the crude oil, and at the top of the column by taking a certain amount of the product as it leaves, condensing it and pumping it back into the top of the column. This is known as the reflux. A certain amount of material will condense, which has a lower boiling point than the bulk of the liquid on a particular tray. To enable separation of these fractions, the liquid from a selected tray is drawn into a smaller auxiliary column, called a ‘sidestripper’. Here it is treated with steam that causes the lightest fractions to vaporise and pass along with the steam into the main column. Issue 2 – April 2003

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Fractioning Tower. Figure 9.1.


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The use of these side-strippers enables kerosene and gas oil to be obtained direct from the plant. Lubricating oil distillate, if such is present, can usually be drawn direct from a tray without the use of a side-stripper, while gasoline leaves the top of the column as a vapour and must be cooled to condense it to liquid gasoline. 9.3 PROPERTIES 9.3.1 EASE OF FLOW

The ease of flow of a fuel is mainly a question of viscosity, but impurities such as ice, dust, wax, etc., may cause blockages in filters and in the fuel system generally. Most liquid petroleum fuels dissolve small quantities of water and if the temperature of the fuel is reduced enough, water or ice crystals are deposited from the fuel. Adequate filtration is therefore necessary in the fuel system. The filters may have to be heated, or a fuel de-icing system fitted, to prevent ice crystals blocking the filters. Solids may also be deposited from the fuel itself due to the solidification of waxes or other high molecular weight hydrocarbons. Distillates heavier than kerosene, such as gas oil, generally have a pour point temperature too high for use in aircraft operating in low temperatures. If these fuels were to be used, some form of heating in the aircraft’s tanks and fuel system would be necessary. Such heating would obviously be an unreasonable complication. 9.3.2 EASE OF STARTING

The speed and ease of starting of gas turbines depends on the ease of ignition of an atomised spray of fuel. This ease of ignition depends on the quality of the fuel in two ways: a)

The volatility of the fuel at starting temperatures.

b)

The degree of atomisation, which depends on the viscosity of the fuel as well as the design of the atomiser.

The viscosity of fuel is important because of its effect on the pattern of the liquid spray from the burner orifice and because it has an important effect on the starting process. Since the engine should be capable of starting readily under all conditions of service, the atomised spray of fuel must be readily ignitable at low temperatures. Ease of starting also depends on volatility, but in practice the viscosity is found to be the more critical requirement. In general, the lower the viscosity and the higher the volatility, the easier it is to achieve efficient atomisation. 9.3.3 COMPLETE COMBUSTION

The exact proportion of air to fuel required for complete combustion is called the theoretical mixture and is expressed by weight. There are only small differences in ignition limits for hydrocarbons, the rich limit in fuels of the kerosene range being 5:1 air/fuel ratio by weight and the weak limit about 25:1 by weight. Flammable air/fuel ratios each have a characteristic rate of travel for the flame which depends on the temperature, pressure and the shape of the combustion chamber. Flame speeds of hydrocarbon fuels are very low and range from 0.3 – 0.6 m/sec. These low values necessitate the provision of a region of low air velocity within the flame tube, in which a stable flame and continuous burning are ensured. Issue 2 – April 2003

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Flame temperature does not appear to be directly influenced by the type of fuel, except in a secondary manner as a result of carbon formation, or of poor atomisation resulting from a localised over-rich mixture. The maximum flame temperature for hydrocarbon fuels is roughly 2,000C. This temperature occurs at a mixture strength slightly richer than the theoretical, owing to dissociation of the molecular products of combustion, which occurs at the theoretical mixture. Dissociation occurs above about 1,400C and reduces the energy available for temperature rise. The problem of the flame becoming extinguished in flight is not perfectly understood, but it appears that the type of fuel is of relatively minor importance. However, wide cut gasoline’s are more resistant to extinction than kerosene and engines are easier to relight using wide cut fuel. This is due to the higher vapour pressure of these fuels. 9.3.4 CALORIFIC VALUE

The calorific value is a measure of the heat potential of a fuel. It is of great importance in the choice of fuel, because the primary purpose of the combustion system is to provide the maximum amount of heat with the minimum expenditure of fuel. The calorific value of liquid fuels is usually expressed in megajoules (MJ) per litre. When considering calorific value, it should be noted that there are two values which can be quoted for every fuel, the gross value and the net value. The gross value includes the latent heat of vaporisation and the net value excludes it. The net value is the quantity generally used. The calorific value of petroleum fuels is related to their specific gravity. With increasing specific gravity (heavier fuels) there is an increase in calorific value per litre but a reduction in calorific value per kilogram. Thus, for a given volume of fuel, kerosene gives an increased aircraft range when compared with gasoline, but weighs more. If the limiting factor is the volume of the fuel tank capacity, a high calorific value by volume is the more important. 9.3.5 CORROSIVE PROPERTIES

The tendency of a turbine fuel to corrode the aircraft’s fuel system depends on two factors:a)

Water.

b)

Other corrosive substances, notably sulphur compounds.

The water which causes corrosion is dissolved water. It leads to corrosion of the fuel system, which is particularly important with regard to the sticking of sliding parts, especially those with small clearances and only small or occasional movement. Corrosion can also be caused by secondary effects, such as biological corrosion caused by plant spores, which are not killed off by the cracking process. Kerosene and diesel suffer from this form of contamination.


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9.3.6 EFFECTS OF BY-PRODUCTS OF COMBUSTION

Carbon deposition in the combustion system indicates imperfect combustion and may lead to:a)

A lowering of the surface temperature on which it is deposited, resulting in buckled flame tubes because of the thermal stresses set up by the temperature differences.

b)

Damage to turbine blades caused by lumps of carbon breaking off and striking them.

c)

Disruption of airflow through the turbine, creating turbulence, back-pressure and possible choking of the turbine, resulting in loss of efficiency.

It appears that carbon deposition depends on the design of the combustion chamber and the aromatic content of the fuel. (Aromatics are a series of hydrocarbons based on the benzene ring). The higher the aromatic content, the greater the carbon deposits. Sulphur will affect the turbine. Every effort is made to keep the sulphur content as low as possible in aviation turbine fuels. Unfortunately, removal of the sulphur involves increased refining costs and decreased supplies and so some sulphur is therefore permitted. 9.3.7 FIRE HAZARDS

There are three main sources of fire hazard; these arise from:a)

Fuel spillage with subsequent ignition of the vapour from a spark, etc.

b)

Fuel spillage on to a hot surface causing self-ignition.

c)

The existence of inflammable or explosive mixtures in the aircraft fuel tanks.

The first hazard depends on the volatility of the fuel. The lower the flash point, the greater the chance of fire through this cause. It is more difficult to ignite kerosene than to ignite gasoline or wide cut fuel in this way. The second hazard depends on the spontaneous ignition temperature of the fuel. In this respect, gasoline has a slightly higher spontaneous ignition temperature than kerosene, but if a fire does occur, the rate of spread is much slower in kerosene owing to its lower volatility. The third hazard depends upon the temperature and pressure in the tank and the volatility of the fuel. For any fuel there are definite temperature limits within which a flammable fuel vapour/air mixture will exist. If the temperature falls below the lower limit, the mixture will be too weak to burn, while if the temperature rises above the upper limit, the mixture is too rich to burn. At ground level the comparative temperature limits of flammability for gasoline and kerosene is as follows: a)

Gasoline. Upper limit -10C. Lower limit -42C.

b)

Kerosene. Upper limit +90C. Lower limit +43C.


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At higher altitudes the temperatures are somewhat lower. This information indicates that explosive conditions in the vapour space will occur with the low volatility turbine fuel under extremely hot weather conditions and with gasoline under extremely low temperature conditions. 9.3.8 VAPOUR PRESSURE

The vapour pressure of a liquid is a measure of its tendency to evaporate. The saturated vapour pressure (SVP) of a liquid (ie. the pressure exerted by vapour in contact with the surface of the liquid) increases with increasing temperature. When the SVP equals the pressure acting on the surface of the liquid, the liquid boils. Thus, the boiling point of a liquid depends on a combination of SVP, the pressure acting on its surface and its temperature. 9.3.9 FUEL BOILING AND EVAPORATION LOSSES

At high rates of climb, fuel boiling and evaporation is a problem which is not easily overcome. A low rate of climb permits the fuel in the tanks to cool and thus reduce its vapour pressure as the atmospheric pressure falls off. However, the rate of climb of many aircraft is so high that the fuel retains its ground temperatures, so that on reaching a certain altitude the fuel begins to boil. In practice this boiling has proved to be so violent that the loss is not confined to vapour alone. Layers of bubbles form and are swept through the tank vents with the vapour stream. This loss is analogous to a saucepan boiling over and is sometimes referred to as slugging. The amount of fuel lost from evaporation depends on several factors: a)

Vapour pressure of the fuel.

b)

Fuel temperature on take-off.

c)

Rate of climb.

d)

Final altitude of the aircraft.

Fuel losses as high as 20% of the tank contents have been recorded through boiling and evaporation. 9.3.10 METHODS OF REDUCING OR ELIMINATING FUEL LOSSES

Possible methods of reducing or eliminating losses by evaporation are: a. Reduction of the rate of climb. b. Ground cooling of the fuel. c. Flight cooling of the fuel. d. Recovery of liquid fuel and vapour in flight. e. Re-design of the fuel tank vent system. f.

Pressurisation of the fuel tanks.


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Reduction of the Rate of Climb Reducing the rate of climb imposes an unacceptable restriction on the aircraft and does not solve the problem of evaporation loss. This method is, therefore, not used. Ground Cooling of the Fuel This is not considered a practical solution, but in hot climates every effort should be made to shade refuelling vehicles and the tanks of parked aircraft. Flight Cooling of the Fuel The use of a heat exchanger, through which the fuel is circulated to reduce the temperature sufficiently to prevent boiling, is possible. High rates of climb, however, would not allow enough time to cool the fuel without the aid of heavy or bulky equipment. At a high true airspeed speeds TAS, the rise in airframe temperature due to skin friction increases the difficulty of using this method. On small highspeed aircraft the weight and bulk of the coolers becomes prohibitive. Recovery of Liquid Fuel in Flight This method would probably entail bulky equipment and therefore is unacceptable. Another method would be to convey the vapour to the engines and burn it to produce thrust, but the complications of so doing would entail severe problems. Redesign of the Fuel Tank Vent System The loss of liquid fuel could be largely eliminated by redesigning the vents, but the evaporative losses would remain. However, improved venting systems may well provide a more complete solution to the problem. Pressurisation of the Fuel Tanks There are two ways in which fuel tanks can be pressurised: a. Complete Pressurisation. Keeping the absolute pressure in the tanks greater than the vapour pressure at the maximum fuel temperature likely to be encountered eliminates all losses. However, this means that with gasoline type fuels, a pressure of about 8 psi absolute would have to be maintained at altitude and the tank would be subjected to a pressure differential of 6.5 psi at 50,000 feet. The disadvantage is that this would involve stronger and heavier tanks and a strengthened structure to hold the tanks. b. Partial Pressurisation. This prevents all liquid loss and reduces the evaporative loss. It involves strengthening the tanks and structure and the fitting of relief valves.


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9.3.11 FUEL ADDITIVES

Additives are added to fuel to improve its characteristics. Lubricity Additive. This is added to the fuel to reduce wear in fuel pumps, FCUs etc. when the fuel does not have sufficient lubricating properties of its own. Ice Inhibitor. Added to fuel to reduce/prevent ice crystals forming in the fuel and subsequently blocking fuel filters. This additive may also have biocide properties. Biocide. This is added to the fuel to prevent microbiological growth at the margins of free water within the aircraft fuel tanks. It can also be used as a shock treatment if contamination is suspected or as a preventative measure. 9.3.12 SAFETY PRECAUTIONS

All fuel will burn! Wide cut fuel is easier to ignite than kerosene. Strict No Smoking areas should be established around aircraft when any fuel system components are removed or fuel tanks are opened. This is important during refuelling and tank venting as fuel vapour present in the vent gasses produce an extremely explosive mixture. Fuel produces a very high static charge when flowing through pipes and meticulous care must be taken with bonding or grounding of pipes etc. The charge built up is dependant on flow rate, which is exceptionally high during refuel. Care must be taken when draining fuel from a component, as there is a chance of a static discharge occurring. Fuel soaked clothing is a great fire risk as the vapours given off are combustible. Fuel can also cause serious damage to the body. It degreases the skin which can cause dermatitis; the additives can increase the damage. Fuel also attacks sensitive areas of skin causing fuel burns (chemical burn) which can be extremely uncomfortable and may require hospitalisation. The chance of fuel burns to the skin is also increased if clothing becomes soaked, because of the proximity and rubbing action. Wash hands prior to going to the toilet. Eye protection may be required when entering systems that may contain fuel or fuel vapours. Avoid touching around your eyes if fuel is on your hands, you will only do it once! Fuel can be harmful if ingested, therefore hands should be thoroughly washed prior to eating. Spilt fuel on the floor or aircraft skin is very slippery and can even melt the soles of some types of shoe. Spills should be mopped up and disposed of in accordance with company procedures. Fuel spills should not be washed into domestic drains or sewers Spills to grass areas where as there is a chance of the fuel entering and polluting the water table below ground must be reported.


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9.4 GAS TURBINE OIL PROPERTIES AND SPECIFICATIONS Introduction There are two basic types of lubrication, they are Hydrodynamic (or film) lubrication, where the surfaces concerned are separated by a substantial quantity of oil, and Boundary lubrication, where the oil film may be only a few molecules thick. Before describing the types of lubrication in depth, it is necessary to explain viscosity. 9.4.1 VISCOSITY

The coefficient of viscosity, also known as dynamic viscosity, is a measure of the internal resistance of a fluid to relative movement, ie. its thickness, or film strength. Viscosity decreases with increase of temperature, the rate depending on the particular fluid considered. It is important for a lubricating oil that this rate of change of viscosity is predictable and is as small as possible. The viscosity index (VI) is an empirical number devised to indicate this change of viscosity with temperature, so than an oil with a high VI is preferable to one with a low VI. 9.4.2 HYDRO-DYNAMICS OR FLUID FILM LUBRICATION

Fluid film lubrication is the most common form of lubrication. It occurs when the rubbing surfaces are copiously supplied with oil and there is a relatively thick layer of oil between the surfaces (may be up to 100,000 oil molecules thick). The oil has the effect of keeping the two surfaces apart. Under these conditions the coefficient of friction is very small and may be as low as 0.001. The lubrication of a simple bearing (such as supports a rotating shaft) is a good example of fluid film lubrication (see figure 9.2.). The rotating shaft carries oil around with it by adhesion and successive layers of oil are carried along by fluid friction. As the shaft rotates it moves off-centre resulting in a narrow wedge of oil within which the pressure increases as the wedge narrows. For efficient lubrication this wedge, and the resulting increase of pressure, is essential as this keeps the surfaces apart. If this steady pressure increase breaks down, efficient film lubrication ceases and boundary lubrication occurs.

Lubrication of a Simple Bearing. Figure 9.2.


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In film lubrication, viscosity is the important factor because it controls the ability of the oil to keep the surfaces apart. A shaft revolving at high speed in a bearing must be free to carry oil round with it, with as little drag as possible. The rapid movement of one layer of oil slipping over another, with minimum drag, can only be achieved with a low viscosity oil. As the rotational speed decreases, the rate of deformation of the oil decreases, therefore the drag decreases and consequently an oil of higher viscosity may be needed if it is to be successfully carried round the bearing. The running temperature of the bearing is equally as important as the speed of rotation, as it controls the viscosity of the oil to be used. Bearing temperatures may vary, hence the need for oils with high VIs. 9.4.3 BOUNDARY LUBRICATION

If a shaft carries an appreciable load and rotates very slowly it will not carry round sufficient oil to give a continuous film and boundary lubrication will occur in which the friction is many times greater than in fluid film lubrication. Boundary lubrication is said to exist when the oil film is exceedingly thin and may only consist of a very few layers of molecules. It occurs due to high bearing loads, inadequate viscosity (possibly due to excessive bearing temperatures), oil starvation or loss of oil pressure. The friction is independent of the viscosity of the oil, but depends on the load and the “oiliness” of the lubricant. When a lubricating oil reduces the friction in a bearing to a lower value than that given by another lubricant of the same viscosity at the same bearing temperature, it is said to have a greater oiliness. It is thought that the reduction in friction is achieved by the fatty acids in the oil combining chemically with the bearing metal to form a “soap” which gives a boundary layer between the thin oil film and the bearing material to protect the metals from welding together. Boundary lubrication is not a desirable phase of lubrication as rupture of the thin film means wear, a very high surface temperature and possible seizure; therefore lubrication is designed to be hydro-dynamic if possible. However, boundary lubrication often occurs during starting conditions and may occur in piston engines at the end of reciprocating strokes. There is no precise division between boundary and fluid film lubrication although each is quite distinct in the way in which lubrication is achieved. In practice both forms occur at some time giving mixed film lubrication. 9.5 LUBRICATING OILS General Viscosity and VI are the factors which decide the lubricant for a particular purpose. The desirable viscosity for a given purpose is decided by bearing loads and clearances, sliding speeds, oil pump capacity, operating temperatures, etc. Therefore, in a lubricating oil specification, the desired viscosity is specified, together with VI and other safeguards to prevent the use of oil, which would deteriorate excessively or corrode the engine. Special engine tests are also carried out in test engines for each main batch of lubricating oil.


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Extreme Pressure Lubricants Extreme pressure lubricants are designed to work under boundary lubrication conditions. Certain chemicals known as extreme pressure (EP) additives (eg. sulphur, chlorine) give the lubricant the necessary quality. They appear to work in the same way as fatty acids, in that they combine chemically with the surface of the bearing metals. Additives Additives are substances added in small quantities to a lubricating oil to give it more desirable properties. Additives to lubricating oils are of the following main types:h. Extreme Pressure, as discussed. They are not in general use except in certain helicopter applications. i.

Anti-corrosion, which is used to protect some part of the engine.

j.

Detergents, which are used in piston engine oils to keep the engine clean.

k. Viscosity Index improvers, which make the VI as large as possible. l.

Pour Point Depressants, that permit oils to flow at lower temperatures than they were previously able.

m. Anti-foaming additives, that minimise foaming by increasing the surface tension of the oil. n. Anti-oxidants, which may be used to reduce the breakdown of the oil due to oxidation. 9.6 TURBINE OILS Introduction For lubrication of a high-speed turbine shaft running in contact bearings, an oil with good boundary lubrication properties and low viscosity is required. Because of the small amount of oil in circulation and the high bearing temperatures, good resistance to oxidation is essential. The earliest gas turbine engines were developed using straight mineral oils, but the operational requirements for low temperatures either on the ground or at a high altitude, led to the development of a range of straight mineral oils with viscosity’s far lower than those of conventional aircraft engine oil of that time. Mineral turbine oils are very rarely used now.


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9.6.1 FIRST GENERATION SYNTHETIC OILS

With the progressive development of the gas turbine engine to provide a higher thrust and compression ratio, mineral oils were found to lack stability and to suffer from excessive volatility and thermal degradation at the higher temperatures to which they were subjected. At this stage, a revolutionary rather than evolutionary oil development took place concurrently with engine development; lubricating oils derived by synthesis from naturally occurring organic products found an application in gas turbine engines. The first generation of synthetic oils were based on the esters of sebacic acid, principally dioctyl sebacate. As a class these materials exhibited outstanding properties which made them very suitable as the basis for gas turbine lubricants. Unlike straight mineral oils, the synthetic oils relied on additives (and in later formulations on multi-component additive packages) to raise their performance. This was particularly necessary to improve resistance to oxidation and thermal degradation (important properties which govern long term engine cleanliness). 9.6.2 SECOND GENERATION SYNTHETIC OILS

The introduction of the by-pass or turbo-fan engine raised further problems; in this engine the by-pass air acts as an insulating blanket and increases heat rejection to the lubricant. Therefore the requirement arose for an oil with an even greater resistance to thermal and oxidative stress. Several synthetic oils which meet this requirement have been developed. Known as Type 2 lubricants, they are blended from more complex esters and an additive package consisting of anti-oxidants, load-carrying additives, corrosion inhibitors, metal deactivators and foam inhibitors. 9.6.3 THIRD GENERATION SYNTHETIC OILS

Sustained flight at speeds in excess of Mach 1 aggravates the lubricant problem still further as the kinetic heating of the fuel reduces the effectiveness of fuel-cooled oil coolers. At Mach 2, oil temperatures may reach 260 - 316C, at which level standard ester-based oils degrade rapidly. In some military aircraft, Type 1 and Type 2 ester oils are still used under these conditions, but at greatly increased oil change frequencies. This procedure is expensive to operate as ideally the oil should remain in the engine for full engine life, with only periodic replenishment. More complex chemicals have been discovered which are more thermally stable than esters. However, they have various deficiencies such as poor low temperature properties or poor steel-on-steel lubricity. All are more expensive than esters. High temperature lubricants blended from specially developed ester oils, with new additives to limit oxidation degradation and corrosiveness and of increased load carrying ability, appear to offer the most practical solution for lubricating the jet engines in commercial supersonic transport (SST) aircraft. Many firms have been active in developing lubricants of this type and, after many submissions, two lubricants have been adopted for the Olympus 593 engines which power the BACAerospatiale Concorde.


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9.6.4 SAFETY PRECAUTIONS

There is much less risk of fire with oil, however it will burn if the conditions are right. The main risk with oil is to the body; prolonged contact with oil can cause dermatitis and/or cancer. The use of barrier cream and gloves cannot be overstated. Washing of hands before going to the toilet or eating is important, as is the reapplication of protection afterwards. Oil spills should be cleaned up as soon as possible and waste disposed of in accordance with company procedures.


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Intentionally Blank


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10 LUBRICATION SYSTEMS 10.1 INTRODUCTION There is always friction when two surfaces are in contact and moving, for even apparently smooth surfaces have small undulations, minute projections and depressions and actually touch at only a comparatively few points. Motion makes the small projections catch on each other and, even at low speeds when the surface as a whole is cool, intense local heat may develop leading to localised welding and subsequent damage as the two surfaces are torn apart. At higher speeds and over longer periods, intense heat may develop and cause expansion and subsequent deformation of the entire surface; in extreme cases large areas may be melted by the heat, causing the metal surfaces to weld together. The gas turbine engine is designed to function over a wider environment and under different operating conditions from its piston engine equivalent and therefore special lubricants have been developed to cope with the following main problems: a. High rpm compared with piston engines. b. Cold starting in winter can mean initial bearing temperatures of -54C which rapidly increases after starting to 232C. Therefore a good viscosity index and adequate cooling are required. On the other hand, the following advantages can be claimed for the gas turbine: a. There are fewer bearings and gear trains. b. Oil does not lubricate any parts directly heated by combustion and therefore oil consumption is low. c. There are no reciprocating loads. d. Bearings are generally of the rolling contact type and therefore only low oil pressures are needed (40 psi is normal). Turbo-prop engine lubrication requirements are more severe than those of a turbojet engine because of the heavily loaded reduction gears and the need for a highpressure oil supply to operate the propeller pitch control mechanisms. (For example, a twin relief valve in the Dart provides 35 psi for engine lubrication and 70 psi, which is fed to the propeller controller and boosted by a further pump to a pressure of 600 psi). 10.2 BEARINGS The early gas turbines employed pressure lubricated plain bearings but it was soon realised that friction losses were too high and that the provision of adequate lubrication of these bearings over the wide range of temperatures and loads encountered was more difficult than for piston engine bearings. As a result, plain bearings were abandoned in favour of the rolling contact type as the latter offered the following advantages: a

Lower friction at starting and low rpm.

b

Less susceptibility to momentary cessation of oil flow.


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c

The cooling problem is eased because less heat is generated at high rpm.

d

The rotor can be easily aligned.

e

The bearings can be made fairly small and compact.

f

The bearings are relatively lightly loaded because of the absence of power impulses.

g

Oil of low viscosity may be used to maintain flow under a wide range of conditions and no oil dilution or pre-heating is necessary.

The main bearings are those which support the turbine and compressor assemblies. In the simplest case (a single spool engine), these usually consist of a roller bearing at the front of the compressor and another in front of the turbine assembly, with a ball bearing behind the compressor to take the axial thrust on the main shaft. “Squeeze film” main bearings have been introduced to reduce transfer of rotor vibration to the aircraft. In this type of bearing pressure oil is fed to a small annular space between the bearing outer track and the housing. Figure 10.1. shows that the bearing will therefore “float” in pressure oil, which will damp out much of the vibration. Squeeze film bearings are fitted to the Spey and all subsequent aero engines produced by Rolls-Royce (1971) Ltd. They have also been fitted retrospectively to existing engines. In addition to the main bearings, lubrication will also be required for the wheelcase, tacho-generator, CSDU, alternator, starter and fuel pump drives.


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Squeeze Film Bearing. Figure 10.1.


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Single Spool Turbojet

Twin Spool Turboprop Engine. Bearing Location Comparison. Figure 10.2.


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10.3 ENGINE LUBRICATION SYSTEMS There are basically two types of lubrication system at present in use in gas turbine engines:a)

Recirculatory. In this system, oil is distributed and returned to the oil tank by pumps. There are two types of recirculatory system:(i) Pressure relief valve system. (ii) Full flow system.

b)

Expendable. The expendable or total loss system is used on some small turbo-jet engines, eg. RB 162 in which the oil is spilled overboard after lubricating the engine.

10.3.1 PRESSURE RELIEF VALVE RE-CIRCULATORY SYSTEM

In the pressure relief valve type of recirculatory lubrication system the flow of oil to the various bearings is controlled by a relief valve which limits the maximum pressure in the feed line. As the oil pump is directly driven by the engine (by the HP spool in the case of a multi-spool engine), the pressure will rise with spool speed. Above a pre-determined speed the feed oil pressure opens the system relief valve allowing excess oil to spill back to the tank, thus ensuring a constant oil pressure at the higher engine speeds. A typical relief valve type of recirculatory lubrication is shown in the figure 10.3. The oil system for a typical turbo-prop engine is similar but, as it supplies the propeller control system, it is more complicated. The oil supply is usually contained in a combined tank and sump formed as part of the external wheelcase. Oil passes via the suction filter to the pressure pump, which pumps it through the air-cooled oil cooler to the pressure filter. A pressure regulating valve upstream of the filter controls the oil pressure. Both oil pressure and temperature indications are transmitted to the cockpit. The oil flows through pipes and passages to lubricate the main shaft bearings and wheelcases. The main shaft bearings are normally lubricated by oil jets and some of the heavier loaded gears in the wheelcases are also provided with oil jets, while the remaining gears and bearings receive splash lubrication. An additional relief valve is fitted across the pump in the lubrication system of some engines to return oil to pump inlet if the system becomes blocked.


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A Pressure Relief Valve Lubrication System for a Two Shaft Turbojet. Figure 10.3. Issue 2 – April 2003

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A Turboprop Full Flow Oil System. Figure 10.4. Issue 2 – April 2003

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10.3.2 RECIRCULATORY OIL SYSTEM – FULL FLOW TYPE

The full flow lubrication system is an alternative to the pressure relief valve oil system and full flow systems are in use as a means of lubricating many modern high power gas turbine engines. The full flow system is similar in many ways to the pressure relief system just discussed – i.e. oil is drawn from a tank by a pump and delivered, via a pressure filter, to various parts of the engine; the oil is then returned by scavenge pumps, via the oil cooler to the tank; also, air is separated from the oil by a de-aerator and centrifugal breather. The major differences from the pressure relief type of recirculatory system are as follows:

The flow of oil to the bearings is determined by the speed of the pressure pump, the size of the oil jets and the pressure in each of the bearing housings.



A metered spill of feed oils is fed back to the tank. This spill is calibrated to match the pump output to ensure that the oil flow to the bearings, via the oil jets, is the same at all engine speeds.



The relief valve in this system is set to prevent excessive oil pressure in the feed side of the system.



A filter by-pass is not normally fitted. The pressure drop across the filter is sensed by a differential pressure switch, any increase in the pressure difference being indicated to give advance warning of a blocked filter.

10.3.3 ADVANTAGES OF FULL FLOW LUBRICATION

The advantages of full flow lubrication are those of suitable oil flow and oil pressure at all engine speeds. A study of the graph will reveal a difference in oil pressure between the pressure relief system and the full flow system and, it will also show that the pressure difference continues throughout the speed range of the engines, with a crossover point at cruising speed. The relief valve system provides too much oil pressure at idle rev/min, but because of the relief valve, the oil pressure is below optimum at maximum engine speed. In contrast the pressure provided by the oil pump of a full flow system rises progressively with increased engine speed and is nearer to the optimum value throughout the rev/min range of the engine.

Comparison of Full Flow and Relief Valve Systems. Figure 10.5. Issue 2 – April 2003

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Full Flow Oil System ( RR Gem). Figure 10.6.


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10.3.4 EXPENDABLE SYSTEM

An expendable system is generally used on small engines running for periods of short duration. The advantage of this system is that it is simple, cheap and offers an appreciable saving in weight as it requires no oil cooler, scavenge pumps or filters. Oil can be fed to the bearing either by a pump or tank pressurisation. After lubrication the oil can either be vented overboard through dump pipes or leaked from the centre bearing to the rear bearing after which it is flung onto the turbine and burnt.

An Expendable Oil System. Figure 10.7. Issue 2 – April 2003

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10.4 MAIN COMPONENTS In any aircraft oil system, we have a number of components that may be thought of as the main components and we have some that are incorporated to safeguard the system (ie. to act as safety devices). The main components, on which the operation of the system depends, include the oil tank, the oil pump and the oil cooler; these are considered in the paragraphs immediately following. The safety devices, which include the various valves and filters, are considered later. 10.4.1 OIL TANK

The oil tank is usually mounted on the engine; it may be a separate unit or part of an external gearbox called the sump. It has provision to allow the system to be filled and drained and a sightglass or dipstick to allow the oil contents to be checked. Usually, the oil level sightglass on the side of the tank is graduated in half-pint or in litre increments, between LOW and FULL marks. The tank is replenished either by pressure or by gravity feed. The pressure filler connection contains a non-return valve and a bayonet adapter to which the oil replenishment trolley pipe is connected.

A de-aerator tray is mounted in the top half of the tank and receives the return oil from the scavenge pumps. The oil in its passage through the system will become aerated and steps must be taken to remove the air. As the oil/air mixture flows over the tray, the oil separates and drains down into the sump, whilst the air is vented to atmosphere.

Typical Oil Tank. Figure 10.8.

Typical Oil Tank Figure 10.8


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10.4.2 OIL PUMPS

The oil pumps fitted in a recirculatory system are normally gear-type or Gerotor type pumps. The pumps are usually mounted in a pack containing one pressure pump and several scavenge pumps. They are driven by a common shaft through the engine gear train. Gear type pumps (Fig.10.10. ) require suitable machining of the gear teeth, or the provision of a milled slot in the casting (adjacent to the delivery side of each pump), to prevent pressure locking of the gears. Gerotor type pumps (Fig.10.11.) use an inner and outer rotor, where the inner rotor is driven by the engine, and the outer rotor which has an extra gear tooth rotates with it. The inner rotor is eccentric to the outer and it is the stepping of the teeth that pumps the oil. The pump also requires kidney shaped slots as inlet and outlet ports. The scavenge pumps have a greater capacity than the pressure pump to ensure complete scavenging of the bearings in a dry sump system. Furthermore, air tends to leak into the bearing housings from the air pressurised seals and this aeration of the oil means that the scavenge pumps have to pump an increased oil/air volume. As we saw in the previous paragraph the air is subsequently removed by the deaerator.

Typical Gear Type Oil Pump. Figure 10.9.


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Gear Type Pump. Figure 10.10.

Gerotor Type Pump. Figure 10.11.


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10.4.3 OIL COOLING

All engines transfer heat to the oil by friction, churning and windage within a bearing chamber or gearbox. It is therefore common practice to fit an oil cooler in recirculatory oil systems. The cooling medium may be fuel or air and, in some instances, both fuel-cooled and air-cooled coolers are used. Some engines which utilise both types of cooler may incorporate an electronic monitoring system which switches in the air-cooled oil cooler (ACOC) only when it is necessary. This maintains the ideal oil temperature and improves the overall thermal efficiency. The fuel-cooled oil cooler (FCOC) has a matrix which is divided into sections by baffle plates. A large number of tubes convey the fuel through the matrix, the oil being directed by the baffle plates in a series of passes across the tubes. Heat is transferred from the oil to the fuel, thus lowering the oil temperature. The fuel-cooled oil cooler incorporates a bypass valve fitted across the oil inlet and outlet. The valve operates at a pre-set pressure difference across the cooler and thus prevents engine oil starvation in the event of a blockage. A pressure maintaining valve is usually located in the feed line of the cooler which ensures that the oil pressure is always higher than the fuel pressure. In the event of a cooler internal fault developing, the oil will leak into the fuel system rather than the potentially dangerous leakage of fuel into the oil system.

Typical Fuel Cooled Oil Cooler. Figure 10.12.


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The air-cooled oil cooler is similar to the fuel-cooled type both in construction and in operation – except, of course, that air replaces the fuel as the cooling agent. On some engines, the airflow through the matrix is controlled by a flap valve, which is automatically operated when the temperature of the return oil rises to a predetermined value. A turbo-propeller engine may be fitted with an oil cooler that utilises the external airflow as a cooling medium. This type of cooler incurs a large drag factor and, as kinetic heating of the air occurs at high forward speeds, it is unsuitable for turbo-jet engines. 10.4.4 PRESSURE FILTER

The pressure oil filter housing contains a wire-wound or mesh, Paper or felt elements and incorporates a by-pass valve. The filter housing can be drained independently of the main oil system. This is done through a drain valve in the housing base. When drained, the filter can be removed for examination, servicing, or replacement, as necessary, without disturbing the rest of the system. Typical pressure filters are illustrated in figure 10.13.

Wire Wound and Paper Type Oil Filters. Figure 10.13.


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Filters are usually fitted with an impending by-pass indicator. This is usually a red pop out indicator which will pop out and stay out it the differential pressure across the filter element exceeds a predetermined value. This value will be less than the by-pass valve value, to allow the filter to be replaced before the filter goes into bypass. The pop out usually has a thermal lock on it, which prevents the pop out extending when the oil is cold and thick.

Filter Bowl with Pop Out Indicator. Figure 10.14.


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10.4.5 LAST CHANCE FILTER

Some of the gears in the gearboxes and also the main bearing of the engine are lubricated through oil jets. These jets are usually protected by thread-type oil filters. These are often referred to as last chance filters. You may also find small mesh filters doing this job.

Thread Type Last Chance Filter Figure 10.15. 10.4.6 SCAVENGE OIL STRAINERS

When the oil has been distributed to all parts of the engine and has done its job, it is returned to the oil tank by either gravity or pressure from the scavenge pumps. Each pump returns the oil from a particular part of the engine and is protected by a coarse filter (or strainer) in the return line. This arrangement protects the pump gears. It also gives an indication of impending component failure if the strainers are examined for metal particles during periodical inspection.


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10.4.7 MAGNETIC CHIP DETECTOR

Magnetic detectors may be fitted into the oil system at various points to collect and hold ferrous debris. They are normally fitted in gearboxes and in the scavenge pump return lines to the tank. The collection of ferrous particles on the chip detector provides a warning of impending (or incipient) failure of a component. Some detectors are designed so that they can be removed for periodical examination without having to drain the oil system; others may be checked externally by connecting a suitable test circuit to the plug; finally, some are connected to a cockpit warning system to give an in-flight indication of failure. The chip detector (see figure10.15.) fits into a self-sealing housing and has a bayonettype fitting for easy removal.

Magnetic Chip Detector. Figure 10.16. 10.4.8 DE-AERATOR

We have already noted that air from the bearing sealing system mixes with the oil and causes frothing. If the air is allowed to remain in the oil it may cause a lubrication failure. To prevent this, a de-aerating device may be installed; this removes air from the oil before the oil is re-circulated round the engine by the pressure pump; the air can be vented to atmosphere via the centrifugal breather. De-aerators are usually tray types fitted in the oil tank or centrifugal type as a separate item.


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10.4.9 CENTRIFUGAL BREATHER

When the oil/air mixture returns to the tank the air is separated by the de-aerator tray and passes through to the gearbox via a vent line. It carries some of the oil with it in the form of a fine mist. The oil/air mist in the gearbox can then pass to the centrifugal breather (see figure 10.17). As the vanes of the centrifugal breather rotate, the oil in the mixture is caught in the vanes and thrown back into the gearbox; the air being vented to atmosphere.

Centrifugal Breather. Figure 10.17. 10.4.10

PRESSURE RELIEF VALVE

The pressure relief valve shown in the figure 10.18. controls the oil pressure at the pre-set value demanded by the system. The valve is normally integral with the pump assembly and protects the system from excessive pressure. It is usually a spring-loaded plate-type valve, and can on some engines provide adjustment of pressure setting.

Simple Pressure Relief Valve Figure 10.18. Issue 2 – April 2003

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It is more usual to find a pressure relief valve that varies the pressure with engine speed or breather pressure. These valves are usually adjustable but usually only effect the max speed oil pressure see Figure 10.19.

Pressure Relief Valve That Uses Breather Pressure to Vary Pressure. Figure 10.19

This type of valve uses the oil system breather pressure and an adjustable spring to balance the oil pressure in the main oil feed line to the engine bearings. Consider Fig. 10.19. With the engine running, the breather pressure plus the spring push the sliding valve to the left and restrict the pump spill back to return. This is balanced by the pressure from the main feed line trying to move the slide valve to the right. Should the pressure in the main feed line fall, the breather pressure and spring will move the slide valve further to the left and restrict the oil spill still further. This will allow more oil to flow to the system, and the oil pressure in the main feed line will increase. The slide valve will then move to the right, and the oil spill to the return will be controlled by the main feed line pressure balancing the spring and breather pressure. 10.4.11

BY-PASS VALVE

This is similar in construction to the normal pressure relief valve just discussed. It is connected in the system in such a way that, should the oil cooler or the pressure filter become blocked (so that the oil flow is restricted), the appropriate by-pass valve will operate to re-route the oil. Although the cooling or the filtering has now been by-passed, oil starvation of the oil bearings is prevented. Pop–out indicators are used to warn of an impending by-pass. Issue 2 – April 2003

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The oil cooler will usually have a thermal by-pass valve which will by-pass the cooler when the oil is cold, thus ensuring that the oil gets up to running temperature quickly. 10.5 INDICATIONS AND WARNINGS Indications and warnings vary from aircraft to aircraft, in both the warnings given and the priority that they are given. 10.5.1 LOW PRESSURE WARNING LAMP

If the oil pressure drops below the safe operating value for the particular system, a pressure-sensing switch will initiate a visual warning; the warning usually consists of a red or amber lamp switching on in the cockpit accompanied by an audio warning. The sensing switch may be a differential pressure switch which senses the pressure difference between the feed oil pressure and the return oil pressure or a simple pressure switch. When the pressure or difference falls below a pre-determined level, the switch operates to activate the warning circuit. To reduce the cockpit noise during taxiing, the audio warning may be inhibited, as engines are often shut down before reaching the stand. Although this system is simple, its warning factor may not be quick enough to prevent serious damage to the engine. This is due to the fact that the warning pressure must be below the normal oil pressure at idle RPM. This means that the engine could be running for some time with a low oil pressure before the warning occurs. To overcome this problem multiple pressure switches are used and activated at differing engine RPM’s. For instance, above 85% RPM the low oil pressure warning will come ‘ON’ at 50 psi, below 85% the warning will come on at 20psi. This is a serious warning and the engine must be shut down as soon as possible. 10.5.2 OIL PRESSURE, TEMPERATURE AND QUANTITY INDICATION

See section 14 engine indications for details of these systems. 10.6 OIL SEALS Oil seals have been covered in section 8. 10.7 SERVICING The engine oil level is usually checked after flight or after an engine run. It is not checked straight after shut-down, as entrained air will give a false reading. It cannot be checked accurately if left too long as the oil may run out of the tank into the gearbox. So it is normally checked between 20 minutes and 2 hours or as defined in the aircraft maintenance manual. The oil system magnetic chip detectors will be checked at the periodicity defined in the maintenance schedule. Spectrometric Oil Analysis Program (SOAP) samples of the oil may be taken when required. Issue 2 – April 2003

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Filters are replaced when required by the maintenance schedule or if the pop out indicator is out.


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11 ENGINE FUEL CONTROL SYSTEMS 11.1 INTRODUCTION The thrust of a turbo jet is controlled by varying the amount of fuel that is burnt in the combustion system and in order to operate the safe temperature limits, the amount of fuel that is burnt must be governed by the amount of air that is available at the time. The air supply is dependent upon the RPM of the compressor and the density of the air at its inlet, so under a constant set of atmospheric conditions, the RPM of the compressor is an indication of the engine thrust. The pilot has control of the fuel flow to the combustion system and is able to select any compressor RPM, between ground idling and the maximum permissible which is required for take off conditions, by the operation of a cockpit lever. In the normal operational environment of an aircraft engine, atmospheric conditions can vary over a wide range of air temperatures and pressures resulting in changes of air density at the compressor inlet. A reduction in air density will cause a reduction in the amount of air delivered to the combustion system at a selected RPM, with a consequent increase in the combustion chamber temperature. If the fuel flow is not reduced, a rise in compressor RPM will occur accompanied with overheating of the combustion and turbine assemblies. An increase in air density will result in an increase in the amount of air delivered to the combustion system at a selected RPM and unless the fuel flow is increased, a reduction in RPM will occur. Changes in air density at the compressor inlet are caused by:a)

Altitude. The density of the air gets progressively less as the altitude is increased, therefore less fuel will be required in order to maintain the selected RPM.

b)

Forward Speed. The faster the aircraft flies then the faster the air is forced into the aircraft intake. A well designed aircraft intake will slow down this airflow, converting its kinetic energy into pressure energy, so that it arrives at the compressor inlet at an optimum velocity (0.5Mach) with an increase in pressure and hence density. This is known as Ram Effect and plays an important part in the performance of a turbo-jet. Within certain limits the greater the ram effect, the greater the air mass flow and more fuel can be burnt at the selected RPM, producing more thrust.

11.2 PURPOSE OF THE ENGINE FUEL SYSTEM The purpose of the engine fuel system is to deliver to the combustion system, in a readily combustible form, the correct amount of fuel over the whole operating range of the engine, under the control of the pilot.


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Block Diagram of a Fuel Control.(JT9D) Figure 11.1.


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11.3 LAYOUT OF TYPICAL SYSTEM COMPONENTS The figure 11.1. illustrates the layout of components of a representative fuel system. 11.3.1 AIRCRAFT MOUNTED COMPONENTS

a)

Fuel Tanks. Stores sufficient fuel for the aircraft’s designed flight duration.

b)

Booster Pump. Ensures a constant supply of fuel at low pressure to the inlet of the engine driven HP Fuel Pump.

c)

Low Pressure Cock. Isolates the engine fuel system from the aircraft fuel system in the event of engine fire or for maintenance.

NOTE: These aircraft mounted components will be dealt with in greater detail during the Aircraft Systems Phase. 11.3.2 THE ENGINE LP FUEL SYSTEM

LP Fuel Pump. Form the LP Cock fuel passes to an engine driven LP Fuel Pump which serves two purposes: a. To boost pressure of the fuel to prevent cavitation of the HP pump. b. To provide means of drawing fuel from the fuel tanks in the event of failure of the fuel boost pump in the tank. These are normally centrifugal type pumps which will boost pressure in the region of 5-10 psi. Fuel/air heat exchanger. To reduce the possibility of low temperatures forming ice, in the fuel heating is applied . Fuel heating is achieved by passing the fuel through a form of radiator which uses hot air (or hot oil) to control and maintain fuel temperature above freezing. LP Fuel Filter. The filter element may be made of felt, paper or in some cases wire wound. Its purpose is to prevent foreign particles from entering the engine fuel system. An indication of the filter ‘clogged’ may be provided on the flightdeck. Not withstanding this a by-pass will be incorporated to ensure that the fuel supply , albeit possibly contaminated is still available. 11.3.3 THE ENGINE HP FUEL SYSTEM

HP Pump. Fuel from the LP Fuel filter passes to the HP pump depending on RPM and FCU in the region of 600-800 psi. This HP fuel is then fed to the fuel control unit (FCU).


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Fuel Control Unit. The FCU will meter the engines fuel requirements based upon a given set of conditions at any given time: a. Throttle position. b. Ambient pressure (Pamb) c. Ambient temperature (T12) d. HP compressor RPM (N2) e. Compressor discharge pressure (CDP) Fuel in excess of that required is returned to the inlet side of the HP pump. Metered fuel is then fed to the flowmeter via a throttles and HP cock. Throttle and HP cock. The fuel control operating levers can be a combined throttle and HP cock lever or separate levers. The position of the throttle lever determines the power required, the HP shutoff cock controls the supply of fuel from the FCU to the burners, when closed the engine will be shut down, when open fuel will be available to the burners. Fuel Flowmeter. The fuel flowmeter will measure the amount of fuel being fed to the burners and relay this information to the flightdeck. A gauge calibrated in either pounds or kilograms will indicate to the operator how much fuel is being consumed an hour. A second window within this gauge may also indicate how much fuel the engine has consumed by the engine during the flight. Fuel/oil Heat Exchanger Similar to the heat exchanger used to heat the fuel, this heat exchanger will use the HP fuel supply to cool the engine oil. Pressurising and Dump Valve. From the fuel/oil heat exchanger HP metered fuel passes to the pressurising and dump valve. It function is to: b. Prevent fuel flowing to the burners during the starting phase until such time as fuel pressure is sufficient to give good atomisation of the fuel thus ensuring good light-up. c. Allow sufficient pressure to build up within the Fuel Control Unit (FCU) servo/hydraulic control systems ensuring correct metering of fuel supply is achieved during starting. d. Enable a rapid dump of fuel remaining in the pipelines to the burners on shutdown.


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Burners. The type of burners employed will vary with design. Two basic types are in common use, atomisers and vaporisers, and their common purpose is to supply fuel in a readily combustible form over the whole operating range of the engine. 11.4 FACTORS GOVERNING FUEL REQUIREMENTS The factors that determine the quantity of fuel that constitutes ‘the correct amount’ to be delivered to the combustion system at any one time are:a)

The RPM selected.

b)

The density of the air at the compressor inlet.

c)

The rate at which the engine can accept the fuel into the combustion system under conditions of engine acceleration.

11.5 REQUIREMENTS OF THE ENGINE FUEL SYSTEM a)

The selection of the RPM must be under the control of the pilot and the system must ensure that the maximum permissible RPM is not exceeded.

b)

The fuel must be introduced into the combustion system in a readily combustible form and the system must be able to automatically adjust the fuel flow to match the air available in order to maintain the selected RPM under all operating conditions.

11.6 ENGINE FUEL SYSTEM COMPONENTS In order to achieve its purpose, the engine fuel system will incorporate the following components:a)

High pressure fuel pump.

b)

Fuel flow-controlling devices.

c)

Burners.

11.7 FUEL PUMPS The type of fuel pump used may vary from one engine type to another and their common purpose is to supply the correct amount of fuel to the burners at a sufficient rate of flow to ensure operation over the whole range of engine operation. The pump is driven by the engine via a suitable gear train. 11.7.1 FUEL PUMP REQUIREMENTS

Because the fuel flow requirements of an engine running at a constant RPM will vary with changing atmospheric conditions, the fuel pump must be capable of delivering fuel at flow rates in excess of the maximum engine demand at any particular RPM, eg. its output must be variable independently of its speed of rotation. The output of the engine driven fuel pump is dependent on engine RPM and controlling signals from various fuel flow controlling devices. Issue 2 – April 2003

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There are two basic types of fuel pump, the plunger-type pump and the constant delivery gear-type pump; both of these are positive displacement pumps. Where lower pressures are required at the burners (spray nozzles), the gear-type pump is preferred because of its lightness. 11.7.2 PLUNGER-TYPE FUEL PUMP

The pump shown in the figure 11.2. is of the single-unit, variable-stroke, plunger type; similar pumps may be used as double units depending upon the engine fuel flow requirements. The fuel pump is driven by the engine gear train and its output depends upon its rotational speed and the stroke of the plungers. A single-unit fuel pump can deliver fuel at the rate of 100 to 2,000 gallons per hour at a maximum pressure of about 2,000 lb/in2. The fuel pump consists of a rotor assembly fitted with several plungers, the ends of which project from their bores and bear on to a non-rotating camplate or swashplate. Due to the inclination of the camplate, movement of the rotor imparts a reciprocating motion to the plungers, thus producing a pumping action. The stroke of the plungers is determined by the angle of inclination of the camplate. The degree of inclination is varied by the movement of a servo piston that is mechanically linked to the camplate and is biased by springs to give the full stroke position of the plungers. The piston is subjected to servo pressure on the spring side and on the other side to pump delivery pressure; thus, variations in the pressure difference across the servo piston cause it to move with corresponding variations of the camplate angle and, therefore, pump stroke.

Plunger Type Fuel Pump or Swash Plate Pump. Figure 11.2. Issue 2 – April 2003

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11.7.3 GEAR-TYPE FUEL PUMP

The gear-type fuel pump (see figure11.3.) is driven from the engine and its output is directly proportional to its speed. The fuel flow to the spray nozzles is controlled by re-circulating excess fuel delivery back to inlet. A spill valve, sensitive to the pressure drop across the controlling units in the system, opens and closes as necessary to increase or decrease the spill.

Gear Type Fuel Pump System. Figure 11.3. 11.8 FUEL FLOW CONTROL Control of the fuel flow to the burners is by two main methods:a)

Manual control by the pilot.

b)

Automatic adjustment of fuel flow to correct for basic engine requirements. (i) Changes in intake pressure. (ii) Excessive fuel to air ratio during engine acceleration. (iii) Additional controlling devices as determined by specific engine requirements.


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11.8.1 BASIC FLOW CONTROL SYSTEM

Principle of Fuel Metering The flow of a fluid through an orifice (jet) depends on the area of the orifice and the square root of the pressure drop across it, ie:Fuel Flow = Orifice Area x Pressure Drop

Principle of Fuel Metering Valve. Figure 11.4. Thus it is possible to vary fuel flow by changing orifice area or the pressure drop across the orifice. In a fuel system the orifice is variable and is in fact the throttle valve. 11.8.1.1

Application to Flow Control System

In the flow control system the fuel flow required to give a selected RPM is selected by throttle area under the control of the pilot (manual control). Compensation for air density variation is superimposed on this selection by the altitude sensing control unit (pressure drop control unit) varying the pressure difference across the throttle valve. 11.8.1.2

Control Principle

The controlling principle of a flow control system is that a constant throttle pressure drop is maintained irrespective of throttle area (position) for a given height and speed.


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Constant Pressure Drop. Figure 11.5. 11.8.1.3

Principle of Flow Control System (See Figure11.6.)

If however, height and speed change, then the altitude sensing unit will vary the pump output and fuel flow (thus throttle pressure drop) by changing the pump output at constant throttle setting.

Principle of Barometric Flow Control. Figure 11.6.


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11.9 HYDRO-MECHANICAL CONTROL UNITS In hydro-mechanically operated flow control units (FCUs), the method of control is to use servo fuel as a hydraulic fluid to vary fuel flow (eg. by varying pump swashplate angle). The pressure of the servo fuel is varied by controlling the rate of flow out of an orifice at the end of the servo line; the higher the outflow, the lower will be servo pressure and vice versa. There are two types of variable orifice: the half-ball valve and the kinetic valve.

Half Ball Valve System. Figure 11.7. 11.9.1.1

The Half-Ball Valve.

In this arrangement, a half-ball on the end of a pivot arm is suspended above the fixed outlet orifice (see figure). Up and down movement of the valve varies servo fuel outflow and thus servo pressure and pump output. 11.9.1.2

The Kinetic Valve. Figure 11.8.

A line containing pump output fuel is so placed as to discharge on to the face of the servo outflow orifice and the kinetic energy so produced restricts servo fuel bleed. A blade can be moved downwards to interrupt the high-pressure flow; this reduces the impact onto the servo orifice, thus causing a greater outflow and a reduction in servo pressure (see figure). The kinetic valve is less prone to dirt blockage than the half-ball type, although it is more complex.


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Condition 1. With the kinetic valve in the open position, the blade separates the opposing flows from pump delivery and the servo cylinder. As there is no opposition to the servo flow, the volume of servo fluid reduces and the piston moves against the spring under the influence of pump delivery pressure. The movement of the piston reduces the pump stroke and therefore it’s output. Condition 2. With the valve fully closed, the kinetic energy of the pump delivery fuel prevents leakage from the servo chamber. Servo fuel pressure therefore increases and, with the assistance of the spring, overcomes the pump delivery pressure, thus moving the piston to increase the pump stroke and output.

Kinetic Valve Figure 11.8. System

Condition 3. Under steady running conditions, the valve assumes an intermediate position such that the servo fuel and spring pressure exactly balances the pump delivery pressure.

11.9.2 BAROMETRIC CONTROLS

The function of the barometric control is to alter fuel flow to the burners with changes in intake total pressure (P1) and pilot’s throttle movement. Several different types of hydro-mechanical barometric control are available. Three of the most common types are described. For simplicity, the description and operation of each type of flow control is related to the half-ball valve method of controlling servo fuel pressure.


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Simple Flow Control. The Simple Flow Control Unit (see figure 11.9.) comprises a half-ball valve acting on servo fuel bleed, whose position is determined by the action of an evacuated capsule (immersed in P1 air) and a piston subjected to the same pressure drop as the throttle valve. Fuel from the pump passes at pressure P pump through the throttle, where it experiences a pressure drop to burner pressure P burner. The response to P1 and throttle variations can now be examined.

Simple Flow Control. Figure 11.9. Throttle Variations. If the pilot opens the throttle, the throttle orifice area increases, throttle pressure drop reduces and therefore PPUMP falls, PBURNER rises and the piston moves down, allowing the spring to lower the half-ball valve against the capsule force, increasing servo pressure and pump output. The increased fuel flow increases the throttle pressure drop to its original value, returning the half-ball valve to its sensitive position. P1 Variations. If the aircraft climbs, P1 will fall, causing the capsule to expand and raise the halfball valve against the spring force. Servo pressure will fall, swashplate angle will reduce and fuel pump output will reduce. The reduced flow will cause a reduced throttle pressure drop. Thus Simple Flow Control keeps the throttle pressure drop constant, regardless of throttle position. At very high altitude the system becomes insensitive and it is not used on large turbo-jets. Nevertheless, it is fitted on the Adour and Dart and has proved to be a reliable and fairly accurate control unit.


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11.9.3 PROPORTIONAL FLOW CONTROL.

The Proportional Flow Control Unit (see figure 11.10.) was designed for use on large engines with a wide range of fuel flow. The problem of accurate control over this wide range was overcome by operating the controlling elements on a proportion of the main flow. The proportion varies over the flow range, so that at low flows a high proportion is used for control and at high flows, a smaller proportion. Fuel passes into the controlling (or secondary) line through a fixed secondary orifice and flows out through another orifice to the LP side of the pump. Secondary flow is controlled via the proportioning valve and sensing valve, which maintains an equal pressure drop across the throttle valve and secondary orifice. Servo pressure is controlled by a half-ball valve operated by P1 and by secondary pressure.

Proportional Flow Control. Figure 11.10.

Throttle Variations. If the throttle is opened, its pressure drop is reduced and the proportioning valve closes until the pressures across the diaphragm are equalised. Thus secondary flow and pressure are reduced, the piston drops, the half-ball valve closes and pump stroke increases. The increased fuel flow increases secondary pressure until the half-ball valve resumes its sensitive position, but the proportioning valve remains more closed than previously, taking a small proportion of the increased flow.


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P1 Variations. Variations in P1 will cause the capsule to expand or contract, thus altering the position of the half-ball valve and altering fuel flow. This tends to cause rapid changes in secondary pressure with resultant instability; damping is provided by the sensing valve, which adjusts to control the outflow to LP, thus damping secondary pressure fluctuations. The valve is contoured to operate only over a small range of pressure drops so that during throttle movements it acts as a fixed orifice. 11.9.4 ACCELERATION CONTROL UNITS

The function of the Acceleration Control Unit (ACU) is to provide surge-free acceleration during rapid throttle openings. There are two main types of hydromechanical ACU in service. The Flow Type ACU. With the flow type ACU (see figure 11.11.) all the fuel from the pump passes through the unit, which compares fuel flow with compressor outlet pressure (P 3), which is proportional to engine speed. The fuel from the pump passes through an orifice containing a contoured plunger; the pressure drop across the orifice is also sensed across a diaphragm. When the throttle is opened, the pump moves towards maximum stroke and fuel flow increases. The increased flow through the ACU orifice increases the pressure drop across it and the diaphragm moves to the right, raising the half ball valve and restricting pump stroke. The engine now speeds up in response to the limited overfuelling and P3 rises, compressing the capsule. The plunger servo pressure drops and the plunger falls until arrested by the increased spring force. The orifice size increases, pressure drop reduces and the diaphragm moves to the left, closing the half-ball valve and increasing fuel flow. Fuel flow will increase in direct proportion to the increase in P3.


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Acceleration Control Using Compressor Discharge Pressure. Figure 11.11.


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The Air Switch. In order to keep the acceleration line close to the surge line, it is necessary to control on “Split P3 air” (a mix of P3/P1) initially and then on full P 3 at higher engine speeds. This is achieved by the air switch (or P 1/P3 switch) shown in the figure 11.12. At low speeds, P3 passes through a plate valve to P1 and the control capsule is operated by reduced, or split P3 until P3 becomes large enough to close the plate valve and control is then on full P3.

Air Switch. Air Switch Figure 11.12.11.12. Figure


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The dashpot Type ACU. The dashpot ACU uses two co-axially mounted throttle valves, The inner one is moved by the pilot, the outer (main) throttle valve will move but is controlled by a dashpot which slows the valve movement down to limit the acceleration fuel flow. When closing the throttle the pilot pushes both sleeves in together.


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ENGINE PROTECTION DEVICES

Described below are typical protection devices that will override any excessive demands made on the engine by the pilot or by the control units. 11.10.1

TOP TEMPERATURE LIMITER.

Turbine gas temperature is measured by thermocouples in the jet pipe. When maximum temperature is reached, these pass a signal to an amplifier, which limits pump stroke by reducing pump servo pressure or moves the throttle valve in series with the pilot. 11.10.2

POWER LIMITER.

A power limiter is fitted to some engines to prevent over-stressing due to excessive compressor outlet pressure during high-speed, low altitude running. The limiter (see figure 11.14) takes the form of a half-ball valve which is opened against a spring force when compressor outlet press (P3) reaches its maximum value. The half-ball valve bleeds off air pressure to the ACU control capsule, thus causing the ACU to reduce pump stroke.

Power Limiter. Figure 11.14.


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OVERSPEED GOVERNOR.

The engine is protected against over-speeding by a governor, which, in hydromechanical systems, is usually fitted on the fuel pump and acts by bleeding off pump servo fuel when the governed speed is reached. On two-spool engines, the pump is driven from the HP shaft and the LP shaft is protected by either a mechanical governor or an electro-mechanical device, again acting through the hydro-mechanical control system. There are two types of pump-driven governors: 11.10.3.1 Centrifugal Governor.

The centrifugal type of governor uses the centrifugal pressure of fuel in radial drillings in the fuel pump rotor to deflect a diaphragm at maximum speed. The diaphragm operates on a half-ball valve to reduce pump servo pressure and thus pump stroke. The disadvantage of this type is that it needs to be reset if fuel specific gravity changes. It is seldom used on modern engines.

Centrifugal Governor Figure 11.15.


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Centrifugal governors using bob weights are used as LP shaft governors on some engines. They will return fuel to low pressure when the LP shaft overspeeds see figure 11.16.

Centrifugal LP Figure 11.16. Governor 11.10.3.2

Hydro-mechanical Governor.

In the hydro-mechanical governor the pump drive shaft rotates a rotor containing a half-ball valve on a lever arm (shown in the figure 11.17.). As engine speed increases, centrifugal force closes the valve, increasing the pressure of fuel in the governor housing (governor pressure) by restricting its flow to LP. When the maximum speed is reached, governor pressure is high enough to deflect a diaphragm, which opens the half-ball valve acting on pump servo. A hydro-mechanical governor does not require adjustment for changes in fuel specific gravity.


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ENGINES

HP Hydro-Mechanical Governor. Figure 11.17.

11.11 11.11.1

BURNERS ATOMISER BURNERS

This type of burner presents the fuel in a finely atomised spray by forcing the fuel to pass through a small orifice. The size of the orifice is critical because it must atomise the fuel effectively over a wide range of fuel flows, from idling to take off RPM. Some engines have such a wide range of fuel flow requirements that a single orifice is unable to perform the task effectively unless extremely high fuel pressures are used and to combat this a burner with two different sized orifices are used. During low fuel flow requirements, only the small or primary orifice is supplied with fuel and at higher flow rates both primary and secondary orifices are in operation.


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Simplex Burner Nozzle Detail. Figure 11.18.

Both types of atomiser burners incorporate an air shroud, which directs some of the primary air into the burner to assist atomisation and to cool the burner head to prevent the formation of carbon. The usual method of atomising the fuel is to pass it through a swirl chamber where tangentially disposed holes or slots impart swirl to the fuel by converging its pressure energy to kinetic energy. In this state, the fuel passes through the discharge orifice where the swirl motion is removed as the fuel atomises to form a cone-shaped spray. The shape of the spray is an important indication of the degree of atomisation; thus, the rate of swirl and therefore the pressure of the fuel at the burner are important factors in good atomisation.


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The simplex burner

A Simplex Burner. Figure 11.19. The Simplex burner shown in the figure 11.19. was first used on early jet engines. It consists of a chamber, which induces a swirl into the fuel and a fixed area atomising orifice. This burner gave good atomisation at the higher fuel flows, that is at the higher burner pressures, but was very unsatisfactory at the low pressures required at low engine speeds and especially at high altitudes. The reason for this is that the Simplex burner was by the nature of its design a “square law” burner, that is the flow through the burner is proportional to the square of the pressure drop across it. This meant that if the minimum pressure for effective atomisation was 30 lbf/in 2, the pressure needed to give maximum flow would be about 3,000 lb/in 2.


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A Duple or Duplex Burner. Nozzle. Figure 11.20. 11.20. Figure



The Duplex burner. The Duplex burner or Duple burner require a primary and a main fuel manifold and have two independent orifices, one much smaller than the other. The smaller orifice handles the lower flows and the larger orifice deals with the higher flows as the burner pressure increases. A pressurising valve may be employed with this type of burner to apportion the fuel to the manifolds (see figure 11.20.). As the fuel flow and pressure increase, the pressurising valve moves to progressively admit fuel to the main manifold and the main orifices. This gives combined flow down both manifolds. In this way, the Duplex and the Duple burner are able to give effective atomisation over a wider flow range than the Simplex burner for the same maximum burner pressure. Also, efficient atomisation is obtained at the low flows that may be required at high altitude. In the combined acceleration and speed control system the fuel flow to the burners is apportioned in the FFR.

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11.11.1.1 The Spray nozzle.

The spray nozzle (see figure11.21.) carried a proportion of the primary combustion air with the injected fuel. By aerating the spray, the local fuel-rich concentrations produced by other types of burner are avoided, thus giving a reduction in both carbon formation and exhaust smoke. An additional advantage of the spray nozzle is that the low pressures required for atomisation of the fuel permits the use of the comparatively lighter gear-type pump.

A Spray Nozzle. Figure 11.21.


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VAPORISING BURNERS

This type of burner presents the fuel in the combustion system in the form of a rich fuel vapour or gas. This is achieved by delivering the metered flow of fuel to “J” shaped vaporising tubes, which protrude into the combustion chamber. The fuel passes down the vaporising tubes in a coarse spray and mixes with the primary air that enters concentrically to the fuel supply pipe. The fuel and air is mixed thoroughly by pins that protrude into the primary airflow and the heat of the flame surrounding the tube causes the mixture to vaporise before it emerges in the combustion chamber. The introduction of the primary air into the vaporising tubes aids the process of vaporisation and also helps to cool the tubes to prevent the formation of carbon. With this type of burner, the flame points towards the incoming airflow and this helps to stabilise the flame in the vaporising tubes, preventing it being blown away by the secondary air, thus allowing a relatively short combustion system.

A Vaporising Combustion Chamber. Figure 11.22.

The advantages of this type are:a)

Pre-vaporising gives complete combustion within a short length of flame tube.

b)

A complete ring of flame around the annular chamber.

c)

Even pressure and temperature around the chamber.


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Start Nozzle System for a Vaporiser Combustion System. Figure 11.23.


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Starting Fuel Solenoid Valve This solenoid valve is fitted on the starting fuel feed line. It is a two-position valve spring loaded to the closed position. During starting, the solenoid is energised and the valve opens. The flow is directed to the check valve. During the starting cycle the solenoid is de-energised and the spring force closes the valve and the fuel flow to check valve is stopped. Check Valve A check is fitted in the starter jet line downstream of the Priming Solenoid Valve to prevent fuel dribbling into the combustion chamber on shut down. It is a spring-loaded valve, which is closed at rest and opens when fuel pressure reaches a pre-determined value. Starter Jets As vaporisers do not atomise the fuel sufficiently for combustion until they become heated, for starting purposes initial heating during start is provided by four jets, two of which are combined with High Energy Igniters. The starter jets ensure that, even at the low flows encountered during start, the fuel is atomised as required for light up. Pressurising Valve A pressurising valve is fitted in the main gallery feed line. It is spring-loaded which functions to build up and stabilise the metering system servo pressures before any flow to the main gallery. Thereby it ensures the correct delivery of fuel to the vaporisers during start. Main Gallery and Vaporisers The main gallery connects with delivery tubes, each feeding one vaporiser head through a distribution orifice. The delivery tubes are fitted in pairs on the combustion chamber outer case. Fuel is mixed with air in the vaporiser tubes. As the mixture passes through the heated tube, the fuel becomes vaporised so that it is delivered in combustible form. A single unit houses the check and pressurising valves. A purge flow tapped upstream the pressurising valve is connected to the check valve via a purge restrictor. This ensures a continued fuel flow through the starter jets to avoid formation of carbon in this area. 11.11.3

COMBUSTION AND AIRFLOW

The addition of fuel to compressor air and the resulting continuous combustion gives a release of heat and an increase in volume, which is converted to an increase in velocity. In the combustion chamber the heat release (combustion efficiency) may be as high as 99%.


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More power and efficiency result from “rich” mixtures, but these are limited by maximum turbine temperatures. Therefore fuel supplies must be limited so that an overall air/fuel ratio of about 60:1 at maximum rpm is achieved. At other rpm the ratio will change due to changing efficiencies of turbine and compressor. The “correct” mixture strength is 15:1 hence only about a quarter of the air passing through the engine is used for combustion. (15% - 25% is the typical range). In the flame area the ratio is about 13:1 and around the flame centre a weaker ratio of 18:1 is used to ensure complete combustion with no carbon formation. The flame rate at an atomising burner is 2-10 ft/sec and at a vaporiser, 60 ft/sec. Both figures are low compared with the air velocity through the combustion zone, hence the requirement for a low velocity zone at the burner to (a) aid ignition and (b) maintain the flame at the burner. Theoretically, combustion in a gas turbine is at “constant pressure”, ie. the pressure along the combustion chamber does not change due to combustion but could alter due to changes in rpm and air intake pressure. In practice the combustion chamber shape affects the pressure and they are designed to minimise this and a drop of 4% along its length is usual. Flame temperature is high; a constant 2,000C at the centre. Flame size, however, can change and the bigger the flame becomes the higher goes Turbine Entry Temperature and Jet Pipe Temperature (TET and JPT). “Over-fuelling” gives a larger flame and “Under-fuelling” a smaller; the significance of these will be seen in a later note.


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ELECTRONIC ENGINE CONTROL SYSTEMS

Advances in gas turbine technology have demanded more precise control of engine parameters than can be provided by hydromechanical fuel controls alone. These demands are met by electronic engine controls, or EEC, of which there are two types: supervisory and full-authority. 11.12.1

SUPERVISORY ELECTRONIC ENGINE CONTROL

The first type of EEC is a supervisory control that works with a proven hydromechanical fuel control. The major components in the supervisory control system include the electronic control itself, the hydromechanical fuel control on the engine, and the bleed air and variable stator vane control. The hydromechanical element controls the basic operation of the engine including starting, acceleration, deceleration, and shutdown. High-pressure rotor speed (N2), compressor stator vane angles, and engine bleed system are also controlled hydromechanically. The EEC, acting in a supervisory capacity, modulates the engine fuel flow to maintain the designated thrust. The pilot simply moves the throttle lever to a desired thrust setting position such as full takeoff thrust, or maximum climb. The EEC adjusts the fuel flow as required to maintain the thrust compensating for changes in flight and environmental conditions. The EEC control also limits engine operating speed and temperature, ensuring safe operation throughout the flight envelope. If a problem develops, control automatically reverts to the hydromechanical system, with no discontinuity in thrust. A warning signal is displayed in the cockpit, but no immediate action is required by the pilot. The pilot can also revert to the hydromechanical control at any time. Electronic Engine Control A typical example of an EEC system is that used in many of the Pratt and Whitney 100 series engines currently in service. A brief explanation of how the system works, both in automatic and manual modes follows.


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Pratt & Whitney 100 Series Fuel Control System Schematic. Figure 20.24.


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Automatic Operation (EEC mode) The EEC receives signals from various sources: a. Power Management Switch, enabling take off thrust, maximum continuous thrust, climb thrust or cruise thrust settings to be selected. b. Engine inlet pressure and temperature. c. Ambient pressure. d. Air data computer inputs. (a computer that senses pitot pressure, static pressure and total air temperature) e. Engine RPMs – N1 and N2. f. Power lever position. (via a potentiometer) g. Failure signals. Based on these input signals the EEC will output command signals to adjust and control: a. The Hydromechanical Fuel Control Unit via a stepper motor which adjusts the throttle metering valve. b. Ignition circuits. c. Bleed valves d. Torque gauge 11.12.2

FUEL CONTROL

11.12.3

GENERAL

The fuel control is provided by the hydro-mechanical unit (HMU) The HMU is supplied by the HP fuel pump and provides the required fuel quantity to the nozzles. In normal operation the fuel control is managed by the Electronic Engine Control (EEC). This enables accelerations and decelerations without engine surge or flame out whatever the displacement sequence of the power lever. The HMU is also mechanically connected to the power lever thus ensuring fuel control in case of failure of the EEC. Hydro-mechanical Unit (HMU) The HMU comprises: A stepper motor controlled by the EEC. A lever which controls fuel shutoff. A lever which controls the fuel flow.


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PW100 Series Fuel System Auto/Normal Figure 20.25. Mode. Issue 2 – April 2003

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Operation The fuel flow supplied to the nozzles is mainly obtained through two valves: a bypass valve a metering valve. The fuel enters the HMU from pump outlet with a constant flow. This flow is split by the bypass valve into two flows, one for the nozzles (via the metering valve) and one bypass return flow to the pump. The position of the bypass valve is a function of the loss of fuel pressure caused by the metering valve. The metering valve is pneumatically actuated. In the pneumatic servo block, the reference pressure is the HP compressor outlet pressure, P3. A controlled reduction of the P3 pressure results in a variable Py pressure which when opposed to a bellows device, moves the piston of the metering valve. The pneumatic servo block is managed: in normal operation by the EEC in manual operation, by the power input lever. Normal Operation (EEC Mode) According to the input data (pressures, temperatures, speeds) and to the commanded power (power lever), the EEC controls a stepper motor located in the HMU. The stepper motor regulates Py pressure thus modulating the fuel flow as requested. A governor acts on the Py pressure, thus setting an NH speed limit function of the compression of a spring by a cam (EEC cam) connected to the power lever. Manual Operation (Manual Mode) Py pressure is not regulated by the stepper motor but by the simultaneous actions of the NH speed governor and the spring, compressed by a second cam (manual cam) connected to the power lever. Transfer from the EEC Mode to the Manual Mode. In normal operation the EEC manages the fuel regulation. The manual operation is automatically connected when the operation in the EEC mode is switched off. A solenoid in the HMU selects the manual cam instead of the EEC cam and cancels the regulation control through the stepper motor. Operation of the HMU in the fail mode In case of failure of the EEC, the position of the stepper motor is "frozen". Whatever the increase of power through the power lever, the last N H speed remains unchanged (the load applied by the spring on the N H speed governor increases).For any power reduction through the power lever, the N H speed decreases according to the curve of the EEC cam (decreasing spring load).


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PW 100 Series Fuel System in Manual Mode. Figure 11.26. Issue 2 – April 2003

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FULL-AUTHORITY DIGITAL ELECTRONIC CONTROL (FADEC)

The supervisory control was a step toward the full-authority, fully redundant EEC. It controls all engine functions and eliminates the need for the backup hydromechanical control used in the supervisory system. The modern full authority EEC is a digital electronic device called a full-authority digital electronic control, or FADEC. One of the basic purposes of the FADEC is to reduce flight crew workload. This is achieved by the FADEC's control logic, which simplifies power settings for all engine operating conditions. The throttle position is used to achieve consistent engine settings regardless of flight or environmental conditions. The FADEC establishes engine power through direct closed-loop control of the engine ratio thrust-rating parameter. The required thrust is calculated as a function of throttle lever angle, altitude, Mach number, and total air temperature. The air data computer supplies altitude, Mach number, and total air temperature information, and sensors provide measurements of engine temperatures, pressures, and speeds. This data is used to provide automatic thrust control, engine limit protection, transient control, and engine starting. FADEC uses a pre-programmed schedule to obtain the correct thrust for the various throttle lever angles, and it provides the correct thrust for any chosen angle during changing flight or environmental conditions. To get the desired thrust, the pilot has only to set the throttle lever to a position which aligns the thrust command from the control with the reference indicator from the aircraft thrust management computer. The control system automatically accelerates or decelerates the engine to the desired level without the pilot having to continually monitor the thrust gauge. Once a power setting has been selected, the FADEC maintains it until the throttle lever position is changed. A constant throttle lever angle setting can be used for takeoff and climb. In addition, since the pilot sets engine thrust , and the system controls the thrust by using a given throttle lever angle, the same thrust rating will be obtained on each engine at the same throttle position. This eliminates throttle stagger. The FADEC has many advantages over both the hydromechanical and supervisory EEC. Some of these are: 

It requires no engine trimming



It ensures improved engine starts



It provides a constant idle speed with changes in atmospheric conditions and changing service bleed air requirements



It saves fuel by providing improved engine bleed air management



It fully modulates the active clearance control (ACC) system (if fitted)



It ensures more repeatable engine transients due to the higher precision of its digital computer


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It provides engine limit protection by automatically limiting critical engine pressures and speeds

A typical FADEC system is that used in some of the Pratt and Whitney 4000 series engines currently in service. A brief explanation of how the system works follows. Fuel Distribution and Control Components (Figure 11.27.) Components controlling and distributing the fuel to the burners include: 

FADEC/EEC



Fuel/oil cooler and by pass valve



Fuel metering unit



Fuel distribution valve



Fuel injector supply manifolds



Fuel injectors

Fuel Distribution System of a FADEC Engine Figure 11.27.


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Fuel Distribution During operation, fuel flows from the aircraft fuel tank to the fuel-pump boost-stage inlet. The pressurised fuel from the boost stage of the engine-driven fuel pump then leaves the pump and is delivered to the fuel/oil cooler, whose purpose is to keep the fuel sufficiently warm to prevent ice from forming in the fuel, and at the same time, keep the maximum temperature of the oil within the correct limits. This engine is also equipped with an air/oil heat exchanger, which uses fan air and 2.5 bleed air to prevent the fuel from getting too hot. From the fuel/oil cooler, the fuel is returned to the fuel pump, where it is filtered and sent to the main pump stage to be further pressurised before it is sent to the fuelmetering unit, which actually does the metering on the basis of information it receives from the FADEC. The fuel-metering unit sends fuel to the fuel-flow transmitter, and then to the fuel distribution valve. (Servo fuel, used as an actuation pressure to some interface components, also comes from the fuel-metering unit.) Bypass fuel not sent to the fuel distribution valve or servo supply is returned to pump interstage flow. From the fuel distribution valve, the metered fuel flows through the fuel manifolds to the fuel injectors. The FADEC is the primary interface between the engine and the aircraft. The FADEC contains two channels that are called "A" channel and "B" channel. Each time the engine starts, alternate channels will automatically be selected. The channels are linked together by an internal mating connector for crosstalk data transmission. Much more is accomplished by this control than simply sending a signal to the fuel-metering unit to establish a fuel flow to the nozzles.


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FADEC Interface with the Aircraft. Figure 11.28.

Interface with Aircraft The FADEC receives several refereed (a validated reference used to confirm correct input) inputs and delivers several outputs. Inputs to the FADEC come from the following: 1. The power levers. Two analogue signals come from each power-lever resolver. (The resolver is an electromechanical device to measure angular movement.) 2. The air-data computers (ADC) in the form of a. Total pressure b. Pressure altitude c. Total air temperature 3. The flight-control computer (FCC) for adjusting the engine pressure ratio (EPR) for all engines as a part of the engine thrust trim system (ETTS). The ETTS logic starts when the engine pressure ratio (EPR) on any two engines is above 1.2. 4. Seven discrete (electrical signals) inputs: a. Pt2/Tt2 probe heat b. Fire c. Alternate mode select c. External reset (fuel-control switch) Issue 2 – April 2003

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d. Bump rate selector e. Maintenance (data retrieval) f. Engine location identification 5. Two sources of 28 VDC power (DC bus and ground test power) Out puts from the FADEC are as follows: 

Engine pressure ratio (EPR)



Low-speed spool (NI). There is a backup N1 speed output from channel "B."



Exhaust gas temperature (EGT)



High-speed spool (N2)

Flap/slat position and weight-on-wheels status is also sent to the FADEC. The flight-control computer (FCC) acts as a backup for the air-data computer (ADC). FADEC Interface with Engine All data input to the FADEC is validated through a series of comparisons and checks .For example, compressor rotor speeds are compared to each other and checked to ensure the proper range (0 -120 percent). Inputs to the FADEC from the engine are as follows: 

N2 rpm, Power comes from the FADEC alternator and is used for limiting, scheduling systems, and setting engine speeds.



N1 rpm, which comes from the FADEC speed transducer (a transducer is a device used to transform a pneumatic signal to an electrical one) and is used for limiting and scheduling systems. It is also used as an alternate mode.



Compressor-exit temperature (Tt 3 ), which comes from the diffuser case, is used to calculate starting fuel flow. • Exhaust-gas temperature (Tt 4.95 ), which comes from the exhaust case, is used for indication.



Fuel temperature (Tfuel), which comes from the fuel pump, is used to schedule the fuel heat-management system.



Oil temperature (Toil), which comes from the main gearbox, is used to schedule the fuel heat-management system and to schedule the integrated drive generator (IDG) oil-cooling system.



Inlet total temperature (Tt 2), which comes from the inlet cowl on the wing engines and the bellmouth on the tail engine. It is used to calculate fuel flow and rotor speed.



Inlet total pressure (Pt 2), which comes from the same sources as Tt 2, is used to calculate EPR.



Exhaust gas pressure (Pt4.95), which comes from the exhaust case, is also used to calculate EPR.


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

The engine electronic control (EEC) programming plug is used to determine the engine thrust rating and EPR correction.



Burner pressure (Pb), which comes from the diffuser case, is used for limiting and surge detection. • Ambient pressure (Pamb), which comes from the inlet cowl, is used to validate altitude and Pt2.

FADEC Interface With Engine. Figure 11.29.

Based on information received from its various sources the FADEC will: 1. Monitor, control and protect:  Anti surge bleed valves/variable stator vanes  Cooling airflows  Engine oil cooling and IDG oil cooling  Nacelle cooling  Fuel heating  Starting  Idle speed  Acceleration/Deceleration Issue 2 – April 2003

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 Stabilised engine operation  Thrust control including overboost  Critical speeds and pressures 2. Improve reliability of the engine by:  A two channel system of control  An automatic fault detection and logic system  An automatic fault and compensation system 3. Make maintenance easier by:  Engine monitoring  Self test  Fault isolation Control Modes The FADEC has two modes for setting the power of the engine. The EPR mode is the rated or normal mode, while the N1 mode is the alternate or fault mode. Normal Mode. When a thrust-level request is made through the thrust lever, the throttle-resolver angle (TRA), input causes an EPR command. The FADEC will then adjust fuel flow so that EPR actual equals EPR command. The normal or rated power levels are 

Maximum power available (takeoff or maximum continuous)



Maximum climb

At approximately 78 degrees TRA maximum power available is calculated by the FADEC. If the altitude is less than approximately 14,100 ft, the FADEC calculates a takeoff power rating. But if the altitude is greater than 14,100 ft, the FADEC calculates a rating for maximum continuous power. At approximately 68 degrees TRA, the FADEC calculates the maximum climb-power rating. To get all other power levels, except idle, it is necessary to set the thrust lever. Alternate or N1 Mode. If the FADEC cannot control in the EPR, or normal mode, it will go to the N1 mode and a fault is enunciated . In the N1 mode, the FADEC schedules fuel flow as a function of the thrust-lever position, and the TRA input will cause the FADEC to calculate an N1 command biased by Mach number, altitude, and Tt 2. In reverse thrust, the FADEC goes to the N1 mode, and N1 is biased by Tt2. Control in the N1 mode is similar to that of a hydromechanical fuel-control system. Moving the thrust lever fully forward will cause an overboost of the engine. N1 mode may be manually selected, but the logic that keeps the thrust at the same level as it would be in the EPR mode is removed. Issue 2 – April 2003

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Parameters Sensed and Controls Actuated by an Electronic Engine Control. Figure 11.30. Faults

The FADEC has dual electronic channels, each with its own processor, power supply, program memory, selected input sensors, and output actuators. Power to each electronic control channel is provided by a dedicated, engine gearbox-driven alternator. This redundancy provides high operational reliability. No single electronic malfunction will cause an engine operational problem. Each control channel incorporates fault identification, isolation, and accommodation logic.


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While electronic controls are highly reliable, malfunctions can occur. A hierarchy of fault-tolerance logic will take care of any single or multiple faults. The logic also identifies the controlling channel, and if computational capability is lost in the primary channel, the FADEC automatically switches to the secondary channel. If a sensor is lost in the primary channel, the secondary channel will supply the information. If data from the secondary channel is lost, the FADEC will produce usable synthesised information from the parameters that are available. If there is not enough data available for synthesising, the control modes switch. For example, if EPR is lost, the engine will be run on its N1 ratings. In the unlikely event both channels of electronic control are lost, the torque motors are spring-loaded to their fail-safe positions. The fuel flow will go to minimum flow, the stator vanes will move to fully open, the air-oil cooler will open wide, and the ACC will shut off. The FADEC includes extensive self-test routines which are continuously actuated. BITE, or built-in test equipment, can detect and isolate faults within the EEC and its input and output devices. The fault words of the control are decoded into English messages by a maintenance monitor, and they identify the faulty line-replaceable unit (LRU). In-flight fault data is recorded so it can be recalled during shop repair. The FADEC is able to isolate problems and indicate whether the fault is within itself or in a sensor or actuator. In the shop, computer-aided troubleshooting can identify a fault at the circuit-board level. EEC Programming Plug The EEC programming plug located on the FADEC "A" channel housing, selects the applicable schedules within the FADEC for the following: 

Engine thrust rating



EPR modification data



Engine performance package



Variable-stator-vane schedule



2.9 bleed-valve thermocouple selection

The EEC programming plug data is input to the FADEC "A" channel, while the "B" channel EEC programming-plug input is crosswired and crosstalked from the "A" channel. During test-cell operation, the EPR/thrust relationship is compared, and the engine gets a correct EEC programming plug. If the FADEC must be replaced, the EEC programming plug must remain with the engine. If the engine is started without the EEC programming plug installed, the FADEC goes to the N1 mode. But nothing will happen with the FADEC operation if the EEC programming plug disconnects in flight.


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EEC Programming Plug. Figure 11.31.


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Pneumatic and Electrical Connectors As shown in Figs:11.32. there are several pneumatic and electrical connectors to the FADEC. The four pneumatic inputs are as follows: 1. Pt 4.95 This input comes from two combination Pt4.95/Tt4.95 probes, located on the turbine exhaust case, and goes to FADEC port "P 5." For all pressure inputs a transducer in the FADEC changes the pressure signal into an electric signal and sends this signal to both channels. 2. Pt 2 This input comes from the Pt2/Tt2 probe located in the inlet duct. 3. Pb This input comes from a static pressure port in the diffuser case to measure burner pressure. 4. Pam-This input comes from two screened static pressure ports located on the inlet cowl outer surface.

FADEC Electrical and Pneumatic Connections. Figure 11.32.


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Alternator. The alternator provides the FADEC with power and an N2 speed signal. It also sends N2 information to the flight deck.

FADEC Alternator Figure 11.33.


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Speed Transducer. The speed transducer supplies the FADEC "A" and "B" channels with the N1 signal by sensing the frequency at which the 60 teeth on the lowpressure compressor/low-pressure turbine (LPC/LPT) coupling pass by them.

FADEC Speed Transducer Figure 11.34. Temperature Probes. A dual-element, alumel-chromel thermocouple, located on the top right side of the fuel pump, provides the FADEC with information relating to fuel heating and engine oil cooling. Oil Temperature Probes. Two other similar devices inform the FADEC about scavenge oil temperature and No. 3 bearing-oil temperature, and provide input for engine oil cooling-system control, oil-temperature warning indication, and IDG oil-cooling override. Tt3 Temperature Probe. This dual-element probe is located on the diffuser case and provides the FADEC with information for heat-soaked engine start logic.


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FADEC Fuel and Oil Temperature Thermocouples. Figure 20.35. Tt14.95 Temperature Probes. Four thermocouples measure EGT and send their signal to the thermocouple junction box and then to the FADEC. The temperature sense is used only for input to the indication system. There is no EGT limiting function in the FADEC. Exhaust Gas Pressure Probes. The two probes measure Pt14.95 pressure, are manifolded together, and send their averaged pressure to the FADEC.


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FADEC T6 Probe and Exhaust Gas Temperature Junction Box Figure 11.36.

FADEC Exhaust Gas Temperature and Pressure Probes. Figure 11.37. Issue 2 – April 2003

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Pt2/Tt2 Probe. The inlet pressure/temperature probe supplies the FADEC with engine-inlet pressure and temperature information. The pressure sensor is a total pressure probe that sends its signal to both FADEC channels. The temperature sensor is a dual-element resistance type. One element sends its signal to the "A" channel, while the other sends its signal to the "B" channel. The probe is continuously electrically heated.

Pt2/Tt2 Probe. Figure 11.38.


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Automatic Turbine Rotor Clearance Control System The automatic turbine rotor clearance control system also known as the turbine case cooling system, controls and distributes fan air to cool and shrink the HPT and LPT cases. This process increases efficiency by reducing turbine tip clearance for takeoff, climb, and cruise operation. The FADEC commands the system operation to a schedule determined by altitude and N2.

Turbine Case Cooling System. Figure 11.39.

Turbine Vane and Blade Cooling System The turbine vane and blade cooling system (TVBCS) optimises engine performance during cruise by controlling 12th-stage cooling airflow to the HPT and LPT areas. This system is also controlled by the FADEC as a function of altitude and N 2. Additionally, the FADEC receives a feedback signal from the TVBCS right valve.


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FADEC Controlled Active Tip Clearance System Figure 11.40.


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Turbine Vane and Blade Cooling System. Figure 11.41.


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A Pressure Control System for a Turbo –Prop Engine (Dart) Figure 11.42


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A Pressure Control System for a Turbo-Jet Engine (Adour). Figure 11.43. Issue 2 – April 2003

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A Proportional Flow Control System (Avon). Figure 11.44. Issue 2 – April 2003

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Combined Acceleration and Speed Control.(Spey & Tay). Figure 11.45.


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Combined Speed and Acceleration Control with Air Bleed Control. (ALF502.) Figure 11.46. Issue 2 – April 2003

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Figure 11.47. Issue 2 – April 2003

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12 AIR SYSTEMS 12.1 INTRODUCTION In the working cycle and airflow section we discussed the main airflow and working cycle of a gas turbine engine and found that a major function of the airflow through the engine was to act as a cooling medium and that only a small proportion of the air was used to support combustion. In fact, because of the intense heat developed, gas turbine engines only became practical power units when it was discovered that the airflow could be used to ‘insulate’ the structural materials and thus provide acceptable working temperatures for the materials. Many parts of the engine, made from light alloy or ferrous metals, have to be protected from the very high temperatures. To achieve this, an efficient and effective cooling system is needed and this is provided by ducting cooling air from the main gas stream.

Internal Cooling Air Flow. Figure 12.1. In addition to its function of cooling, the airflow is also used to pressurise oil seals and bearings to prevent oil leakage. We thus have the two functions of cooling and sealing to consider. In general, independent airflow’s are taken from the engine compressors to provide:

Low pressure for sealing.

 

Intermediate pressure air for some cooling functions. High-pressure air for the remainder of the cooling functions.

These are considered in the paragraphs that follow.


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12.2 INTERNAL COOLING AIRFLOW Because of the different design features of different gas turbine engines, the cooling airflow varies considerably from one engine type to another. However, the basic principles remain the same and can be explained by using an example. Figure 12.1. shows the cooling and sealing airflow of a two-spool, low ratio by-pass engine. To show the cooling airflow more clearly, the by-pass and main air-stream air paths have been omitted. A study of the figure will show that air is supplied from the low-pressure compressor and also from the high-pressure compressor. This gives the range of pressures required, as mentioned in the previous paragraph. After doing its job, the air is either vented directly to atmosphere or fed into the exhaust gas flow. 12.2.1 LOW PRESSURE AIR

Air is taken from the low-pressure compressor outlet and ducted through the engine to become both a sealing and cooling airflow. This airflow:

Pressurises the main bearing oil seals to prevent oil leakage.



Provides cooling for the low-pressure compressor shaft, the front half of the high-pressure compressor shaft and the turbine shaft.

12.2.2 INTERMEDIATE PRESSURE AIR

This airflow is taken from an intermediate stage of the high pressure compressor and passes through transfer ports to cool the rear half of the high pressure compressor shaft and also the rear face of the last disc of the compressor; it then flows outwards through tubes to mix with the by-pass airstream. 12.2.3 HIGH PRESSURE AIR

This airflow is taken from the high-pressure compressor outlet and is ducted to all faces of the turbine discs to maintain the temperature within the required limits. The pressure of the cooling air is greater than that of the hot gases and since the air is directed outwards across the faces of the turbine discs, it prevents the hot exhaust gases flowing inwards across the discs. Overheating of the turbine discs is thus prevented.


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12.2.4 DIFFERENTIAL PRESSURE SEALS

We know that we require high pressure cooling air at the turbine discs (to reduce the flow of hot exhaust gases across the discs) and low-pressure air at bearing seals (to prevent leakage of oil without undue aeration of the oil). The air at these different pressures must be prevented from mixing and thus, becoming equalised in pressure. This is done by inserting differential pressure seals at appropriate points in the system; these seals are of a multi-groove rotating type. 12.3 SEALING Air at low pressure is used to seal the main shaft bearings and prevent oil from leaking into the engine casing. For effective sealing, the air pressure must always by greater than that of the oil. However, it must not be too much greater, otherwise an excessive amount of air will enter the oil system. De-aeration by means of the de-aerator and the centrifugal breather (see lubrication) may then become difficult. Figure 12.2. shows that the mechanical seals used in air pressure oil sealings are designed to reduce clearance to a minimum; air is fed into the seal at the end remote from the oil feed.


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Air and Oil Seals. Figure 12.2. Issue 2 – April 2003

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12.4 COOLING. Figure 12.3. illustrates the turbine cooling airflow of a typical gas turbine engine. The outward flow of cooling air is controlled by air seals of multi-groove construction and the arrangement is such that the turbine discs obtain the maximum possible cooling from the airflow. Interstage seals are incorporated and they are made in such a way that the front sections provide less restriction to the passage of air than the rear sections do. The result is that the rate at which the cooling air flows across the seals is sufficient to prevent any inward flow of hot gases. The front face of each disc receives a greater airflow than the rear.

Turbine Cooling Airflow. Figure 12.3.


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High pressure cooling air is also directed to the engine’s nozzle guide vanes and turbine blades. These components, which are externally heated by the high temperature gas stream, are cooled by ducting air through air passages formed inside the items themselves. After completing its task, the air is exhausted into the engine exhaust gas flow and thence to atmosphere.

Nozzle Guide Vane Cooling Air. Figure 12.4.


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HP Turbine Nozzle Guide Vane Cooling

LP Nozzle Guide Vane Cooling Nozzle Guide Vane Cooling. Figure 12.5.


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Development of Turbine Blade Cooling. Figure 12.6.


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12.5 TURBINE CASE COOLING – DESCRIPTION AND OPERATION 12.5.1 PASSIVE CLEARANCE CONTROL SYSTEM. FIGURE 12.7.

Compressor discharge air and HP compressor air provide cooling airflow to protect the turbine casing against rapid temperature changes. The stationary parts in the high-pressure turbine section expand and contract more rapidly than the rotor due to pressure and temperature changes. The rotor also has a radial expansion due to rotational speed. The turbine casing incorporates temperature controlled casing flanges with cooling air passages for the passive case clearance control system. The cooling air controls the expansion and contraction of the case to match the rotor and thus maintain desired clearances throughout all temperature ranges and operating conditions.

Cooling Air tubes (Bird Cage) Figure 12.7. Figure 12.7. shows (highlighted) air tubes (Bird Cage) that cools the HP and LP turbines. The air is taken from just aft of the fan and ducted through the cowls (not shown).


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12.5.2 ACTIVE CLEARANCE CONTROL SYSTEM. FIGURE 12.8.

The system provides fan discharge air for cooling the core compartment and the low-pressure turbine case. At low altitudes the core engine requires more cooling and the LPT case requires less cooling to prevent rub. At high altitude the core requires less and in the LPT core requires more to close clearances. By means of a Y manifold and two shut-off valves, cooling air can be selectively directed to the core compartment or to the LPT case. The valves are not positively shut, but permit a required minimum flow at all altitudes and when activated added flow is directed. The valves are controlled by an altitude sensor which activates the core compartment valve below 19,000 feet +5000 feet and the LPT case valve above 19,000 feet +5000 feet. Increased cooling airflow causes the cases to cool and shrink. closes blade tip to case clearances producing improved efficiency.

This shrinkage

LPTACC (Low Pressure Turbine Active Clearance Control). HPTACC (High Pressure Turbine Active Clearance Control). Active Tip Clearance Control. Figure 12.8.


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12.5.3 LOW PRESSURE TURBINE CLEARANCE CONTROL VALVE

Operation At take-off and low altitude the valve is in its normal closed position allowing cooling airflow to the core compartment. When an altitude of 19,000 feet +5000 feet is reached, the altitude sensor switches to supply compressor discharge pressure to the signal port of the valve, causing the valve piston to move to the open position, thus allowing cooling airflow to the low pressure turbine cooling manifold. During descent, at approximately 15,000 feet +1500 feet, the altitude sensor switches back and cuts off the compressor discharge signal pressure to the valve and the positioning spring in the valve returns the piston to its normal closed position. Operation can be monitored by the electrical position indicator switch and a disagree flightdeck light.

Active Tip Clearance Details. Figure 12.9.


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Active tip Clearance control. Figure 12.10.


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12.6 EXTERNAL COOLING 12.6.1 EXTERNAL SKIN OF AERO-ENGINE.

Cooling of the external skin of an aero-engine is achieved by suitable design of the aircraft airframe; the layout will depend upon where the engine is fitted and what kind of engine compartment is used. Normally, the cooling and ventilating of an engine bay or pod is achieved by ducting atmospheric air round the engine and spilling it back to atmosphere through suitably placed outlets (see figure 12.11.). The air is usually taken from a ram inlet but provision is also made to provide a cooling and ventilating airflow during ground running periods. Another function of the cooling airflow is to remove flammable vapours from the engine compartment to reduce the fire risk.

External Cooling. Figure 12.11.


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12.6.2 COOLING OF ACCESSORIES

A number of aircraft accessories produce sufficient heat in normal use to require a cooling system to prevent overheating. A good example is the aircraft electrical generator, which produces considerable heat under normal operating conditions. Such accessories can be cooled by ram airflow when the aircraft is flying, but will require an alternative cooling airflow when the aircraft is on the ground. For ground running and taxiing, the generator for example, is cooled by an airflow that is taken from the engine compressor. This air is blown through nozzles to produce a venturi effect area of low pressure. The low pressure then induces a continuous cooling flow of atmospheric air through the normal ram air passages. This is adequate for cooling most accessories during ground running. Figure 12.12. illustrates a generator cooling system. These are sometimes referred to as ejectors or eductors

HP Air Powering a Jet Eductor to Draw Air Through a Generator at Low Speed. Figure 12.12.


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12.7 HP AIR FOR AIRCRAFT SERVICES. Air is drawn from the compressor at various places to provide air for Airframe needs such as cabin pressurisation and wing and tail anti/de ice. It can also be used within the fuel control system to meter fuel, and in the compressor bleed valve system to control the bleed valves. It can provide heating air for fuel heaters and muscle air to drive air motors in pumps (both for the engine and the airframe) and it can power thrust reversers.

External Air System Schematic.(JT9-D) Figure 12.13. 12.7.1 EXTERNAL AIR TAPPINGS

Engines vary as to the number of external air tappings and their usage. The following notes are taken from the Pratt and Whitney JT9D but have been simplified to provide a more generic coverage. 12.7.1.1

Fan Air

Utilised for the pre-cooling of air conditioning air, cooling the ignition system and on some engines, the Passive and Active tip clearance control. 12.7.1.2

HP Compressor – IP Air (8th and 9th Stage)

Utilised for pneumatic cabin bleeds at concise RPM’s on the JT9D, this can also supply air for nose cowl anti-icing on other engines. The nose cowl anti-icing may have a separate manifold from another compressor stage.


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Pressure Relief

Should the high pressure stage bleed valve fail in the open position, a pressure relief valve is provided to protect the pre-cooler from over-pressure damage. The valve normally would include a pressure switch connected to a PRESS RELIEF warning on the pneumatics display on the flight deck. The operating pressure would be in the region of 100 psi. If the valve opens the vented air escapes through a spring-loaded door on the cowl (blow out panel). 12.7.1.4

Temperature Control

The system normally consists of a pre-cooler temperature sensor and controller, pre-cooler and control valves. This system stabilises the air going to the airframe system, by keeping it constant at a value that the engine can achieve at all power settings. The valves are normally part of the pre-cooler and flow of the fan air is regulated by the opening or closing of the valves. When temperature at the bleed air outlet of the pre-cooler exceeds its limit (160180C) the pneumatic pressure is vented from the actuators to move the cooling air valves toward the open position.


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Bleed Air Temperature Control Valves. Figure 12.14.


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12.8 ANTI-ICING SYSTEMS Generally on gas turbines the engine anti-icing system prevents the formation of ice in the engine intake and on the aircraft structure by the circulation of hot air from the engine. It is normally taken at a midway point along the HP compressor at an approximate temperature of 300C and controlled by a switch on the flight deck. Air is taken via the control valve mounted near the manifold on the HP compressor and directed to an annular manifold around the air intake casing, then through hollow intake guide vanes, tangential struts and nose cone exhausting into the airstream or, as in the case of large fan engines, directly overboard. Control of the nacelle anti-ice system is by means of flight deck switches. These valves may fail safe, i.e. to the open position, if electrical power is lost. On some systems a tapping of hot air also feeds the intake pressure probe.

Typical Intake Anti Ice System. Figure 12.15.


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Intake Anti Ice Control Panel. Figure 12.16.

Spinner Anti Icing Thermal anti-icing of the spinner is often provided by using hot oil. Ice formation can also be minimised by the shape of the spinner and a flexible rubber coating which tends to shed any ice that forms. On a large number of turbo fan engines there are no support struts to the spinner, which rotates with the fan. Thermal anti-icing of the spinner is often provided by using hot oil. Ice formation can also be minimised by the shape of the spinner and a flexible rubber coating which tends to shed any ice which forms.


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Intentionally Blank


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13 STARTING AND IGNITION SYSTEMS 13.1 BASIC PRINCIPLES OF GAS TURBINE ENGINE STARTING SYSTEMS 13.1.1 PURPOSE

The purpose of a gas turbine engine starting system is to: a. carry out a normal ground start. b. relight the engine should flame out occur during flight. c. enable certain components of the system to be isolated for ground servicing purposes (eg. wet runs and dry runs). 13.1.2 ESSENTIAL STARTING REQUIREMENTS

In order to effect a start, the engine must be supplied with:a. Air b. Fuel c. Ignition 13.1.2.1

Air Supply.

The air supply is provided from the engine compressor which must be accelerated from rest to self sustaining rpm by means of a starter motor. In flight the engine may be “Windmilled” by the forward speed of the aircraft, this has to be within an envelope of speed, where the engine rotation is fast enough for the engine to start and not so fast that the flame will be blown out by the airflow. 13.1.2.2

Fuel Supply.

The fuel required for starting is supplied from the normal engine fuel system. It is usually initiated by the pilot opening the HP cock at around 10% HP Compressor speed. If vaporiser type burners are used, the fuel is supplied in the initial stages of starting via a starting solenoid valve and starting atomisers. Once the fuel has been ignited and the vaporisers are heated, the solenoid valve closes to divert the fuel to the vaporiser tubes, normal combustion takes place and fuel supply to the starting atomisers ceases. 13.1.2.3

Ignition.

Ignition of the air fuel mixture is provided by high energy plugs fitted in the combustion chambers. They are positioned close to the fuel spray and operate for a timed period during the starting cycle. HE Ignition units supply the high energy electrical supply to the ignitor plugs. The same ignitor plugs are used to provide relight (restarting) in the air and also as continuous ignition for operation when rain, snow or standing water is present and may cause the engine to flame out. The figure 13.1. illustrates a typical starting sequence applicable to most gas turbines. Issue 2 – April 2003 Page 13-1

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Typical Engine Start Sequence. Figure 13.1.

13.2 STARTER MOTORS There are a number of basic types of starter motors:a. Electric starters. Electrical starter generator. b. Turbo starters (Air Starters). You may also hear of other starter systems such as cartridge and AVPIN starters, these are explosive starters which were once common on older military aircraft. They were never used on commercial aircraft and therefore will not be covered in this book. Gas turbine starters, where a small gas turbine engine like an APU directly drive the engine to start the main engine have been used, but again it is unlikely that you will come across them. Hydraulic starters where hydraulic pressure is applied to one of the hydraulic pumps to drive it as a motor can be used, again usually in military applications.


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13.2.1 ELECTRICAL STARTER MOTOR

Typical Starting Control System. Figure 13.2.

This usually consists of a heavy duty, compound wound, DC motor, which draws its electrical supply from an external source. The motor works in conjunction with a starter control panel, the sequence of events during a start being precisely controlled. To allow the starter motor to overcome the initial inertia of the rotating assembly, the supply to the motor is via a series of resistors, this allows the motor to build up to full speed gradually, reducing the chance of failure within the drive system. The drive from the starter motor to the engine is through suitable reduction gearing and some form of clutch is fitted to disengage the drive when the engine is running. The start master switch does not just switch the starting system ‘ON’. On some aircraft will prepare the aircraft electrical system for the start operation i.e. starter motors require a very high current for starting which is usually too much for a single Transformer rectifier (TRU), so it will parallel the DC systems. To ensure that a start is not carried out on a single TRU, it will place all the AC power systems onto one generator, so if it fails the start is aborted. It will also ensure that the engine gauging systems are all powered for the start in all conditions. 13.2.2 ELECTRIC STARTER/GENERATOR

On some smaller aircraft (eg. Jetstream), an electric starter/generator is employed. The starter /generator initially functions as a starter. When the engine is running it automatically becomes a generator. The drive is through a suitable reduction gearing hence there is no requirement for any form of clutch. Main advantage is the reduction in weight. Issue 2 – April 2003

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13.2.3 SAFETY INTERLOCKS

On some helicopter electric starting systems, a series of safety interlocks are incorporated in the control circuit. The purpose of the interlocks is to prevent the starter relay from closing should an unsafe condition exist.

A Typical Electric Starting System. Figure 13.3.


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13.2.4 AIR TURBO STARTERS

Sources of Air Supply. The air starter can be supplied with air from one or more of the following sources:a. Ground air starting trolley. b. Airborne auxiliary power unit (APU). c. Air from another engine (multi-engined aircraft). d. Air cylinders. 13.2.4.1

Operation.

Air is supplied to the starter via an electrically operated air valve. This is controlled by the starter control unit and is activated by pressing the starter button in the flightdeck. The air is fed to a manifold around the turbine and then directed onto the turbine blades by nozzles or guide vanes. The turbine revolves at very high speed and through reduction gearing and a one way clutch (sprag) mechanism, drives the engine compressor rotor. After a timed period of operation, the control unit closes the air valve. The starter is often mounted on the external gearbox.

An Air Start System. Figure 13.4.


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An Air Starter Figure 13.5. 13.2.4.2

Sprag Clutch

Sprag clutches are used to provide the disconnect mechanism between the starter motor and the engine. The clutch will transmit drive from the starter motor, but will disconnect the drive when the engine speed exceeds the starter. The clutch consists of two smooth concentric drive faces and between them a cage containing many elongated figure of eight shaped cams called “sprags”. All the surfaces are hardened to reduce wear, and are lubricated by oil. The sprag are spring loaded in contact with the starter drive so that when the shaft starts to rotate the sprags stand up and contact the engine drive due to the cam action of their shape. See Figure 13.6. As engine RPM accelerates its drive will be faster than the starter motor and the clutch will automatically dis-engage as sprags get pushed back to their minimum height position. Sprag clutches are used on most types of starter motor or in drives where one way drive is required such as helicopter gearboxes.


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Sprag Clutch. Figure 13.6. 13.2.4.3

Speed Switch

The speed switch can give warning of an over-speed of the starter (engine driving starter) and/or an auto shut-down. As the starter speeds up towards an over-speed, the ball weights centrifuge out forcing up the bell housing breaking the micro-switch to give an over-speed signal.


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13.3 A300 STARTING SYSTEM The following example of an engine start is taken from the training manuals for an A300-134 fitted with GE 6-50 engines. 13.3.1 GE 6-50 STARTING PROCEDURE

The engines are equipped with air starters. The air to start the engine is provided by:

The APU, the ground connectors, or the other engine, if it is already running.

The starting system has provision for:

Engine start.



Engine crank.



Continuous ignition.

The A300 Starting System –Simplified Figure 13.8.

13.3.1.1

The control panel

The control panel is located on the overhead panel. Figure 13.9. shows the start panel with, at the top, the ignition selector which controls the two ignition systems of each engine. The selector has three positions: CRANK in the vertical position, then ground START ignition A or B when turned to the left and continuous RELIGHT when turned to the right. At the bottom of the panel is the master switch with ARM and START/ABORT positions. Finally on each side, one yellow push-to-start button for each engine with its corresponding start valve position light, which is blue and is marked OPEN. The ignition system is supplied by two different electrical circuits.


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Engine Start Panel Figure 13.9. 115 VAC is used to energise the exciter and is controlled through the HP fuel shut off valve lever, the ignition selector and the ignition relay. The ignition relay is energised by 28 Vdc when the master switch is in the ARM position and the start button is pushed. Starting is achieved in the following manner:Set the ignition selector to A or B. Set the master switch to “ARM”. This arms the ignition circuit and closes the air conditioning system if it is open. The amber lights in the push-to-start buttons will illuminate during this transit. When the air conditioning valves are closed, the lights in the push-to-start buttons extinguish and the operator can push the start button which will latch. This increases the APU rpm to 100% to provide sufficient air for starting. It also arms the ignition circuit and finally, provided that pneumatic power is available, it opens the start valve and the blue OPEN light illuminates.


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When the Start Button is Pressed the APU goes to 100% Figure 13.10.

When engine N2 reaches 10% the HP Fuel Shut Off Valve must be opened.

At 10% N2 the HP Fuel Valve is opened. Figure 13.11.

This supplies fuel to the engine and energises the ignition exciters. The engine should light up and EGT should increase. Issue 2 – April 2003 Page 13-10

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When N2 reaches 45% the engine will be self-sustaining so the ignition is switched off, the push-to-start button pops out and the APU demand goes back to normal. Engine rpm should now increase to Ground Idle, which is approximately 65% N 2 and 24% N1.

At 45% The Starter Sequence is cancelled. Figure 13.12.


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13.4 IGNITION SYSTEMS 13.4.1 HIGH ENERGY IGNITION UNIT 13.4.1.1

Basic Operation

The outline of a high energy ignition system is illustrated in the figure. Each high energy ignition unit has a low voltage supply which is controlled by the control unit in the starting system. Depending upon the engine and installation, the supply voltage may be either direct current (DC) or alternating current (AC). If the supply is DC, either a trembler mechanism or a transistor generator is used to convert the dc input to low voltage ac. Thereafter, the operation is the same as that of the system supplied with AC:

The low value of AC is stepped up to a high value by a transformer.



The high value alternating voltage is then ‘rectified’ to provide a high value of DC voltage which is used to charge a capacitor.



DC Ignition Unit Block Diagram. Figure 13.13. When the capacitor voltage is high enough, it breaks down a discharge gap and the discharge is applied to the igniter plug where the energy (high voltage, high current) is converted to a spark across the face of the igniter plug.

13.4.1.2

Construction

A modern transistorised version of a high-energy ignition unit is illustrated in figure 13.14. Although the construction varies according to the type of ignition unit, the basic operation is as described. A choke is fitted to extend the duration of the discharge and safety resistors are fitted to ensure dissipation of energy in the capacitors. 13.4.1.3

Lethal Warning

The electrical energy stored in the HE ignition unit is potentially lethal and, even though the capacitor is discharged when the electrical supply is disconnected, safety precautions are necessary. Before handling the components, the associated circuit breaker should be tripped, or the fuse removed. Never rush in; at least one minute must be allowed between disconnecting the power supply and touching the ignition unit, HT lead or igniter plug. Issue 2 – April 2003

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Transistor generator

A Transistorised Ignition Unit. Figure 13.14.


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13.4.2 IGNITER PLUG

There are two basic types of igniter plug; the constricted or constrained air gap type and the shunted surface discharge type. (fig. 13-15) The air gap type is similar in operation to the conventional reciprocating engine spark plug, but has a larger air gap between the electrode and body for the spark to cross. A potential difference of approximately 25,000 volts is required to ionise the gap before a spark will occur. This high voltage requires very good insulation throughout the circuit. The surface discharge igniter plug has the end of the insulator formed by a semiconducting pellet which permits an electrical leakage from the central high tension electrode to the body. This ionises the surface of the pellet to provide a low resistance path for the energy stored in the capacitor. The discharge takes the form of a high intensity flashover from the electrode to the body and only requires a potential difference of approximately 2000 volts for operation. The normal spark rate of a typical ignition system is between 60 and 100 sparks per minute. Periodic replacement of the igniter plug is necessary due to the progressive erosion of the igniter electrodes caused by each discharge. The igniter plug tip protrudes approximately 0.1 inch into the flame tube. During operation the spark penetrates a further 0.75 inch. The fuel mixture is ignited in the relatively stable boundary layer which then propagates throughout the combustion system.

Ignitor Plugs Figure 13.15. 13.4.3 SERVICING THE IGNITION SYSTEM

Before any servicing is carried out on an ignition system, you must read the relevant Safety Notes together with the Maintenance Manual relating to this work. You must, in particular, understand the lethal warning notice regarding handling high energy ignition equipment and the safety precautions you are to observe. Issue 2 – April 2003

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14 ENGINE INDICATION SYSTEMS 14.1 INTRODUCTION. Engine indications are very important to the crew of a powered aircraft, as they indicate one of the primary parameters needed for flight. There are three types of indications: 1.

Performance indications such as thrust (Engine Pressure Ratio EPR) and Revolutions Per Minute (RPM).

2.

Operation indications such as Turbine Temperature indications, fuel flow, oil pressure and temperature. Discrete indications which put ‘ON’ a warning annunciator such as low oil pressure, fuel low pressure engine overspeed etc.

3.

The engine instruments on most modern commercial aircraft will invariably be located on the main instrument panel in the centre, so that they are visible to both pilots. The instruments are laid out in a logical pattern so that the main thrust indicator is at or near the top of the indications. The indicators will be in vertical columns for each engine and like indicators in rows. When a flight engineer is carried he will have a panel with some of the primary indications and all of the secondary and discrete indicators. He may also have a duplicate set of thrust levers so that he can trim engines when required. Until fairly recently the majority of aircraft used analogue gauges (sometimes referred to as clockwork gauges) These had moving pointers or strips which indicated the parameter being monitored. The modern trend is to replace the analogue instruments with electronic instruments that use LED, liquid crystal or cathode ray screens to display the engine parameters, often not displaying continuously all the information, but to highlight when a fault has occurred or when asked for by the crew. These types of instrument do not usually retain the last indication after an accident, however the electronic box powering them will inform the flight data recorder and/or retain the information in its own memory.


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An Analogue Engine Indication Panel Figure 14.1.


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14.2 ENGINE SPEED INDICATORS. All engines have their rotational speed (R.P.M.) indicated. On a twin or triple-spool engine, the high pressure assembly speed is always indicated; in most instances, additional indicators show the speed of the low pressure and intermediate pressure assemblies. Where Engine Pressure Ratio (EPR) is not indicated then Low Pressure RPM is indicated as this can be corrected to give Thrust. Engine speed indication can be electrically transmitted from a small tacho-generator, driven by the engine, to an indicator that shows the actual revolutions per minute (r.p.m.), or a percentage of the maximum engine speed (fig. 14.2.). The engine speed is often used to assess engine thrust, but it does not give an absolute indication of the thrust being produced because inlet temperature and pressure conditions affect the thrust at a given engine speed. The tacho-generator supplies a three phase alternating current, the frequency of which is dependent upon engine speed. The generator output frequency controls the speed of a synchronous motor in the indicator, and rotation of a magnet assembly housed in a drum or drag cup induces movement of the drum and consequent movement of the indicator pointer.

A Tachometer Generator and Indicator (RPM). Figure 14.2. Issue 2 – April 2003

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Simplified Tachometer System. Figure 14.3.

ROTOR (SYNCHRONOUS WITH SQUIRREL CAGE START)

ROTOR (MAGNET)

N

N S

S

GENERATO R

INDICATO R

Schematic Circuit Diagram of Tachometer System. Figure 14.4.


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Tachometer generator systems have largely been replace by speed probes. A variable-reluctance speed probe, in conjunction with a phonic wheel, is used to induce an electric current that is amplified and then transmitted to an indicator (fig. 14.5.). This method can be used to provide an indication of r.p.m. without the need for a separately driven generator, with its associated drives, thus reducing the number of components and moving parts in the engine.

A speed Probe and Phonic Wheel. Figure 14.5 The speed probe can be positioned on the compressor casing in line with the phonic wheel, which can be a machined part of the compressor shaft. A gear wheel in an external gearbox can also be used. The teeth on the periphery of the wheel pass the probe once each revolution and induce an electric current by varying the magnetic flux across a coil in the probe. The magnitude of the current is governed by the rate of change of the magnetic flux and is thus directly related to engine speed. On some engines one of the teeth is bigger than the others, and will give a bigger response. This can be used for Fan blade balancing or synchronising and/or synchrophasing. SQUARER

MAXWELL BRIDGE

TACHO CIRCUIT

GAUGE

DC  N

Block Diagram of Pulse Probe Circuit. Figure 14.6.


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In addition to providing an indication of rotor speed, the current induced at the speed probe can be used to illuminate a warning lamp on the instrument panel to indicate to the pilot that a rotor assembly is turning. This is particularly important at engine start, because it informs the pilot when to open the fuel cock to allow fuel to the engine. The lamp is connected into the starting circuit and is only illuminated during the starting cycle.

Eddy Current Fan Speed Sensor Figure 14.7. A variation of this system uses an eddy current sensor on the fan casing that senses the fan blades rotating (see figure 14.7.). Sensors similar to these can be used for active tip clearance control, where it senses the gap between the casing and the blade. Modern speed gauges usually have an analogue type display, i.e. a pointer, and also a digital readout below the pointer axis. A target speed indicator is usually fitted which on a clockwork gauge is a pointer outside the numbers, and on an electronic gauge as a coloured marker, this usually has a digital readout of its set position within the gauge normally above the pointer axis.


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14.3 THRUST INDICATION 14.3.1 ENGINE PRESSURE RATIO.EPR.

Although rpm gauges give an indication of the rotational speed of the compressor, they do have one drawback. They do not normally indicate the thrust output or power output of the engine. Any distress within a compressor may cause the engine to have a reduction in thrust output. Some means must be provided, therefore, to indicate the engine's power output. This is done by using an engine pressure ratio system, which is commonly known as EPR. The system consists of pitot type pressure heads located in the engine inlet, which are averaged together and a series of pitot type pressure heads located at the turbine exhaust which are averaged together. Both feed into a pressure ratio transmitter. On a high bypass engine the sensed pressure at the rear of the engine can be the by pass or cold flow or a combined input from both the hot and cold flows. The transmitter receives the pressure inputs from the inlet, and from the exhaust gas pressure probes. The probes are connected in to a common manifold, thus providing an average gas pressure. Both pressure tubes to the transmitter are provided with water drain traps that must be drained during maintenance checks. The formula used by the transmitter in determining the EPR signal is:- EPR = exhaust pressure inlet pressure Sometimes it can be expressed by using engine station configuration numbers, i.e. inlet PT2 or Exhaust PT7 (PT= pressure total), therefore EPR can be expressed as:PT7 PT2 As EPR is used as a thrust parameter, the flight crew must determine the maximum EPR for the barometric/temperature conditions. Take off EPR or maximum EPR can be determined by checking trim charts for engineers, or take off charts for flight crew. The EPR gauge in Fig. 14.8. that there is an EPR set knob. Once the EPR target figure has been calculated, then by turning the knob 'a reference target bug can be set at the take off EPR setting. This indicates to the crew the maximum amount of EPR required. Exceeding this figure could possibly overboost the engine. Modern aircraft use aircraft sensors to make this correction and will set the bug for the pilot if required.


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EPR Used in a Low Bypass Engine Figure14.8


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14.3.2 TORQUE INDICATION

Turboprop and turboshaft engines do not provide significant thrust through their jet pipes, so EPR would not be of any use in determining the thrust being produced by the engine. Engine torque is used to indicate the power that is developed by these engines, and the indicator is known as a torquemeter. The engine torque or turning moment is proportional to the horsepower and is transmitted through the propeller or rotor reduction gear. A torquemeter system is shown in fig. 14.9. In this system, the axial thrust produced by the helical gears is opposed by oil pressure acting on a number of pistons; the pressure required to resist the axial thrust is transmitted to the indicator.

Oil Pressure Type Torquemetering System. Figure 14.9. In addition to providing an indication of engine power, the torquemeter system may also be used to automatically operate the propeller feathering system if the torquemeter oil pressure falls due to a power failure. It is also used, on some installations, to assist in the automatic operation of the water injection system to restore or boost the take-off power at high ambient temperatures or at high altitude airports. Issue 2 – April 2003

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Operation The helical gear form used in the reduction gearbox develops an axial thrust in its three layshaft assemblies. This thrust is proportional to the power which is being transmitted through the reduction gearbox. The axial thrust is balanced by an opposing oil pressure, which is therefore proportional to engine power. This oil pressure is referred to as torquemeter pressure and is indicated on a flight deck instrument. Each of the layshafts operates against a piston that is supplied with oil pressure from a torquemeter pump. The torquemeter supply comes from the pressure side of the engine lubricating system. To balance any changes in axial thrust, or engine power changes, the oil pressure is regulated by a control valve that is incorporated in the lower piston assembly. The piston on the lower layshaft assembly is drilled centrally and operates over a stationary control valve. Flats on the control valve align with radial drillings in the piston. This is oil spill to the engine oil scavenge system as shown in Fig. 14.10. With the engine running at a stabilised power setting the lower piston will be in a sensitive position, allowing a constant spill of oil to engine scavenge. In this situation oil pressure is balancing the axial thrust. With an increase in engine power the layshaft pushes the piston further over the control valve. The oil spill is reduced, the oil pressure then increases giving an increased thrust indication on the flight deck instrument. With a decrease in engine power the oil pressure pushes the piston and the layshaft rearwards. The control valve now increases the oil spill, and the oil pressure decreases until it balances the axial thrust on the layshafts. If an engine fails the torquemeter pressure rapidly decreases below its normal operating range, this condition is referred to as a negative torque signal. The negative torque signal activates a low torque switch, which will in turn could activate the automatic feathering sequence.


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Oil Pressure Torquemetering Schematic. Figure 14.10. Issue 2 – April 2003

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14.3.3 PHASE COMPARISON TORQUEMETER

This method of torquemetering is often used in helicopters and modern turboprop engines. The shaft transmitting the load to the propeller or rotor has a second coaxial shaft splined to it, this shaft is not loaded. At the other end of this shaft there are two sets of pulse probes and phonic wheels. Normally the pulses will be in phase with one another, but as the drive shaft is loaded it will twist very slightly and the pulses will move out of phase with one another, the time difference being proportional to the torque.

A Phase Comparison Torquemeter System Figure 14.11.


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14.4 EXHAUST GAS TEMPERATURE Monitoring of the temperatures within the engine core is performed by the exhaust gas temperature system. The operating limits of the engine, and monitoring of the mechanical integrity of the turbines during operation, is vital for the continuing serviceability of the engine. Exhaust gas temperature, abbreviated to EGT, is only one of the terms relating to gas temperature. It can also be known as:

turbine gas temperature (TGT)



jet pipe temperature (JPT)



turbine inlet temperature (TIT)



turbine blade temperature (TBT)



Intermediate turbine temperature (ITT)

The EGT system consists of a series of thermocouples arranged radially in the exhaust- section of the engine. The exact location is decided by the engine manufacturer; other components within the system are:

a thermocouple junction box



a balance resistor box (junction box)



indicators on the flight deck.

A typical system lay-out is illustrated in Fig. 14.12 . 14.4.1 THERMOCOUPLES

The thermocouple itself consists of two dissimilar metals joined together within the probe body. Gas inlet holes are provided in the outer casing to allow hot gases to circulate around the sensing elements. The most common types of dissimilar sensing wires used are chromel and alumel. The probes may contain more than one thermocouple to sense the temperature at different lengths into the exhaust duct, or adjacent probes may be of different lengths. Some engines may have more than one EGT system. One for FADEC or for temperature limiting. The junction of the two wires (within the probe) is known as the hot or measuring junction; the indicator end is known as the cold or reference junction. The operation is fairly simple, as the thermocouple is a self-generating electrical system. Assuming that the reference end is kept at a constant temperature (flightdeck) and the hot end is subjected to high gas temperatures, then an electromotive force (emf), created by the dissimilar metals. The Seebeck effect causes the indicator to move in proportion to the difference in temperature between the two junctions. Issue 2 – April 2003

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A Thermocouple system with temperature compensation in the gauge and overall resistance compensation. Figure14.12. The thermocouples are connected electrically in parallel to provide an average gas temperature. The two wires (chromel and alumel) from each thermocouple terminate at the junction box. The chromel wires are connected together to form a parallel circuit, the alumel wire is common to all thermocouples. The junction box can also be used to check the thermocouple continuity during maintenance checks. From the junction box, the chromel and alumel wires are routed to the indicator on the flight deck. In some installations the cold junction is not in the gauge, but is a separate thermocouple located in the intake. The benefit of this system is that when a top temperature system is used to trim the fuel control unit, the majority of the components in the temperature system are located on the engine. It will also indicate the temperature difference across the engine.


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EGT System With Intake Thermocouples. Figure 14.13

Thermocouples are designed in two basic forms:

surface contact - used mainly on piston engines



immersion - Used in Gas Turbines

The immersion type thermocouple can be further divided into two categories:

stagnation type



rapid response type.

The main difference between the two examples shown in Fig. 14.14.is the position of the outlet holes in relation to the gas flow Inlet holes. The main reasons for these arrangements relate to the velocity of the exhaust gases. Issue 2 – April 2003

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The stagnation type is fitted to pure jet engines where the exhaust velocity is high, allowing the larger inlet hole to let the gas circulate around the couple, with the offset outlet hole reducing the outward velocity of the air. In this way the probe receives a good sampling of the gas temperature.

Types of Thermocouple Figure 14.14.

The rapid response type will be fitted mainly to turboprop engines where the gas flow is not as high as the jet turbine flow. In this arrangement the inlet and outlet holes are the same, creating no restriction, so a rapid response of EGT indication is achieved. Finally if we consider the EGT gauge (Fig. 14.15.) you will see that there are similarities to the rpm indicator.


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The indicator shown in Fig. 14.15. is a fairly modern type, although you may experience older instruments with a pointer only. Normally EGT is expressed in degrees centigrade. A red line limit indicates the maximum permissible temperature the engine is allowed to run at. And on some a red dot shows the maximum overswing allowed for a very short time. Finally, in addition to the maximum red line limits, most engines have an engine start EGT limit that is much less than the max. limit. this lower limit protects a cold engine from thermal shock (overtemping) during initial engine start.

TGT Gauge. Figure 14.15.

14.5 FUEL FLOW METERING Fuel flowmeters are fitted in aircraft to give an accurate indication of the rate at which fuel is being used and the total amount of fuel that has been used at any point during the flight. From the rate of fuel consumption the pilot is able to determine the performance of his engines, and from the indication of the total fuel consumed, can calculate the total flying hours that the aircraft can remain in the air. There are a number of different types of fuel flowmeters in use on various aircraft and it is beyond the scope of this publication to describe them all. Some of these flowmeters indicate only the total fuel consumed, but the majority give indications of both rate of flow and total fuel consumed.


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The volumetric flowmeter shown in figure 14.16. has a turbine rotor with a magnet inset into one of the vanes. When the rotor rotates it induces a pulse in the induction coil. The bore of the unit is calibrated to cause the rotor to rotate 32 times for every pound of fuel passing through it. The pulses are passed through a system of circuits similar to the speed probes mentioned earlier. This type of A Simple Volumetric Flowmeter flowmeter can indicate flow Figure 14.16. in gallons or litres. Although 16 it is calibrated in pounds per hour, this figure is only accurate at one S.G. or temperature. A similar system using a moving vane in a toroidal chamber is available, again only accurate at one S.G. To indicate mass flow accurately a flowmeter that compensates for changes of S.G. is required.

A Mass Flow Type Flowmeter system Figure 14.17.


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The mass flow type of flowmeter gives a reading of the mass flow rate in pounds or kilograms per hour rather than a volumetric reading in gallons per hour. The mass flow rate is a more useful indication for most types of aircraft. Refer to figure 14.17. for a mass flowmeter. The mass flowmeter consists of a motor-driven impeller, a turbine and a synchro system to transmit the data to a flightdeck gauge. In order to give accurate readings, the impeller must be driven at a constant speed. This is accomplished with an AC synchronous motor or a similar device. As the fuel flows through the impeller, it is given a spin or rotation by the spinning impeller. When the fuel leaves the impeller, it strikes the turbine, which is rotated against a restraining spring by the spin energy of the fuel. Because a denser fuel would impart more spin energy to the turbine the degree of rotation of the turbine is a measure of mass flow rate. The turbine is connected to the transmitter rotor of a synchro system which will cause the pointer on the flightdeck gauge to rotate to the proper position to indicate the correct mass flow rate. The sensor for this and other types of flowmeters is installed in the fuel system downstream of the fuel control device so that the flow rate represents the fuel consumption rate for that engine. There are other type of mass flow transmitters, that use swirl vanes to cause the rotation and have a different type of detection system, or vane type with complicated S.G. correction. The flowmeter gauge will have a flow indicator and usually a fuel used indication. The fuel used indicator is usually a digital read-out that is derived by integrating the fuel used with time. The gauge can be calibrated in pounds per hour of kilograms per hour.

Fuel Flowmeter Gauge Figure 14.18.


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14.6 OIL 14.6.1 THE OIL PRESSURE INDICATOR

The oil pressure indicator has a dial normally calibrated in pounds per square inch (psi). The indicator may have max. limit markers, but will always show the minimum pressure that the engine is allowed to run at. The reason that some engines have an upper limit is dependant upon the type of oil supply system. Some systems may be regulated, therefore needing an upper limit, or be based upon flow where an upper limit is not required.

Oil Pressure Gauge. Figure 14.19.

An example of an oil pressure indicating system is given in Fig.14.19.; the pressure indicator has no upper oil pressure limit, however, the low pressure limit is shown as 15 psi. There is also a precautionary band, normally yellow in colour, that is set just above the lower limit in the case in Fig. 14.19., an indication of between 15 and 25 psi in this yellow band during engine operation may require corrective action.

Any change in oil pressure introduced into the synchro transmitter causes an electrical signal to be transmitted through the interconnecting wiring to the synchro receiver. This signal causes the receiver rotor and the indicator pointer to move a distance proportional to the amount of pressure exerted by the oil. Most oil pressure transmitters are composed of two main parts, a bellowsor diaphragm mechanism for measuring pressure and a synchro assembly (Fig. 14.20.) The pressure of the oil causes linear displacement of the synchro rotor. The amount of displacement is proportional to the pressure, and varying voltages are set up in the synchro stator. These-voltages are transmitted to the synchro indicator. The vent tube to atmosphere prevents a build up of pressure within the transmitter that may interfere with the operation of the diaphragm at high altitudes.

Oil Pressure Gauging System Figure 14.20. Issue 2 – April 2003

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14.6.2 OIL PRESSURE WARNING LIGHT

Oil pressure is also monitored by an oil pressure switch (figure 14.21) that puts a light on when the oil pressure reaches a low level. The light is usually red and will be incorporated into the aircraft warning systems to alert the pilot. On later aircraft the pressure switch may have two pressure switched within it. A speed comparator will decide which switch to monitor. The idea being that a low oil pressure of say 20 psi is fine at low engine speed, however at higher engine speeds the engine could be sustaining damage due to insufficient oil pressure even though it is above 20 psi. The second pressure element would be activated when the engine speed was greater than say 80% and the oil pressure less than 50 psi.

Low Oil Pressure Warning. Figure 14.21.


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14.6.3 OIL TEMPERATURE.

Oil temperature can be detected by a temperature probe. The sensing element of the probe is a resistance wire. When heated the resistance of the wire will change. This can be measure by a wheatstone bridge system. However the wheatstone bridge power supply will also vary the gauge reading so making this method Wheatstone Bridge Temperature System inaccurate. Figure 14.22. It is more usual to use a ratiometer system to measure the resistance. In this instrument the measured resistance and the calibration resistance are in parallel, varying the current flow through two coils which are arranged to provide opposite torque to the pointer. This type of instrument can measure temperature up to 150°C, so is capable of monitoring an engine oil system.

Ratiometer Type System Figure 14.23.


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14.6.4 OIL QUANTITY

Oil quantity indicators are usually found in most aircraft these days. They usually consist of a float and probe. The float has a bush which supports it on the probe, a magnet within the bush sequentially operates reed switches within the probe. These switches change the resistance’s at A & B as the oil level changes, which will be read on a desyn type gauge in the flight deck. 14.7

Oil Quantity Probe (ALF 502) Figure 14.24.


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14.7 VIBRATION A turbo-jet engine has an extremely low vibration level and a change of vibration, due to an impending or partial failure, may pass without being noticed. Many engines are therefore fitted with vibration indicators that continually monitor the vibration level of the engine. The indicator is usually a milliammeter that receives signals through an amplifier from engine mounted transmitters fig.14.25. A vibration transmitter accelerometer is mounted on the engine casing and electrically connected to an amplifier and indicator. The vibration sensing element is usually an electromagnetic transducer that converts the rate of vibration into electrical signals and these cause the indicator pointer to move proportional to the vibration level. A warning lamp on the instrument panel is incorporated in the system to warn the pilot if an unacceptable level of vibration is approached, enabling the engine to be shut down and so reduce the risk of damage. The vibration level recorded on the gauge is the sum total of vibration felt at the pick-up. A more accurate method differentiates between the frequency ranges of each rotating assembly and so enables the source of vibration to be isolated. This is particularly important on multi-spool engines.(Figure 14.26. refers) A crystal-type vibration transmitter, giving a more reliable indication of vibration, has been developed for use on multi-spool engines. A system of filters in the electrical circuit to the gauge makes it possible to compare the vibration obtained against a known frequency range and so locate the vibration source. A multiple-selector switch enables the pilot to select a specific area to obtain a reading of the level of vibration. 14.8 WARNING LIGHTS Warning lights are used to indicate to the pilot if a failure has occurred. These will be red for something that requires immediate action or amber for less urgent items. Lights are also used to indicate when a function has operated. These light are usually white, blue or green. Warning lights may also be provided for L.P. fuel filter blocked, low fuel supply pressure, vibration low oil pressure and any other system the designer or the engineering authority require.


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Vibration Indicating System Figure 14.25.


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Vibration Signal Conditioner. Figure 14.26.


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15 THRUST AUGMENTATION 15.1 INTRODUCTION There are occasions when the maximum thrust from a basic gas turbine engine is inadequate and some method of increasing the available thrust is required without resorting to a larger engine with its concurrent penalties of increased frontal area, weight and fuel consumption. There are two recognised methods of augmenting this maximum thrust: a. De-mineralised Water or water/Methanol injection to restore, or even boost, the thrust from a gas turbine operating from hot and high altitude airfields. b. Reheat (or afterburning) to boost the thrust at various altitudes, especially at high speeds. This is normally for short periods only. 15.2 WATER INJECTION 15.2.1 EFFECTS ON ENGINE POWER

The power output from a gas turbine engine depends upon the weight (air density) of the airflow and the amount that it is accelerated as it flows through the engine. Therefore, it follows that any condition that reduces the air density will reduce also the engine power output. The two main natural causes of reduced air pressure are: 

Increased Altitude



Increased Temperature

When these two causes of reduced air density are combined at a high altitude/ tropical airfield, there is a possibility that engines may not produce sufficient power for a safe take-off and climb out. However, in these circumstances, the engine power can be restored and in some instances increased, by cooling the airflow to increase its density. To date, the addition of water or a water/methanol mixture has proved to be the cheapest practical means of restoring or increasing the power of an engine. Methanol has anti-freezing properties and it is also a fuel; therefore water/methanol increases the density of the airflow and provides the extra fuel necessary to match the increased weight of air. Adjustments to the engine fuel system are, therefore, unnecessary. The addition of water has two effects upon the performance of the engine: the cooling effect of water increases the density of the airflow to increase the thrust and, when the water is converted into steam, it provides a high volumetric expansion that increases the thrust even further. 15.2.2 METHODS OF APPLYING WATER/METHANOL

The following notes describe two methods of using water/methanol as a means of restoring lost engine power, or as a means of increasing the total useful power obtainable from a gas turbine engine. The water/methanol mixture can be 

Injecting as a spray into the compressor air intake.



Injecting direct into the combustion chamber.


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Spraying the mixture into the air intake is more effective for engines with centrifugal compressors than it is for axial compressors. With centrifugal compressors, an even distribution of the mixture is obtained whereas, with an axial flow compressor, even distribution is uncertain. (Turbo propeller engines use intake injection regardless of the type of engine in use). Water/methanol injection into the combustion chamber used to be carried out on older engines where the combustion chambers were relatively long and the methanol had time to separate and burn before entering the turbine. Later engines use water only and increase the fuel flow to gain the extra thrust. 15.2.3 COMPRESSOR INTAKE INJECTION (TURBO PROP)

When water or water/methanol mixes with the air at the compressor intakes, the temperature of the air is reduced and, as a result, the air density, mass airflow and thrust are increased. If water alone were to be injected, it would reduce the turbine inlet temperature and permit an increased fuel flow to be used. When methanol is added, the turbine inlet temperature is partially restored by burning the methanol in the combustion chamber; this restores the engine power without adjusting the fuel flow.


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Water/Methanol Intake Injection System Figure 15.1.

Operation When the system is switched ON, water/methanol mixture is pumped from the aircraft-mounted tank to a control unit which meters the flow of mixture fed to the air intakes ( figure 15.1.). The flow of water/methanol is controlled by a single metering valve and a servo piston that is powered by engine oil. The flow of the engine oil to the servo piston is controlled both by a shut-off cock and the position of a servo valve which, in turn, is moved by a control mechanism. This control mechanism balances propeller torque system oil pressure against atmospheric air pressure upon a capsule assembly within the control. The oil cock is interconnected with the throttle lever in such a manner that until the throttle is moved to the take-off position, the oil cock remains closed and the water/methanol system is inoperative. Moving the throttle lever to the “take-off” position opens the oil cock to motivate the water/methanol system.


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15.2.4 COMBUSTION CHAMBER INJECTION SYSTEM

Injecting a water or water/methanol mixture into the combustion chambers increases the mass flow through the turbine and the high volumetric expansion as the water becomes steam increases the thrust. The pressure and temperature drop across the turbine is reduced and this further increases the thrust. The reduction in turbine inlet temperature due to water injection enables the fuel system to provide an increased fuel flow to restore the maximum speed of the engine, thus providing further additional thrust without exceeding the safe turbine gas temperature limits(See figure 15.2.). When methanol is used with the water the turbine inlet temperature is partially restored without extra fuel from the fuel system. 15.2.4.1

Operation

Water flows from an aircraft-mounted tank to an air turbine driven water pump and is delivered to a water flow sensing unit (see figure 15.3.). From the water sensing unit the mixture is distributed to the burner feed arms where two jets at the base of each arm spray the mixture on to the upstream side of the swirl vanes to cool the air entering the combustion zone. The water pressure between the sensing unit and the discharge jets, is sensed by the fuel system control, which automatically resets the engine speed governor to give a higher maximum engine speed. The water system is brought into operation when the throttle lever is moved into the take-off position where it closes micro-switches to provide an air supply for the air turbine-powered water pump. The water flow sensing valve opens when a correct pressure difference exists between water pressure and compressor delivery air pressure. The valve in the water flow sensing unit also acts as a non-return valve to prevent air pressure feeding back from the water discharge jets and provides for the operation of an indicator to show when water/methanol is flowing.

Fuel Control Unit Speed Governor Reset Device to Increase Fuel Flow Figure 15.2.


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Water injection Into The Combustion Chamber Figure 15.3.


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15.3 RE-HEAT (AFTER BURNING) 15.3.1 PURPOSE

Re-heat is a system fitted to a gas turbine engine as a means of increasing the total thrust. As much as twice the thrust can be obtained using reheat. Unfortunately it is extravagant with fuel so is suitable for brief periods of use only; nevertheless, reheat allows flexibility in handling. The only civil aircraft to have reheat is Concorde. Principle The principle of re-heat is similar to that of the gas turbine engine itself – i.e. thrust is obtained as a reaction from accelerating a mass of air through the engine. Reheat obtains extra thrust from accelerating the exhaust gases in the jet pipe behind the turbine. The exhaust gases contain oxygen provided by the un-burnt cooling air. By adding fuel and burning it, the exhaust gases can be ‘re-heated’ to cause an increase in velocity with a substantial gain in thrust. A ring of fuel burners is mounted in the jet pipe and fed with fuel from the aircraft tanks, so that the exhaust acts like a ram jet. 15.3.2 REVISION OF THRUST

As the air flows through the engine it undergoes many changes in speed, direction and pressure. However, as we learnt in Chapter 1 of this book, the useful thrust depends upon the mass of air passing through the engine and upon the change in velocity between the air at the intake and that at the exit of the propelling nozzle. For a constant mass airflow, anything that increases the difference between the final velocity and the initial velocity will give an increase in thrust. Re-heat does just this; by burning fuel in the exhaust system behind the turbine we are creating a ram jet which increases the final velocity of the airflow; this in turn, increases the effective thrust from the engine. 15.3.3 RE-HEAT AND BY-PASS ENGINES

When re-heat is fitted to a by-pass engine, much greater thrust increase can be obtained. This is because the gas temperature before re-heat is much lower and hence the temperature ratio is much higher. Gains in the region of 70% increase in static thrust are readily obtained, with greater gains in thrust at high forward speeds. The limiting factor is the temperature that the jet pipe can withstand. 15.3.4 THE ADVANTAGE OF RE-HEAT

Re-heat provides the best means of substantially increasing the thrust of an engine for short periods. The advantages are those of improved take-off, rate of climb and air speed. Re-heat can be selected or cancelled at will by moving the throttle lever into or out of the re-heat position.


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15.3.5 THE DISADVANTAGES OF RE-HEAT

Because of the additional fittings, the diameter of the re-heat jet pipe is greater than that of a standard jet pipe for the same engine. Therefore, drag may be increased because the overall frontal area of the engine is increased. There is also a small weight penalty and the maximum continuous thrust is slightly reduced by the drag of the re-heat fittings inside the pipe. Re-heat is grossly extravagant with fuel. 15.3.6 PROPELLING NOZZLES

The design of the jet pipe and nozzle area has a considerable influence upon the overall useful thrust produced by a gas turbine engine. Generally the jet pipe and the propelling nozzle match the gas flow characteristics of the engine so that the final pressure and velocity of the gas produces the greatest amount of useful thrust. Thus the area of the propelling nozzle is as important it must be designed to match the airflow characteristics of the engine if it is to obtain the desired balance between pressure, temperature and thrust. A fixed area propelling nozzle, as fitted to non re-heat engines, is a compromise designed to provide an acceptable amount of thrust without being ideal for all engine speeds. The size of a fixed nozzle is chosen to provide its greatest efficiency at high cruising and maximum power but, a variable area nozzle would be more efficient.


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15.3.7 RE-HEAT NOZZLES

If re-heat was fitted to an engine with a standard sized fixed area propelling nozzle, the expansion of gases caused by the use of re-heat would increase the pressure in the jet pipe and reduce the pressure drop across the turbine (turbine expansion ratio). A reduced turbine expansion ratio will slow down the turbine and consequently lower the engine power. It would also increase the back pressure on the rear stage of the compressor which would cause the compressor to surge. To avoid a rise in pressure at the turbine outlet, the area of the propelling nozzle must be enlarged when re-heat is in use. Thus the propelling nozzle of a re-heat engine must be able to provide a nozzle area suitable for normal running without re-heat and a larger nozzle area when re-heat is used. Re-heat can usually be selected only after the throttle lever has passed through a normal 100% position. Therefore the smallest nozzle area must be efficient at normal maximum power and the large nozzle area must cater for the re-heat gas flow. If the amount of re-heat can be varied, then the re-heat nozzle must change to match the amount of re-heat selected. Variable Area Nozzles The variable propelling nozzle is suitable for use with controllable re-heat systems because it can provide a variable nozzle area to match the amount of re-heat selected. The circular continuity of the nozzle is maintained by a system of hinged flaps. The nozzle area is reduced by positive mechanical means but it is enlarged by the exhaust gas pressure acting upon the inside surface of the flaps. Description A ring of hinged master flaps is interleaved with a ring of hinged sealing flaps to provide a variable area propelling nozzle. Each flap is hinged at its forward edge so that the rear edge can move inwards to reduce the nozzle area, or outwards to increase the nozzle area. Actuation of the nozzle system can be hydraulic using oil or fuel as the fluid medium, or an air motor driving screw jacks. On selection of reheat the nozzle will move first to prevent back pressure on the engine, when it has moved the fuel will be supplied. With any increase in reheat the nozzle moves then the fuel follows. When reheat is reduced the opposite occurs first the fuel reduces then the nozzle closes. This ensures the nozzle area is too large rather than too small for any change in fuel flow.


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An Air Motor and Screwjack nozzle Figure 15.4. Actuation System

Reheat Jet Pipe with Hydraulic Actuation. Figure 15.5


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15.3.8 THE RE-HEAT JET PIPE 15.3.8.1

Description

The afterburning jet pipe is made from a heat resistant nickel alloy and requires more insulation than the normal jet pipe to prevent the heat of combustion being transferred to the aircraft structure. The jet pipe may be of a double skin construction with the outer skin carrying the flight loads and the inner skin the thermal stresses; a flow of cooling air is often induced between the inner and outer skins. Provision is also made to accommodate expansion and contraction, and to prevent gas leaks at the jet pipe joints. A circular heatshield of similar material to the jet pipe is often fitted to the inner wall of the jet pipe to improve cooling at the rear of the burner section. The heatshield comprises a number of bands, linked by cooling corrugations, to form a single skin. The rear of the heatshield is a series of overlapping 'tiles' riveted to the surrounding skin. The shield also prevents combustion instability from creating excessive noise and vibration, which in turn would cause rapid physical deterioration of the afterburner equipment.

Reheat Pipe Cutaway Figure 15.6.


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Re-heat Flame

Before looking at the re-heat burners and fuel supply systems, we must consider the problem of establishing and stabilising the re-heat flame. In the re-heat jet pipe where the flame must burn, the gas flow has a speed of the order of 500 mph (750 ft/sec to 1200 ft/sec). In effect, we are trying to burn fuel in a ‘wind tunnel’ and the problems are a magnification of those already described in chapter 11. Any attempt to establish a flame in the re-heat jet pipe will not succeed unless the airflow can be slowed locally and its pressure increased. Therefore the burner system must include some type of diffuser equipment. 15.3.8.3

The Burner Assembly

The construction of the re-heat burner assembly varies from one manufacturer to another. However, the burner assembly shown in figure 15.7. is typical of those now in use. This assembly consists of three concentric fuel manifolds, two concentric ‘V’ section flame stabilising gutters (vapour gutters) and a number of support struts; it is built upon a tubular centre piece. There are three long struts interspaced with three short struts and welded to the centre tube with 60 spacing. These struts locate and secure the burner assembly into the re-heat pipe. A modern trend is to use vaporisers set into the vapour gutters for the main fuel flow.

Reheat Burner Figure 15.7.


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Fuel Flow

A re-heat fuel pump receives fuel from the engine fuel supply Its operation and flow rate are controlled by a reheat control unit. The fuel is fed to the reheat burner by fuel pipes which run inside the burner support struts. The fuel is divided into main fuel flow and vapour gutter/ignition flow. The ignition fuel flow is used with ignition plugs and catalytic ignition systems. Vapour gutter flow provides a flow into the gutters which provides a stable, slower airflow to allow the flame to stabilise behind the gutters. Interconnectors allow the flame to spread between the vapour gutters. The main fuel flow goes to the spray nozzles that are upstream of the vapour gutters, and this fuel is atomised and vaporised before being ignited by the vapour gutter flame. 15.3.8.5

Re-heat Ignition

The atomised fuel spray is fed into the re-heat jet pipe and ignited by one of three methods:

Spark Ignition



Hot Streak Ignition



Catalytic Ignition

a. Spark Ignition. Spark ignition for re-heat fuel is similar to normal engine ignition. Light-up is obtained by using a pilot fuel burner and an igniter plug. The igniter plug is fitted downstream of the pilot burner in a conical fitting that is a part of the re-heat system. The core provides airflow conditions suitable for light-up and when fuel is sprayed from the pilot burner, it is carried on to the igniter plug and ignition takes place. This method has been superseded by the other methods.

Spark ignition Figure 15.8.


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Hot Streak Ignition With a Relay Figure 15.9. Supply

b. Hot Streak Ignition. The hot streak ignition system is more often called ‘hot shot’ ignition. It consists of one or two fuel injectors; one sprays fuel into the engine combustion system and the other if fitted sprays fuel aft of the turbine as a relay system to keep the flame alight for a longer distance. Spraying additional fuel into the main combustion area causes an elongated flame and a ‘hot streak’ flame reaches and ignites the re-heat fuel. The turbine blades are not damaged because the hot streak flame is of short duration. This method provides a very quick light up, however if it fails to light then reheat has to be reselected. c. Catalytic Ignition. Catalytic ignition is achieved by use of a platinum/rhodium element. Atomised fuel is sprayed over the element and a chemical reaction causes spontaneous ignition.

Catalytic Ignition Figure 15.10.


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INTENTIONALLY BLANK


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16 TURBO-PROP ENGINES 16.1 INTRODUCTION The turbo-prop engine consists of a gas turbine engine driving a propeller. In the turbo-jet engine the turbine extracts only sufficient energy from the gas flow to drive the compressor and engine accessories, leaving the remaining energy to provide the high velocity propulsive jet. By comparison, the turbine stages of the turbo-prop engine absorb the majority of the gas energy because of the additional power required to drive the propeller, leaving only a small residual jet thrust at the propelling nozzle. Turbo-shaft engines work on identical principles, except that all the useful gas energy is absorbed by the turbine to produce rotary shaft power and the residual thrust is negligible; such engines find particular applications in helicopters and hovercraft. The lack of a significant propulsive jet means that these engines can be mounted in any position in the airframe and this flexibility is increased by the very compact design and layout of a modern turbo-shaft engine. Because the propeller wastes less kinetic energy in its slipstream than a turbo-jet in its exhaust, the turbo-prop is the most efficient method of using the gas turbine cycle at low and medium altitudes and at speeds up to approximately 350 knots. At higher speeds and altitudes, the efficiency of the propeller deteriorates rapidly because of the development of shock waves on the blade tips. 16.2 TYPES OF TURBO-PROP ENGINES Current turbo-prop engines can be categorised according to the method used to achieve propeller drive; these categories are: a. Coupled Power Turbine. b. Free Power Turbine. c. Compounded Engine. 16.2.1 COUPLED POWER TURBINE

The coupled power turbine engine is the simplest adaptation from the turbo-jet engine. In this configuration, the gas flow is fully expanded across a turbine which drives the compressor, the surplus power developed being transmitted to the propeller by a common drive shaft via suitable reduction gearing. This arrangement is shown diagrammatically in figure 16.1.


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Coupled Power Turbine Figure 16.1. 16.2.2 FREE POWER TURBINE

In this arrangement, a gas turbine acts simply as a gas generator to supply highenergy gases to an independent free power turbine. The gases are expanded across the free turbine, which is connected to the propeller drive shaft via reduction gearing. The layout of a free power turbine engine is shown in the figure 16.2. The free turbine arrangement is very flexible; it is easy to start due to the absence of propeller drag and the propeller and gas producer shafts can assume their optimum speeds independently.

Free Power Turbine Figure 16.2. Issue 2 – April 2003

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16.2.3 COMPOUNDED ENGINE

The compounded engine arrangement features a two-spool engine, with the propeller drive directly connected to the low-pressure spool.

A Compounded Turboprop Engine. Figure 16.3.

16.3 REDUCTION GEARING The power turbine shaft of a turbo-prop engine normally rotates at around 8,000 to 10,000 rpm, although rpm of over 40,000 are found in some engines of small diameter. However, the rotational speed of the propeller is dictated by the limiting tip velocity. A large reduction of shaft speed must be provided in order to match the power turbine to the propeller. The reduction gearing must provide a propeller shaft speed which can be utilised effectively by the propeller; gearing ratios of between 6 and 20:1 are typical. In the direct coupled power turbine and compounded engines, the shaft bearing the compressor and turbine assemblies drives the propeller directly through a reduction gearbox. In the free turbine arrangement reduction gearing on the turbine shaft is still necessary; this is because the turbine operates at high speed for maximum efficiency. The reduction gearing accounts for a large proportion (up to 25%) of the total weight of a turbo-prop engine and also increases its complexity; power losses of the order of 3 to 4% are incurred in the gearing (eg. on a turbo-prop producing 6,000 eshp, some 200 shp is lost through the gearing).


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ENGINES

16.3.1 SIMPLE SPUR ‘EPICYCLIC’

A gear train consisting of a sun (driving) gear meshing with and driving three or more equi-spaced gears known as ‘Planet Pinions’. These pinions are mounted on a carrier and rotate independently on their own axles. Surrounding the gear train is an internally toothed ‘Annulus Gear’ in mesh with the Planet Pinions.

An Epicyclic Gear TrainFigure 16.4.

If the annulus is fixed, rotation of the sun wheel causes the planet pinions to rotate about their axes within the annulus gear, this causes the planet carrier to rotate in the same direction as sun wheel but at a lower speed. With the propeller shaft secured to the planet pinion carrier, a speed reduction is obtained with the turbine shaft (input shaft) and propeller shaft (output shaft) in the same axis and rotating in the same direction. (Fig.16.5.)


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Epicyclic Gear train with Fixed Annulus Ring Rear. Figure 16.5. If the annulus is free, rotation of the sun wheel causes the planet pinions to rotate about their axles within the annulus gear. With the planet pinion carrier fixed and the propeller shaft attached to the annulus gear, rotation of the planet pinions causes the annulus gear and propeller to rotate in the opposite direction to the sun wheel and at a reduced speed. (Fig.16.6.)

Epicyclic Gear Train with Fixed Planet gear Carrier. Figure 16.6.


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16.3.2 COMPOUND SPUR EPICYCLIC

Compound epicyclic reduction gears enable a greater reduction in speed to be obtained without resorting to larger components. They may be of either the fixed or free annulus type.

Compound Spur Gear TrainsFigure 16.7.


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16.3.3 GEAR TRAIN/EPICYCLIC

Some turbo-props will use a gear train or a combination of gear train and epicyclic.

Garrett 331 Cut away Showing the Combined Epicyclic Gear Train. Figure 16.7.

16.4 TURBO-PROP PERFORMANCE The turbo-prop has a higher propulsive efficiency than the turbo-jet up to speeds of approximately 575 mph and higher than a turbo-fan engine up to approximately 450 mph. Compared with a piston engine of equivalent power, the turbo-prop has a higher power/weight ratio and a greater fatigue life because of the reduced vibration level from the gas turbine rotating assemblies. 16.5 TURBO-PROP ENGINE CONTROL The gas generator element of the turbo-prop engine operates at high rpm for maximum efficiency; any reduction in rpm reduces the pressure ratio across the compressor and therefore adversely affects the sfc. In practice, most turbo-prop engines have gas generators which run at or near 100% rpm and three main methods are used to control the rpm and power absorption of the propeller throughout the normal flight ranges. These are: 

Integrated control of both rpm and fuel flow.



Direct control of gas generator fuel flow.



Direct control of propeller blade angle.


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16.5.1 INTEGRATED CONTROL OF RPM AND FUEL FLOW

The integrated control system is suitable for coupled power turbine and compounded turbo-prop engines. In this system the propeller rpm is selected by a power lever that simultaneously adjusts the fuel flow to ensure that the correct flow is maintained for any selected rpm. Up to maximum rpm, the turbo-prop runs at the selected rpm, increases in rpm demanded by the power lever being automatically accompanied by corresponding increases in fuel flow, blade angle and hence power. At maximum rpm, further increases in power are achieved by increasing the fuel flow; the propeller constant speed unit (CSU) automatically increases the blade angle to absorb the extra power and thus maintain constant speed. 16.5.2 DIRECT CONTROL OF FUEL FLOW

The direct control of fuel flow is suitable for use in a free power turbine engine. In this system, the gas generator is controlled in the same manner as a turbo-jet and the power available to the free turbine assembly is governed by the fuel flow. Through reduction gearing, the free turbine turns the propeller that is maintained at constant rpm by the CSU. 16.5.3 DIRECT CONTROL OF BLADE ANGLE (BETA CONTROL)

This control system can be used for any turbo-prop engine. In this system, the cockpit power lever simply selects a blade angle (B) and various automatic systems are used to maintain the propeller rpm by adjusting the fuel flow (e.g. by a governor in the fuel control system). The Astazou engine in the Jetstream is typical of the direct-coupled engine in which this control system is used. As the propeller blade angle is changed, the propeller speed governor adjusts the fuel flow to maintain constant propeller rpm (and thus constant engine rpm). The direct control of blade angle in a free turbine system is found most commonly in helicopter turbo-shaft engines. Here, the blade angle is selected by the collective lever and the output of the gas generator is automatically adjusted to maintain the rotor rpm within fine limits.


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ENGINES

16.6 ENGINE AND PROPELLER CONTROLS Turbo-props are normally controlled by two levers per power unit, a power lever and a condition lever. The diagrams show a typical installation for a twin turbo-prop. The example shown is a DASH-8.

Power and Conditioning Levers. Figure 16.8.

16.7 CONTROL OUTSIDE NORMAL FLIGHT RANGE Outside the normal flight range and particularly in the reverse thrust range, the engine/propeller combination is normally controlled by the beta system, i.e. by direct control of propeller blade angle. The transition point between the control systems is usually indicated by a stop or detent in the throttle lever quadrant. 16.8 PROPELLER CONTROL The main propeller controls found on the majority of turbo-prop engines are as follows: a. Constant speed unit. b. Manual and automatic feathering controls. c. Fixed and removable stops. d. Synchronisation and synchrophasing units. e. Reverse thrust control. Issue 2 – April 2003

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16.8.1 CONSTANT SPEED UNIT

In the normal flight range, the main control of the propeller is exercised by the propeller control unit (PCU). 16.8.2 MANUAL AND AUTOMATIC FEATHERING CONTROLS

All turbo-prop aircraft are fitted with some form of manual feathering control. In some cases this control is integral with the HP cock for the associated engine; in others the feathering control is operated through the fire protection system which also closes the HP cock. Automatic feathering control is fitted to many turbo-prop engines to avoid excessive drag following an engine failure. The automatic system receives signals from the engine torquemeters and reacts to unscheduled loss of torque by feathering the appropriate propeller. On twin-engine turbo-prop aircraft, the operation of the autofeather system on one engine automatically inhibits the same operation on the other engine, while still allowing the latter to be feathered manually. 16.8.2.1

Power Lever

The power lever operates in a quadrant slot labelled “POWER” with positions (from rear to front) labelled “MAX REV”, “DISC”, “FLT IDLE” and “MAX”. The power lever is connected by cables, pushrods and bellcranks to the control system and PCU of the associated powerplant. The power lever quadrant slot has a lockout gate at the FLT IDLE position, which is controlled by a finger latch below the power lever knob. Raising the latch permits aft movement into the ground range. The power lever controls power in the forward thrust range and blade angle in the flight Beta and ground Beta ranges. The flight Beta range extends from a blade angle of 26° to 19 (minimum in-flight blade angle). The power lever controls blade angle from aft of FLT IDLE to MAX REV. The spring-loaded, detented DISC position produces at 0 blade angle or flat discing; further aft movement increases blade angle in a negative direction until at MAX REV the blade angle is –11.5°. Both of these positions will assist in slowing the aircraft during landing. While operating in the Beta range, the HP fuel control regulates engine power, providing Np underspeed governing between FLT IDLE and DISC and both engine power and blade angle control in the reverse thrust range. When the flight control gust lock lever, labelled “CONT LOCK” is at the on position, the power lever cannot be moved to the MAX position. This lever will also lock the aircraft flight controls.


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Engine and Propeller Controls (Dash 8) Figure 16.9.


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Condition Lever (RPM Control)

The condition lever is connected to the PCU and HP fuel control by cables, pushrods and bellcranks and operates in a quadrant slot labelled “PROP” on the centre console. The condition lever positions are labelled (rear to front) “FUEL OFF”, “START & FEATHER”, “MIN” and “MAX”. The range between START & FEATHER and MIN is labelled “UN-FEATHER”. Inadvertent selections below MIN and START & FEATHER are prevented by detents. The lever must be pulled out for aft movement past these positions. Moving the condition lever from MIN to START & FEATHER feathers the propeller through the PCU and signals the HP fuel system to establish a fuel flow to sustain ground idle rpm. Moving the lever forward of START & FEATHER unfeathers the propeller when the engine is running. When the condition lever is moved from START & FEATHER to FUEL OFF, it mechanically closes the fuel shut-off valve on the HP fuel system and shuts down the engine. The condition lever range between MIN and MAX sets propeller rpm for in-flight constant speed operation. 16.8.2.3

Constant Speed Range

The constant speed range is defined as propeller operation from a fully fine setting (condition lever at MAX RPM) to an increased blade angle pre-selected by a condition lever angle (CLA) setting of a speed-sensitive, flyweight governor in the PCU. The governor operates to obtain and maintain constant speed settings between 900 and 1,200 propeller rpm (Np). Ground range lights indicate at 16.5 and the discing is between 1.5 and 3.0. 16.8.2.4

Beta Range

The term “Beta Range” is used to define propeller operation from a maximum Beta setting (propeller blade angle 26) to a full reverse setting (propeller blade angle – 11.5). The Beta range is divided operationally into two ranges by a gate on the associated power lever which controls blade angle from 16 to 19 above the gate and below the gate to full reverse. Propeller blade angle at full feather is 86 + 5.


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Power Lever and Propeller Ranges. Figure 16.10. Issue 2 – April 2003

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Power lever Quadrant and Associated Typical Blade Angles. Figure 16.11.


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16.8.3 FIXED AND REMOVABLE STOPS

A number of stops or latches can be incorporated in the propeller control system; their purpose is to confine the angular movement of the blades within limits appropriate to the phase of flight or ground handling. The most common stops are described below and typical values are given for the corresponding blade angles (see the figure). a. Feather and Reverse Braking Stops. These two fixed stops define the full range within which the propeller angle may be varied (+85 to -15). b. Ground Fine Pitch Stop. This is a removable stop (-1) which is provided for starting the engine and maintaining minimum constant rpm; the stop also prevents the propeller from entering the reverse pitch range. c. Flight Fine Pitch Stop. This is a removable stop (+14) which prevents the blade angle from fining off below its preset value. Its purpose is to prevent propeller overspeeding after a CSU failure. It also limits the amount of windmilling drag on the final approach. The stop is usually engaged automatically as the pitch is increased above its setting; removal of the stop is, however, usually by switch selection. d. Flight Cruise Pitch Stop. This is a removable stop (+27) which is fitted to prevent excessive drag or overspeeding in the event of a PCU failure. The stop engages automatically as the pitch is increased above its setting and is also withdrawn automatically as the pitch is decreased towards flight idle provided that two or more of the propellers fine off at the same time. Variations on this type of stop include automatic drag limiters (ADL) and a Beta follow-up system. In the first of these, the stop is in the form of a variable pitch datum which is sensitive to torque pressure. If the propeller torque falls below the datum value, the pitch of the propeller is automatically increased. The pitch value at which the ADL is set is varied by the position of the power lever. Thus, as the power is reduced, the ADL torque datum value is also reduced so that the necessary approach and landing drag may be attained, while simultaneously limiting the drag to a safe maximum value. The Beta follow-up stop uses the Beta control (ie. direct selection of blade angle for ground handling) to select a blade angle just below the value controlled by the PCU. In the event of a PCU failure, the propeller can only fine off by a few degrees before it is prevented from further movement in that direction by the Beta follow-up stop. In the flight range, the position of this stop always remains below the minimum normal blade angle and so does not interfere with the PCU governing. e. Coarse Pitch Stop. This stop (+50) limits the maximum coarse pitch obtainable in the normal flight range. A feathering selection normally over-rides this stop.


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16.9 OVERSPEED SAFETY DEVICES A propeller that overspeeds, even by the small amount of, say 5 or 10%, increases the centrifugal forces on the hub by a huge amount. This could cause the blades to separate from the hub with catastrophic results to the aircraft. A gas turbine engine has its own fuel control system, which maintains the engine within its operating speed range. With a turboprop engine it is normally the propeller which acts as a governor by increasing or decreasing its pitch angle to add or remove the loading on the rotating parts of the engine. If a turboprop overspeeds, it is usually due to the fact that the propeller controls have allowed the pitch angle of the propeller to decrease, so that the reduction of load on the engine has caused it to overspeed. This reduction of pitch is as a result of aerodynamic and centrifugal forces acting on the rotating propeller. If the reduction of the propeller pitch has been caused by failure of the propeller control unit, there may be a back-up method, built in to the control system, to drive the propeller back to a coarser angle, thereby slowing it down to a safe value. These back-up systems usually involve the use of centrifugal governors that sense the overspeed. If the propeller control system is damaged or it cannot drive the propeller to a safe, coarser, blade angle, the fuel control of the engine may reduce the flow of fuel to the engine, effectively acting as if the pilot had retarded the throttle. This should bring the hub loading within a safe value. As an example, the system shown is that fitted to the Pratt & Whitney 124 engines on the ATR72 aircraft, which has a combined hydraulic/pneumatic overspeed protection. If the propeller overspeeds above 102.5% NP, (NP = propeller speed), The flyweights move outwards, opening the pilot valve and allowing metered oil pressure to drive the propeller towards coarse. In the event that the above system fails to operate, (propeller continues to accelerate), the air bleed orifice opens at a slightly higher NP. This bleed biases the fuel control system, (H.M.U.) to decrease the fuel flow, reducing the engine speed.


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Propeller Overspeed Governor (ATR) Figure 16.12.


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Whilst the system previously described is rather complex, the engine of a modern, 'free power turbine' design has to have sophisticated protective measures fitted. By comparison, the overspeed protection installed on the Rolls Royce Dart, a 'direct coupled' drive engine designed in the 1940s, is a relatively simple system. The pump case pressure is fed with fuel from radial tappings in the rotating pump assembly. If the engine overspeeds, the fuel is 'centrifuged' into the pump case at a higher pressure. This pressure is fed to a diaphragm in the overspeed governor, which spills the servo pressure and reduces the fuel supply to the engine. This limits the engine, which normally has a governed maximum of 15,000 R.P.M., to an overspeed maximum of 16,400 R.P.M. The illustration below shows the basic system showing how spilling the servo pressure reduces the pump output. Apart from the protection mechanisms already mentioned, which have to react extremely fast to prevent accidents, there are a number of flight deck indications which may be in place of, or in addition to the automatic systems. The simplest is the 'red line' on the tachometer, (revolution counter), or power, (percentage), instrument, which must not be exceeded at any time. If the aircraft has an electronic flight warning system, (F.W.S.) however, then warning lights, captions and audio warnings may be used to get the attention of the flight crew.


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Older Style Governor Built into Fuel Pump. Figure 16.13. Issue 2 – April 2003

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ENGINES

17 TURBOSHAFT ENGINES 17.1 INTRODUCTION. Gas turbine engines that deliver power through a shaft to operate something other than a propeller are referred to as turboshaft engines. In most cases the output shaft (power takeoff), is driven by its own power turbine (free turbine), which extracts the majority of the total power output from the engines gas generator. Turboshaft engines with a reduction gear are used to power boats, ships, hovercraft, trains and cars. They are also used to pump natural gas across country and to drive various kinds of industrial equipment such as air compressors or large electric generators (fig 17.1.)

An Industrial Turboshaft Engine. Figure 17.1.


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In aviation turboshaft engines are used to power many of the modern helicopters in service. They are similar in design to turboprop engines and in some instances will use the same gas generator section design. The turboshaft power takeoff may be coupled to, and driven directly by the turbine that drives the compressor, but is more likely to be driven by a turbine of its own. Engines using a separate turbine for power takeoff are called free power turbine engines, and it is this type of engine that is most commonly used in today’s modern fixed wing and rotary wing aircraft. Atypical example of a turboprop/turboshaft engine is the Pratt and Whitney PT 6. (figure 17.2.)

The Pratt and Whitney (Canada) PT6 turboprop engine is a popular free turbine engine that can be adapted to both turboprop and turboshaft applications. Figure 17.2. A free power turbine engine consists of two main units; the gas generator and the free power turbine. In the example shown in Figure 17.2. air enters the engine and is compressed, then heated in the combustion chamber . The resulting expansion forces the gas at high velocity through the gas generator turbine that drives the compressor. The remaining gas energy is then used to drive the power turbine, which in turn drives the power output shaft. The free power turbine is mechanically independent of the of the gas generator and operates at virtually a constant speed. The power developed by the turbine is varied to meet changing loads imposed on the rotor system, by increasing or decreasing the fuel supplied to the gas generator, thus altering the gas generator speed and the supply of gas energy to the power turbine. As mentioned previously, the turboshaft engine is used to power many of today’s modern helicopters, and to this end we will concentrate on the application of the turboshaft engine in the field of aviation. Issue 2 – April 2003

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The turboshaft engine and the helicopter are ideal companions. The engine is required to respond to frequent and sudden changes in power demands to keep the helicopter rotor revolving at a virtually constant speed (250-300 RPM being typical). The power required to drive the rotor is determined by the pitch angle of the main rotor blades, this angle is being controlled by the pilot using the collective pitch lever. The pilot changes the flight path of the aircraft by using the cyclic pitch control lever, by tilting the rotor head. Control of the tail rotor to compensate for the torque produced by the main rotor is via foot pedals similar to rudder pedals (fig 17.3.). Whenever a control is activated, the resultant force is sensed by the rotor gearbox and in turn sensed by the power output shaft of the engine which means that the engine power must be adjusted to suit.

Flight Controls of a Typical Single Main Rotor Helicopter. Figure 17.3.


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The power output of a free power turbine engine can be changed rapidly because its output speed is independent of the power produced, the latter being dependant on the gas generator speed. The low inertia of the gas generator rotor allows its speed to be changed very quickly, by adjusting the flow of fuel available for combustion. This is achieved in the fuel control system invariably by a computer (electronic or mechanical) controlling the throttling valve. The pilot selects the rotor speed and the fuel control system automatically maintains that speed, within the limits set by the governing characteristics of the system and the operating limitations of the engine. As the fuel control system is automatic, the pilot is relieved of the necessity to constantly manipulate the throttle control. The control parameters being monitored and used for a typical turboshaft engine would include: Parameter Gas generator speed (N2) Free power turbine speed (N1) Power turbine inlet temperature (PTIT) Main rotor speed (Nr) Throttle valve position Torque

Destination Computer and cockpit gauge Computer and cockpit gauge Computer and cockpit gauge Cockpit gauge Computer Cockpit gauge and computer (torque matching engines)

17.2 FUEL CONTROL SYSTEM The computer controls the fuel flow to the engine to maintain a constant rotor RPM. During normal operation the optimum engine/rotor speed is selected by a speed selector lever, and the varying power demands are met thereafter by the automatic fuel computer. The computer varies the rates of fuel flow to the engines to suit the changing power demands occasioned by alterations of rotor blade pitch. The position of the throttle valve is set by an electric actuator controlled by the computer. The speed select lever in the cockpit is directly connected to the computer, and by operating this lever the pilot can select a power turbine speed that is maintained by the computer within built in control laws. In addition to speed selector lever positions , the computer receives signals of power turbine speed N1, gas generator speed N2, power turbine inlet temperature (PTIT), collective pitch angular movement via an anticipator, and throttle position. In the computer the signal representing actual power turbine speed is compared with the sped selector lever position , and any difference causes a signal to be transmitted from the computer to the throttle actuator, which adjusts the throttle opening accordingly. I however this were to cause the PTIT to exceed a predetermined value or to increase at too rapid a rate, the computer signal is modified so that the throttle is held or closed until the PTIT is reduced to a safe level.


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The function of the anticipator is to provide signals proportional to change of collective pitch angle.

Computer Signalling. Figure 17.4.


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17.3 ARRANGEMENTS Because of the need for turboshaft engines to be installed in a variety of aircraft, coupled with the requirement to fit two or more engines, giving more power and adding safety. The turboshaft engine has to be able to output its drive from a variety of different locations. Typical examples of this ability can be seen in Figure 17.5. to 17.9. Figure 17.5. shows the different ways in which the Rolls Royce Gem engine can be configured to suit different aircraft designs.

Different Ways Power can be Taken From the Rolls Royce Gem Engine. Figure 17.5.


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Turboshaft engines can be located forward or behind the main transmission gearbox. The Westland Lynx has two Rolls Royce Gem engines mounted aft of the gearbox driving through couplings at the front of the engines fig 17.6. It can be seen from the illustration how the engine/gearbox unit is quite compact.

The Rolls Royce Gem Installation in the Westland Lynx Helicopter. Figure 17.6.

Another twin engined installation is that which can be found fitted to numerous Sikorsky and Westland helicopters. these are fitted ahead of the main gearbox, so that the output shaft and coupling projects from the rear of each engine. the location of all the previously mentioned layouts permits very easy maintenance and engine changes due to the unobstructed access to the engines. Figure 17.7 shows the S61N model which has two 1400 S.H.P. turboshaft engines.

The Rolls Royce Gnome Engine Installation in a Westland S61N. Figure 17.7.


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Finally there are a few other installations on helicopters, using turboshaft engines, that show the flexibility in the way these engines can be mounted to suit the designers needs. The little Hughes 500 series (fig 17.8.) has a small 400+ S.H.P. engine, installed at an angle, driving upwards at 45° to the main gearbox.

The Engine Installation in a Hughes 500. Figure 17.8. The large E.H. 101 helicopter (fig 17.9.), however has not only three engines, each of 2,000 S.H.P., installed above the decking and all feeding into the main gearbox, but there is an Auxiliary Power unit installed alongside the No.2 engine as well.


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The EH101 Engine Layout. Figure 17.9.


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17.4 DRIVE SYSTEMS Because gas turbine engines rotate at extremely high speeds, and the main rotor of a helicopter needs to rotate at a fairly low, constant speed the output drive of a turboshaft engine must incorporate some form of reduction gearing. Some engines have their reduction gearing installed within the engine so that their output shaft is at a usable speed, which can be further reduced to a rotor speed by the main rotor gearbox. Figure 17.10. is of the reduction gearbox fitted to the front of a Rolls Royce Gem turboshaft engine. The gearbox takes the 27,000 RPM output of the power turbine shaft, and through the two stage epicyclic gear train, reduce it to approximately 6000 RPM, a speed reduction of some 4.5:1. At this speed it can be directly coupled to the main rotor gearbox, which will reduce it further to approximately 250-300 RPM. This reduction mechanism allows the engine to be used not only in helicopters but also in a number of different situations such as powering marine craft, power generating stations and pumping stations etc. This use of the turboshaft engine is very common and even engines as large as the Rolls Royce RB 211 series are used for such purposes. Other types of turboshaft engines will, because their power turbine rotational speed is not so high provide a direct power output to a separate reduction gearbox, in the case of a helicopter, the main rotor gearbox. A typical example of this is the power output shaft of the Rolls Royce Gnome turboshaft engine fitted to the Westland S61N helicopter (fig 17.11.)


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Rolls Royce Gem Engine Reduction Gearbox. Figure 17.10.


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Rolls Royce Gnome Power Turbine and Drive. Figure 17.11. Issue 2 – April 2003

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17.5 COUPLINGS Because of the requirement to make maintenance tasks such as engine removal/refit, gearbox removal/refit easier, it is necessary to have a means of coupling the turboshafts output shaft to the helicopter main rotor gearbox input shaft together. This coupling must possess qualities which will allow movement of both the engine and the rotor gearbox independently of each other i.e. it must be flexible. It must also be finely balanced to reduce vibration. One of the most common couplings in use is the ‘Thomas Coupling’, sometimes referred to as the engine ‘high speed drive shaft’ (fig 17.12.). The engine is joined to the main rotor gearbox by this high speed drive shaft. The shaft is belled at either end , one end being attached to the power take off shaft by means of Thomas flexible steel coupling. Each coupling consists of a number of steel discs, indexed by flats to ensure correct alignment when assembled. Two different numbered discs are used, each disc having a grain running either parallel to the flat or perpendicular to the flat. The discs are assembled alternately with the grains at 90° to each other. The bolts, nuts and washers securing the shaft to the engine are part of the fine balancing of the assembly and must always be replaced in the same position.

Thomas Coupling.

Figure 17.12. Issue 2 – April 2003

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Yet another method of coupling the engines power output to the main gearbox is shown in Figure 17.13. The engine front mounting is bolted with the reduction gearbox to the hub of the air-intake case; it supports the engine in the aircraft and serves as a torque reaction point. The mounting, which is of the gimbal type, is bolted to a gimbal ring, which is bolted to a similar mounting on the aircraft main gearbox, thus forming a gimbal coupling. The engine output drive is transmitted to the aircraft main gearbox by a flanged coupling, which is secured via a flexible laminated disc coupling (Thomas Coupling) to a drive assembly. The drive assembly consists of an engine coupling and an aircraft main gearbox coupling bolted together, with a flexible laminated disc coupling (Thomas Coupling) at each end.


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The Thomas Coupling and Gimbal Mount of a Gem Engine. Figure 17.13.


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Finally as an example of the end product of a typical, turboshaft engines power output Figure 17.14. shows the main rotor gearbox of a Westland S-61N helicopter. The two engines are Rolls Royce Gnome 1400 series turboshaft engines, each producing approximately 1400 S.H.P. Figure 17.15. shows the gearbox together with its monitoring devices and transmission. The free-wheel system enables disconnection of one or both the engines in the event of failure.

S-61N Rotor Gearbox. Figure 17.14.


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Seaking/S-61 Transmission System. Figure 17.15. Issue 2 – April 2003

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18 AUXILLIARY POWER UNITS 18.1 INTRODUCTION The auxiliary power unit or APU as it is commonly known, is a small gas turbine engine as shown in figure 18.1., fitted to aircraft and can provide:

Electric power from shaft driven generators.



Pneumatic duct pressure for air conditioning and engine starting purposes.



Hydraulic Pressure (Some aircraft).

An APU Figure 18.1.

An Auxiliary Power Unit (APU) is an automatic engine, which normally runs at a governed speed of 100%. Some APUs have an idle facility that allows the engine to run at 85% when no loads are applied. As it is an automatic engine the fuel system must control the engine throughout the start and running phases of operation. The engine will be shut down if a critical control function is lost or a serious malfunction such as low oil pressure occurs. APU’s are mainly used on the ground when their main engines are not running and ground carts (electrical and pneumatic) are not available. On most modern aircraft the APU will also be used in the air to provide air-conditioning during take off and landing phases, or to back up the main engines in case of a generator or air system failure. Issue 2 – April 2003

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Operating Altitude for an APU. Figure 18.2.


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Although the APU is usually rated to run at the max cruise altitude of the aircraft it is fitted to, its ability to take load diminishes with altitude. As the major load on any APU is the air load it can be seen from Figure 18.2. that the APU’s ability to provide sufficient air for the aircraft is limited to 15-20,000 ft. Above this height the APU will only provide electrical power, this may also be limited to less than the max cruise height. Most APU’s give shaft priority which means that if air and electric generators are on the generators are given priority. Most Aircraft use constant frequency generators, and their APU’s which run at a constant 100% do not therefore require a constant speed drive unit to maintain a constant output. If the air loads become to high the APU will reach its max EGT and the control system will back off the fuel to prevent damage, this would bring the APU generator off frequency and take the generator ‘off line’. Instead the air load is reduced to maintain a constant APU speed. 18.2 GENERAL ARRANGEMENTS AND CONFIGURATION With the configuration shown in figure 18.3. we can see that air is taken from the compressor via the load control valve (LCV) when pneumatic power is required. Although such an APU layout is acceptable on smaller aircraft where pneumatic power demand is small, it is unacceptable on larger aircraft as the air being drawn from the compressor for pneumatic purposes, reduces the air going to the turbines for cooling purposes. This reduction of cooling air leads to an increase in exhaust gas temperature and a reduction in the life of the turbine.

A Basic Electronically Controlled APU Figure 18.3. Issue 2 – April 2003

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On larger models of APU this problem of reduced turbine life has been reduced by the inclusion of a load compressor. See figure18.4.

Block Diagram of an APU with a Load Compressor. Figure 18.4.

In this configuration, the inlet air is directed into the load compressor as well as into the power section compressor. The load compressor now satisfies all pneumatic loading requirements without extracting any air from the power section. This can best be explained by looking at figure.18.5. This figure represents a typical cross section of an APU with a load compressor. The power section with two centrifugal compressor stages and a centrifugal load compressor both being driven by the turbine. The load compressor produces pneumatic pressure when a demand is made on the system.

Cross-section of an APU with a Load Compressor. Figure 18.5.


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A combination of the previous two examples can also be found, see figure18.6.

A twin shaft APU with Variable output of Air. Figure 18.6.

The location of the APU on the aircraft is generally dictated by the requirements of the manufacturer. Because of the noise factor and the problem of hot exhaust gases, it is located as far away from ground servicing areas as possible. The normal place for it to be fitted is in the tail section of the aircraft, however, this may be impracticable due to the location of a tail mounted engine or airstairs. On some aircraft the APU may be fitted into landing gear bays, engine nacelles, forward fuselage or wing structures. Examples of these are Hercules (U/C bay), Fokker F50 (rear of engine nacelle) and BAe ATP (wing fillet)


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An APU Installation (Airbus A300) Figure 18.7.

Light Alloy APU Intake Duct Without an Intake Door. (BAe 146) Figure 18.8.


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18.2.1 INLET DUCT ARRANGEMENT

Wherever the APU is located, ducting will be required to bring air to the APU inlet. In figure 18.9. we can see that the inlet duct connecting the inlet door to the APU plenum chamber is divided into three parts. The plenum chamber has the APU inlet duct bolted to its structure, thus reducing a complicated duct joint arrangement. These ducts can be manufactured from various materials, but the most common are aluminium, titanium, steel or composite (fibre glass/carbon). Figure 18.8. shows a light alloy side mounted intake duct without an intake door. When the duct length is short, steel or titanium ducts may be used. When ducts cover a large distance an unacceptable weight problem may result. Ducts of this length are therefore manufactured from light alloy or composite materials. One of the main problems of APU’s is the ingestion of foreign objects this can be eliminated by fitting wire mesh grills either in the ducting, or around the APU air inlet (figure 18.8.). The length of the inlet ducts will depend upon the location of the APU and its distance from the inlet. Some APU inlets are fitted with a door, these are usually forward facing or top mounted inlets. The door will open before the APU starts and close after a time delay on APU shut down The duct may be short or fairly long as shown in the figure 18.9.

Long APU Inlet Duct with Intake Door. Figure 18.9.


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GAS TURBINE

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ENGINES Operation of the door opening and closing is achieved by using an electrical actuator, which receives its signal from a command from the flight deck APU switch. In the event of an electrical failure to an actuator, there is normally incorporated into the actuator a means of disengaging the clutch drive mechanism. This enables the actuator to be manually turned to open or close the inlet door. A proximity switch ensures that the door is fully open before the APU start sequence is initiated.

APU Door. Figure 18.10.

APU inlet doors serve three functions: 

They seal off the inlet duct from harmful weather conditions and foreign objects when the APU is not in use.



They open to allow air into the APU when the start sequence is initiated.



They can be used to adjust the intake area when on ground in flight.

A Variable Intake Door. Figure 18.11.

The variable intake door figure 18.11. is used to reduce the ram air entering the APU intake ducting. This could effect the APU fuel system if intake pressure is not taken into the calculation of engine fuel scheduling which is the case with most APU’s . Issue 2 – April 2003

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18.2.2 EXHAUST DUCT ARRANGEMENT

Exhaust ducts are invariably positioned to ensure that on the ground as the hot gases are directed away from the maintenance crews and aircraft structure. This is usually achieved by angling the exhaust duct upwards. Figure 18.12. represents a typical duct arrangement. The exhaust ducts are subjected to high temperatures, so the following design features must be considered: 

Leaf springs are fitted to allow for longitudinal expansion of the exhaust duct.



The flexible bellows allow for slight variations during the assembly of the duct to the engine flange.



Flame traps may be fitted to joints to provide protection if the joint leaks.

The exhaust duct is normally insulated to prevent the heat from affecting the aircraft structure or adjacent components. This can be a double duct with cool air being passed between the ducts or by the use of insulation blankets. An exhaust door may be fitted to reduce cold soak or to prevent rain or snow entering the duct. The door must be open before the engine can start and will close after a time delay on shut down.

An APU Exhaust Duct. Figure 18.12.


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18.3 THE APU ENGINE APU engines usually consist of a centrifugal compressor and a radial turbine however, axial compressors and turbines may be added or used in their own right. Centrifugal compressors are used because of their high compression and small size and when combined with a radial type turbine make the APU very compact. These components are also very robust and require less maintenance than axial flow components. Use is also made of reverse flow combustion chambers that again makes the overall size smaller.

A Honeywell GTCP 36-100 Series APU Figure 18.13.

In most cases there is a design compromise made between the ideal APU for an aircraft i.e. its ability to provide air and electricity throughout the operational envelope of the aircraft, and it weight and size. It is usual therefore to find that air and electricity are limited to various altitudes dependant upon the parameter required. Issue 2 – April 2003

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APU systems are very basic and the APU will shut down if a problem is sensed. Most APU’s will shut down for the following faults: Fault

Comment

Low Oil Pressure Engine Overspeed

110% Honeywell 108% Sundstrand

High EGT

849°C Honeywell

Loss of Speed Signal

Electronically monitored APU’s need this signal to control the APU.

Los of EGT Signal

Electronically monitored APU’s need this signal to control the APU.

Low Speed

Some APU’s shutdown if they drop below 90%. (Some will try and relight at 95%) Loss of Control

Electronic Unit Failure

The APU may also shut down on the ground (not in flight) for the following faults: Fault Fire

Comment May cause a warning horn in the u/c bay to sound

Generator Drive Low Oil Pressure or High Oil Temperature


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18.4 FUEL CONTROL There are three types of APU fuel control, mechanical, electronic and the Electro/mechanical. 18.4.1 MECHANICAL FUEL CONTROL

An APU Fuel System Schematic. Figure 18.14. The basic fuel system is comprised of a fuel pump that receives low-pressure fuel from the aircraft fuel tank via a low pressure fuel valve and pumps it at a higher pressure to the fuel nozzles as shown in figure 18.14. Since the nozzle has resistance to flow, the fuel pressure rises in the fuel line between the pump and nozzle. The fuel is divided into primary and secondary flow by a fuel flow divider before being sprayed into the combustor and, with the addition of a spark, then combustion is initiated. The fuel pump is designed to supply more fuel than required by the APU as shown in the figure 18.15.


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Graph Showing the Excess Fuel Pump Capacity. Figure 18.15. The upper line represents fuel flow from the pump. As pump speed increases so does the pump output capacity. The lower line represents APU fuel requirements. Some means must be available to remove the excess fuel capacity.

Fuel System with a By-pass Valve Figure 18.16. By adding a by-pass valve a method of controlling the fuel pressure and thus the engine. If the by-pass valve is closed, all the fuel is directed to the nozzle. Opening the by-pass valve will allow fuel back to the inlet of the pump, thus reducing the fuel to the nozzle. By controlling the by-pass valve, the operator can vary the amount of fuel to the nozzle. Issue 2 – April 2003

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Such control in fact is done automatically by the APU fuel control unit.

Pneumatic Control of the By-pass Valve Figure 18.17.

Figure 18.17 fuel pressure is applied to the lower part of the by-pass ball valve. An air tapping which protrudes into the compressor airstream, applies pressure to the upper part of the by-pass valve diaphragm, thus holding the valve on its seat. Therefore fuel pressure is limited by the air pressure. When initial ignition takes place within the APU, there is little air pressure, so fuel pressure cannot rise very much without pushing the valve open and allowing the excess fuel to go to the pump inlet. Because of the size of the diaphragm and valve, the air pressure allows the fuel pressure to rise by a proportional amount, thus fuel and air pressure stay in step with each other. As engine speed increases:

Compressor pressure rises.



Fuel pressure rises.

A minimum fuel pressure is required for good fuel atomisation at the fuel nozzle for initial ignition. This is achieved by applying a spring pressure to the by-pass valve, thus keeping it on its seat.


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Figure 18.18. shows a solenoid operated shut-off valve fitted between the FCU and the fuel nozzle. Normally spring-loaded closed; it receives its open and close signals from the APU control unit at certain speeds. On a mechanical APU it is signalled open by the low oil pressure switch when oil pressure is sensed. In an electronic system it is open at speeds above 10%. On receiving a closed signal, the solenoid de-energises and the valve closes, the flow to the combustor is blocked. The build-up in pressure in the fuel line is relieved by the by-pass valve, acting as a pressure relief valve.

A Fuel Shut-off Valve is added Figure 18.18. As the engine accelerates, some means must be provided to enable more fuel to be injected into the combustor. This is achieved by a flow divider, figure 18.19.

Fuel Flow Divider Added. Figure 18.19.


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In the flow divider, one nozzle is placed within the other and separated by a small pressure relief valve. The flow divider is set for a slightly higher pressure than the by-pass valve spring pressure, thus on initial light-up, fuel will only spray from the primary nozzle. After light-up, rising compressor pressure increases the by-pass valve setting and the fuel pressure increases to force the flow divider off its seat. This allows fuel flow through to the secondary nozzle as well as the primary nozzle. During start and acceleration, the APU must produce temperatures that are within certain limits, while at the same time allow the engine to accelerate. Despite the fact that fuel pressure is kept in step with rising compressor pressure (through the by-pass valve), turbine over temperature is possible during certain acceleration phases. As a protection against over temperature, a thermostat (known as acceleration thermostat) is connected to the air pressure line, leading to the by-pass valve, this thermostat is normally closed (see figure 18.20).

An Acceleration Thermostat Prevents Overheat During Acceleration and Overall Temperature Limitation. Figure 18.20

Provided the EGT remains below the thermostat setting, it will remain fully closed. If the EGT exceeds its setting, the thermostat will gradually open and bleed off air pressure that is acting on top of the diaphragm of the by-pass valve.


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This reduced air pressure against the by-pass valve diaphragm will allow the fuel pressure to lift the by-pass valve and direct excessive fuel pressure back to the inlet of the pump. As the fuel pressure drops across the nozzles, the turbine temperature drops until the thermostat closes at a lower safe limit. The acceleration thermostat provides a continuous monitor to prevent the APU engine overtemping. A second pneumatic thermostat is fitted to control the air load valve (see figure 18.29.) which is similar to the acceleration thermostat. The thermostat can be adjusted in two ways, shimming or vernier adjuster. Shimming requires careful calculations to set the correct pressure on the ball. The vernier type adjuster has indications around the top of the thermostat, when it is unlocked the top can be twisted to make the adjustment.

A Pneumatic Thermostat. Figure 18.21


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18.4.2 SPEED CONTROL

Because the APU is designed to run at a constant rpm, some means must be provided to control this speed. Such a control device is known as a speed or rpm governor (see figure 18.22).

A Fuel System with a Governor Fitted. Figure 18.22

The speed governor is linked mechanically to the APU drive. As speed increases above 95%, the bob weights start to move outwards and begin to by-pass the fuel back to the inlet of the pump and as speed increases up to 100% rpm, it causes sufficient fuel to be by-passed by the governor, to maintain this rpm. Increase or decrease in the speed setting is achieved by adjustment of the governor spring. Note that at speeds below 95% rpm the by-pass valve controls the acceleration up to a maximum speed of 95% rpm.


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A Typical Mechanical APU Fuel System. Figure 18.23 18.4.3 MECHANICAL FUEL CONTROL UNIT OPERATION

The fuel flow control unit operation is as follows: 

Fuel is supplied to the pump from the aircraft fuel tank via an electrical shut off valve which opens when start is selected and closes when the APU shuts down.



At a predetermined speed (as dictated by the low oil pressure switch), the fuel shut-off valve opens and fuel is supplied to the combustor (5-10%).



The quantity of fuel supplied is scheduled by the by-pass valve, which senses compressor discharge pressure.



As rpm increases, compressor discharge pressure increases, reducing the bypass flow, hence more fuel to the combustor.



If high gas temperature is sensed, the acceleration thermostat opens and vents compressor pressure from the by-pass valve, thus reducing fuel flow to the combustor.



As the speed approaches 100% the governor backs off the fuel flow to slow the acceleration and to maintain 100%



During normal operation, the governor senses APU rpm and regulates the fuel flow by bypassing some back to the pump, to maintain a constant speed.


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18.4.4 ELECTRONIC APU FUEL CONTROL

Electronic fuel control emulates the mechanical system, however it provides control in a slightly different way. The electronic Control Unit (ECU) monitors the APU speed and EGT continuously and also the low oil pressure switch.

An Electronic APU Fuel Control System Figure 18.24 During start the ECU knows the engine speed so will signal the fuel shut off valve to open at 8-10%. At the same time ignition is selected on and the light up will be sensed by the EGT system. The ECU then enters a timed acceleration schedule where EGT and speed are monitored by the ECU. The ECU deselects the starter at 50% and the ignition at 95%. Once up to speed the ECU keeps the engine at 100% and will monitor the EGT and speed to maintain operation throughout the operating envelope of the APU. The Fuel control unit mounted on the APU gearbox is much simpler than the mechanical FCU. It contains a fuel pump, an electronic servo valve and a pressure drop control valve (  [delta] P valve). The electrical shut of valve and the fuel flow divider are retained and work as they did in the mechanical system.


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ECU Start Schedule. Figure 18.25

The ECU has Built In Test Equipment (BITE) indicators which will indicate why the APU shut down, however these do not retain the information if power is removed. 18.4.5 ELECTRO/MECHANICAL FUEL CONTROL (FIGURE 18.26)

The start fuel valve and ignition are energised as soon as rotation (3%) is sensed by an Electronic Sequence Unit (ESU). At 14% and with rising EGT the main fuel valve is opened. The acceleration rate is controlled by the acceleration schedule adjuster, however this is modified by the differential pressure regulator which uses compressor discharge pressure to vary the fuel flow to the engine. At 50% the starter cuts out. When the engine reaches 85% the start fuel valve closes and the ignition is de-energised. The engine governor then takes over and controls the engine to 100%. As the engine passes 95% plus 3 seconds, the max fuel valve energises open and bypasses the acceleration adjuster and full control of the engine is given to the governor. If the engine is shut down both the Main and Max fuel valves are closed.


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An electrically Sequenced APU Fuel Control. Figure 18.26.

The ESU has indicators that indicate which step of the start sequence the APU is at and the resets at 95% + 3sec to act as BITE indicators.


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18.5 APU OIL SYSTEM A sump at the bottom of the gearbox collects the returning oil, in some APU's the rear face of the sump is finned and let into the intake plenum to act as the oil cooler. The oil is drawn up by the oil pump and pressurised, it then passes through the oil filter before being distributed to the bearings. The oil returns to the sump by gravity. The oil system is monitored by a low oil pressure switch and a high oil temperature switch, either of which can shut the engine down.

Honeywell Oil System. Figure 18.27.


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A Sundstrand Oil System Figure 18.28.


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18.6 APU BLEED AIR SYSTEMS There are two main methods of providing bleed air: 1. Direct from the engine compressor 2. A Separate Load Compressor. 18.6.1 DIRECT FROM ENGINE COMPRESSOR

A Load valve (Figure 18.29) is switched on from the flightdeck, power for the switch is available once the APU has achieved 95% + 3 sec. This energises the switcher valve solenoid, which vents the lower chamber (B) of the control piston and pressurises the top chamber (A). The piston will move down and open the butterfly valve. The bleed air will flow and the EGT will rise, at a predetermined value the Load Thermostat will start to open which will reduce the pressure acting on the top of the piston. This will cause the piston to move up by spring pressure and thus back off the butterfly valve. If the EGT rise is excessive then it could close the valve. The valve will modulate under the control of EGT. The Load thermostat is set at a lower setting than the acceleration thermostat setting to prevent hunting of the system.

A Mechanical Load Control Valve Schematic. Figure 18.29.


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An electronically controlled APU uses the same principle, but the ECU controls a servo valve in the load control valve instead of the load thermostat, see figure 18.30.

An Electronically Controlled Load Valve Schematic. Figure 18.30.

Some APU's do not use load valves, instead they have an air bleed valve which is a simple on/off valve. A flow limiting venturi is used to limit the flow of air from the APU


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18.6.2 SEPARATE LOAD COMPRESSOR

If the APU is fitted with a load compressor either of the previous two methods are used, but instead of controlling a butterfly valve the piston operates a set of variable intake guide vanes for the load compressor, see figure 18.31.

Load Control for APU with Load Compressor Figure 18.31.


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18.7 BAY COOLING There are two methods of bay cooling, they are: 

Ram air cooling



Fan air cooling

18.7.1 RAM AIR COOLING

For ram air cooling, the aircraft has to be moving forward at sufficient speed to enable the cooling air to be picked up by the air scoops in the external skin. This cold air is ducted into the APU bay and passed onto various hot zones to provide a cooling medium. The air is then vented overboard through exhaust ducts. 18.7.2 FAN AIR COOLING

Cooling fans are fitted to the APU gearbox to provide a supply of cooling air to the APU when it is running. The cooling air is pumped into the APU compartment and then vented overboard. The air from the fan is also used to cool the generator drive oil and the exhaust duct on some APU installations.

An APU Cooling Fan System. Figure 18.32.


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engineering

ENGINES

The main components are the:

cooling air fan.



cooling air shut-off valve.

Air is drawn from the normal intake plenum or an external intake and is directed along the cooling air ducts to the cooling fan shut-off valve (when fitted). The shutoff valve closes on APU shutdown to prevent air from entering the compartment to support combustion in the event of an APU fire. The cooling fan is linked to the APU gearbox and as long as the APU is running, the fan is turning. Air is also used to cool the oil within the APU lubricating system (on some APU’s), however, such air is usually ducted overboard and not into the APU compartment. Upstream of the oil cooler the cooling air is ducted into the APU bay an/or the exhaust insulating ducting to provide general cooling. Cooling Fan Shut-Off Valve The cooling valve figure 18.33. is a spring-loaded closed butterfly valve with a pneumatic actuator. When the APU is started, the compressor discharge pressure is ported to the top of the diaphragm. The piston moves down with increasing air pressure and opens the valve against the spring pressure. The cooling air then flows to the compartment. On APU shutdown the air pressure is reduced and spring pressure closes the valve.

Cooling Air Shut Off Valve. Figure 18.33.


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Cooling Fan Arrangement The cooling fan is attached to the APU gearbox, (figure 18.34) and is designed to run at extremely high speeds, the fan boosts the air from the intake plenum (or ambient) into the APU compartment or the coolers etc.

APU Cooling Fan. Figure 18.34.


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Overboard Venting Figure 18.35 represents a typical APU bay overboard vent arrangement. The cooling air is directed into the compartment and also to the oil cooler, this air is then vented overboard along a separate duct. Compartment cooling air is vented overboard, through a louvered door at the rear of the compartment.

Vent System Figure 18.35. .


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18.8 APU POWERPLANT INSTALLATION. The APU engine mounts consist of a number of supports with vibration isolators fitted to the end of each support. The tubular supports are bolted to the plenum chamber and when correctly attached, hold the APU against the air inlet duct in the plenum. The vibration isolators dampen out any vibration effects that the APU would have on the aircraft structure whilst it is running. Attached to the vibration isolator is a cone bolt that passes through a similar hole on the APU mounting bracket. When in position, the bolt is secured by a nut and washer arrangement and torque loaded to the set figure laid down in the Aircraft Maintenance Manual. (Figure 18.36).

APU Mount. Figure 18.36.


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A Shrouded APU. Figure 18.37.

Most APU’s are located in a fire proof box made of titanium. Some aircraft have the APU shrouded in a close fitting Titanium case.


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18.9 APU STARTING SEQUENCE

A Typical APU Start System. Figure 18.38

In figure 18.38, the APU control unit receives its power from the aircraft battery. By moving the APU switch to ‘ON’, power is provided to the intake door actuator and an LP fuel valve which starts to open. When they are both fully open, switches energise the starter system, igniters are energised and the APU accelerates with assistance from the starter motor to idle speed. The starter motor cuts out between 50 & 60%. The fuel system controls the fuel flow during start. Once the engine is at idle (100%, lower if the APU is fitted with an idle power setting), the APU will be ready to load either electrically or pneumatically. This is indicated by a ‘Ready to Load’ light or ‘APU Power Available ‘light coming ’ON’.


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19 POWERPLANT INSTALLATION 19.1 NACELLES OR PODS Nacelles or pods are streamlined enclosures used on multi-engine aircraft primarily to house the engines. They are located below, or at the leading edge of the wing or on the tail of the aircraft. An engine nacelle or pod consists of skin, cowling, structural members, a fire-wall, and engine mounts. Skins and cowlings cover the outside of the nacelle. Both are usually made of sheet aluminium alloy, stainless steel, or titanium. Regardless of the material used, the skin is usually attached to the framework by rivets. The framework can consist of structural members similar to /those of the fuselage. The framework would include lengthwise members, such as longerons and stringers, and widthwise/vertical members, such as bulkheads, rings, and formers. A nacelle or pod also contains a firewall, which separates the engine compartment from the rest of the aircraft. This bulkhead is usually made of stainless steel, or titanium sheet metal. 19.1.1 COWLINGS

Openings in structures are necessary for entrance and egress, servicing, inspection, repair and for electrical wiring, fuel and oil lines, air ducting, and many other items. Access to an engine mounted in the wing or fuselage is by hinged doors; on pod and turbopropeller installations the main cowlings are hinged. Access for minor servicing is by small detachable or hinged panels. All fasteners are of the quickrelease type. A turbo-propeller engine, or a turbo-jet engine mounted in a pod, is usually far more accessible than a buried engine because of the larger area of hinged cowling that can be provided. The accessibility of a wing pylon mounted turbo-fan engine is shown in figure 19.1. and that of wing mounted turbo-propeller engine is shown in figure 19.2.


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Turbofan Nacelle and Cowlings. Figure 19.1.


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Turboprop Engine Nacelle and Cowlings. Figure 19.2.


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19.1.2 FIREWALLS

The firewall is a seal which separates the engine into two zones. Sometimes referred as the “wet zone” and “dry zone”, but more commonly called zone one (front) and zone two (rear). The firewall forms a barrier that prevents combustible fumes that may form in the front section (zone 1), from passing into the rear section (zone 2), and igniting on the hot exhaust section. Dependant upon aircraft/engine design the fire walls design and location will differ, Figures 19.3. and 19.4. refer.

A Turbofan Firewall. Figure 19.3.


Page 19-4 Turboprop Firewall. Figure 19.4.

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19.1.3 COOLING

Turbine engines are designed to convert heat energy into mechanical energy. The combustion process is continuous and, therefore, heat is produced. On turbine engines, most of the cooling air must pass through the inside of the engine. If only enough air were admitted into a turbine engine to support combustion, internal engine temperatures would rise to more than 4,000 degrees Fahrenheit. In practice, a typical turbine engine uses approximately 25 percent of the total inlet airflow to support combustion. This airflow is often referred to as the engine's primary airflow. The remaining 75 percent is used for cooling, and is referred to as secondary airflow. When the proper amount of air flows through a turbine engine, the outer case will remain at a temperature between ambient and 1,000 degrees Fahrenheit depending on the section of the engine. For example, at the compressor inlet, the outer case temperature will remain at, or slightly above, the ambient air temperature. However, at the front of the turbine section where internal temperatures are greatest, outer case temperatures can easily reach 1,000 degrees Fahrenheit. (Figure 19.5.) Cooling Requirements To properly cool each section of an engine, all turbine engines must be constructed with a fairly intricate internal air system. This system must take ram and/or bleed air and route it to several internal components deep within the core of the engine. In most engines, the compressor, combustion, and turbine sections all utilise cooling air to some degree. For the most part, an engine's nacelle is cooled by ram air as it enters the engine. To do this, cooling air is typically directed between the engine case and nacelle. To properly direct the cooling air, a typical engine compartment is divided into two sections; forward and aft. The forward section is constructed around the engine inlet duct while the aft section encircles the engine. A seal or firewall separates the two sections.

Diagram Showing the Temperature That May be Present Around a Turbojet Engine in Degrees Fahrenheit. Figure 19.5. Issue 2 – April 2003

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In flight, ram air provides ample cooling for the two compartments. However, on the ground, airflow is provided by the reduced pressure at the rear of the nacelle. The low pressure area is created by the exhaust gases as they exit the exhaust nozzle. The lower the pressure at the rear of the nozzle, the more air is drawn in through the forward section.

Typical Nacelle Cooling Using Ram Air From the Intake Duct. Figure 19.6.


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19.1.4 ACOUSTIC LININGS

One method of suppressing the noise from the fan stage of a high by-pass ratio engine is to incorporate a noise absorbent liner around the inside wall of the bypass duct. The lining comprises a porous face-sheet which acts as a resistor to the motion of the sound waves and is placed in a position such that it senses the maximum particle displacement in the progression of the wave. The depth of the cavity between absorber and solid backing is the tuning device, which suppresses the appropriate part of the noise spectrum. Figure 19.7. shows two types of noise absorbent liner. Figure 19.8. shows the location of a liner to suppress fan noise from a high by-pass ratio engine and also the use of a liner to suppress the noise from the engine core. The disadvantage of using liners for reducing noise are the addition of weight and the increase in specific fuel consumption caused by increasing the friction of the duct walls.

Two Types of Acoustic Panel Figure 19.7.


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Location of Acoustic Panels in a High ByPass Engine Figure 19.8.

Acoustic Panel Location in a Fan Module. Figure 19.9. Issue 2 – April 2003

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Section Through an Engine Nacelle Figure 19.10.


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19.1.5 ABRADABLE LININGS

Abradable Linings are usually made of a composite material which will be abraded away should the tip of a rotating blade touch the material. In flight the casings of an engine are subject to large changes in ambient temperature, so they will expand or contract. As we know the air temperature at 30,000ft is close to –50°C this would cause the casings to contract onto the rotor and the blades will then rub. To overcome this problem abrasive materials where used on early engines to wear down the tip of the blades, but this may cause balance problems. So most engines now use abradable linings that maintain minimum tip clearance but do not affect balance. They are usually found on the fan as this is the cold area of the rotating assemblies.

Abradable Lining Location in a Fan ModuleFigure 19.11.


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19.2 ENGINE MOUNTS Engine mounts are designed to meet particular conditions of installation, such as the location and the method of attachment of the engine mount and the size, type, and characteristics of the engine it is intended to support. An engine mount is usually constructed quickly and easily from the remaining structure. Engine mounts are commonly made of welded chrome/molybdenum steel tubing, and forgings of chrome/nickel/molybdenum are used for the highly stressed fittings. 19.2.1 WING PYLON MOUNTED ENGINE (TURBOFAN)

Figure 19.12. shows a typical method of mounting an engine onto a wing pylon. The engine is usually suspended on three attachment points. The two front points are located at the lower end of a pylon mounted yoke and engage with the mounting bracket assemblies on the left-hand and right-hand side of the fan casing. The assemblies differ inboard and outboard. The inboard bracket assembly takes side, vertical and thrust loads. The outboard bracket assembly takes vertical and thrust loads. The rear attachment point is an engine mounted lower link assembly bolted to a pylon mounted upper link assembly. This attachment point carries vertical loads only and allows for engine axial expansion.


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Wing Pylon Mounted Engine Mounts. Figure 19.12.


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19.2.2 WING MOUNTED ENGINE (TURBOPROP)

The engine is connected to the structure by means of a flexible attachment system consisting of: 1. 2 forward lateral shockmounts. 2. 1 forward upper shockmount. 3. 2 aft lateral shockmounts on the Left Hand and Right Hand sides. A torque compensation system with a torque tube.


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Wing Mounted Turboprop Engine Mounts. Figure 19.13.


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19.2.3 REAR FUSELAGE ENGINE TURBOFAN.(FIGURE 19.14/15.)

Two crane beams in the nacelle carry the weight of the engine. The crane beams are connected to the frames of the fuselage. Vibration isolators are on the engine mounting Points to absorb vibration. There are three mounting points: 

the rear mount.



the front mount



the trunnion

The trunnion transmits the engine thrust to the airframe. The Trunnion fits in the trunnion housing on the forward crane bean attachment. Between the trunnion housing and the aft beam attachment is a thrust strut, This strut divides the engine thrust between the forward and aft beams attachment. The shear shell between the crane beams makes the engine mounting more rigid.

Rear Fuselage Turbofan Engine Mounts. Figure 19.14.


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Fuselage Mounted Engine Mounts in Detail. Figure 19.15.


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19.3 ENGINE DRAINS. There are two types of drains: 

Controlled drains – the result of normal operation.



Uncontrolled drains – the result of abnormal operation.

19.3.1 CONTROLLED DRAINS

When an engine stops, fuel from the fuel manifold and combustion chamber drains either overboard, or as is more usual into an ’ecology drain tank’. This tank is automatically emptied, (the fuel being fed back into the engine) next time the engine is run. (figure 19.16.)


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Controlled Drains System. Figure 19.6. Issue 2 – April 2003

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19.3.2 UNCONTROLLED DRAINS

Engine driven accessory drive shaft require lubrication. This will be provided by the engine lubrication system. To ensure proper lubrication, the drive shaft bearings are sealed to prevent loss of oil. These bearing seals are monitored for leaks, by the engine drain system which consists of a number of shrouds, enclosing the drive shaft bearing, and pipes leading either an overboard series of drain pipes (figure 19.17.) or a collector tank (figure 19.18.). These drains are often referred to as ‘witness drains or dry drains’ as if they exhibit signs of leakage they bear witness to a potential drive shaft failure.

Uncontrolled Drains With a Drains Mast. Figure 19.17.


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A Typical Drains System. Figure 19.18.


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19.4 ENGINE CONTROLS 19.4.1 THROTTLE CONTROL MECHANICAL

Engine controls are very similar to flying controls, and the same types of equipment are used, such as rods, bellcranks and cables. Most control systems use either one or two systems to control the engine. In a two path system the high pressure cock is controlled separately from the throttle, in a single path system they are combined. 19.4.2 TURBOFAN ENGINE CONTROLS.

Figure 19.19. shows a typical mechanical control system for a turbofan powered aircraft. It uses a single path system to transmit power requirements to the engine. The thrust lever is connected to a rod that transmits the movement down below floor level to a quadrant. The quadrant outputs to two cables which initially run under the floor of the flightdeck and then along the roof of the passenger cabin. They then pass through pressure seals and along the leading edge of the wing before dropping down to a cable compensator in the top of the pylon. The output from the compensator quadrant is a teleflex push/pull cable. This teleflex cable passes down into the engine nacelle to a torque shaft mounted on the nose cowl assembly. The output from the torque shaft moves a rod which provides the input to the fuel control unit. The teleflex cable has a disconnect break mechanism in it to facilitate engine changes. To allow autothrottle functions the quadrants below the thrust levers can be moved by an actuator which drive all four levers via clutches.


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A Typical Mechanical Engine Control System. Figure 19.19


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19.4.3 TURBOPROP ENGINE CONTROLS

Figure 19.20. shows a typical mechanical control system for a turboprop engine. It uses a double path system to transmit power requirements to the power unit,i.e. the power lever controls engine power in the normal operating modes and both power and propeller blade angle in the beta mode. A condition lever controls propeller blade angles in the normal mode, and also controls the feathering of the propeller and the HP shutoff cock.

Power and Condition Levers. Figure 19.20. Issue 2 – April 2003

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Turboprop Power Control System – Cable Routing. Figure 19.21.


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Power Controls. (figure 19.22.) The power lever controls, via the Hydromechanical Control Unit (HMU)the full flow from “MAX” (maximum power) to “REV” (reverse) (Figure 19.23.). Power lever movement is transmitted to the HMU via a series of push/pull rods and cables. A control rod between the HMU and the Propeller Control Unit (PCU) enables control of propeller blade angle in beta mode. Propeller/HP Shutoff Cock Control. (figure 19.22.) The “Condition Lever” controls via the PCU propeller speed from, “Min N P” (minimum propeller speed) to “Max NP” (maximum propeller speed). Condition lever movement is transmitted via a series of push/pull rods and cables, similar to the power lever controls. A second control rod (figure 19.23.) between the PCU and HMU enables control of the HP fuel shutoff cock within the HMU by the condition lever. The condition lever also controls feathering of the propeller (figure 19.22)

Power lever Controls: 1. Power in forward mode (NH or SHP as a function of PLA) 2. NP in reverse. 3. Propeller blade angle in beta. Condition Lever: 1. Fuel “on” or “off”. 2. Feathering or unfeathering the propeller. 3. NP from minimum to maximum.

Power and Condition lever Controls. Figure 19.22.


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HMU to PCU Connections. Figure 19.23.


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Intentionally Blank


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19.5 ENGINE BUILD UNIT When an engine is delivered from manufacturer or overhaul it will not have all the equipment needed for its installation into the aircraft. This is because engines can be fitted into different types of aircraft and the accessories will be type specific. Hydraulic pumps, electrical generators, starters, drains and mounts will have to be fitted during or prior to installation in the aircraft. Although the engines fitted to each wing are the same, the accessories and their fittings may well be handed for the different installations i.e. the BÆ 146 has a generator on the outboard engines and a hydraulic pump on the inboard. These components are referred to as dress items, an engine that is dressed is ready for fitment. For some engines fitting the accessories prior to fit on the aircraft is impractical and the accessories are fitted once the engine is installed. Examples of engine build units are shown in Figures 19.24. to 19.27. together with a list of items and components that must be fitted before the engine is considered ready for release to service prior to installation into the aircraft. 19.5.1 TURBOFAN ENGINE

The manufacturer delivers the engine to fit the no-2 (right) position. Conversion from the no.2 (right) to the no.1 (left) position requires re-position of: 

The front engine mount adaptor.



The trunnion mount.



The HP compressor 7th and 12th stage bleed air ducts.



The electrical harness on the engine.



The external igniter leads on top of the engine.



The engine vibration transducer wiring.


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Power Plant Build Installation.(Tay) Figure 19.24.


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Number

Item

10

Front Mount Adapter

20

Anti-Icing System

30

Vibration Transducer

40

Hydraulic Lines

50

Inlet Cowling

60

Hydraulic Hoses

70

Hydraulic Pump No. 1

80

Hydraulic Pump No. 2

90

Integrated Drive Generator

100

Vent and Drain System

110

Starter System,

120

Air-Starter Duct,

120A

Air-Starter Duct

130

After Cowling

140

Fuel Flow Transmitter

150

Fuel Line

160

Engine Control Rods

170

Power Lever Angle Transmitter

,


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Electrical Harness Installation.(Tay) Figure 19.25.


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Number

Item

10

Igniter Leads

20

Igniter Leads

30

Anti-Ice Electrical Harness

40

Anti-Ice Electrical Harness

50

Electrical Harness on the Hydraulic Pumps No. 1 and 2

60

Electrical Harness on IDG and IDG Oil Temperature Switch

70

Vibration Transducer Electrical Harness, LH-Engine

80

Vibration Transducer Electrical Harness, RH-Engine

90

Electrical Harness on Fuel Flow Transmitter

100

Electrical Harness on PLA-Transducer

110

Fire Detection Element


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Turboprop Build Left Hand Side.(PW125) Figure 19.26. Issue 2 – April 2003

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engineering Number 10. 20. 30. 40. 50. 60. 70. 80. 90 95. 100 110 120. 130. 140. 150 160 170 180 190 200 210


Item Engine Mounts - Forward Isolators Engine Mounts - Forward Frame Assy IDG Assy IDG Support Bracket Pitch Control Unit and Control Rods Lever Bracket and Interconnection Rods Bleed Air - Low Pressure Check Valve Electrical Harness Bleed Air, High Pressure Bleed Valve Heat Shield Installation Back-up Firewall Bleed Air - Low Pressure Off-Take Female Flange - Exhaust Main Fuel Supply Tube Drain Hoses Pipe Lines Installation for Oil Pressure Transducer & Oil Pressure Switch Oil-Pressure Transducer, Oil-Pressure Switch, Oil-Temperature Detector and Fuel-Temperature Detector Heat Exchanger Airduct and LHS & A-Frame Oil-Cooler Assy Propeller Spinner


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Turboprop Build Right Hand Side.(PW125) Figure 19.27


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engineering Number 220 230 240 250 260 270 275 280 290 300 310 320 330 340 350 360 370 370A


Item Vertical Firewall Bleed Air - High Pressure and Low Pressure Fire Extinguisher Tube Starter Motor Hydraulic Hose Assemblies and Hydraulic Pump Feathering Pump Brush Block Drain Tubes Torque Tube Isolator Air Intake Engine Seal Assy Hydraulic Pump Seal Drain Fuel Flow Transmitter Oil Drains Fuel Lines on the Engine Spray Pipe for Air Intake Engine Mounts Engine Mounts - Rear Isolators


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19.6 FIRE PREVENTION – BAYS OR ZONES To prevent the spread of a fire within an aircraft/engine nacelle, it is divided up into sections or zones, each being separated by a fireproof bulkhead. These are made of titanium or stainless steel and prevent the fire from spreading into adjacent areas. The engine nacelle is split into two sections (UK). Zone 1. The cool section contains the: 

Fan



Compressor



Fuel Control



Air system supply



Hydraulic pump



AC generator



Bleed valves and Variable Inlet Guide Vane (VIGV) systems

Zone 2. The hot section contains the: 

Fuel burners



Combustion chamber



Turbines LP & HP



Exhaust

Fire Zones. Figure 19.28.


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All fire zones are sealed from adjacent areas. Fire resistant rubber seals are fitted to the edges of all doors, panels and bulkhead fittings to prevent fire spreading. Each of the zones will be ventilated to prevent the build up gases or pressure and to cool the outer casing of the engine and accessories. Fire break in panels will be built in to allow the use of external fire extinguishers, these may also operate as blow out doors to prevent pressure build up in the zone.


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19.7 INSTALLING AND REMOVING ENGINES. The removal and installation of an aircraft engine follows basically the same principles. However there are differences between turboprop, turboshaft and other engines. Because of the size and complexity of engine replacement there is usually a preprinted job card to ensure the job is carried out correctly. 19.7.1 REMOVAL

To prepare an aircraft for engine removal, check that the aircraft weight and balance will not be adversely effected when the engine is removed. Most engines weigh between 0.5 and 1 ton. Trestles may be required to stabilise the fore and aft axis of the aircraft. The aircraft fuel system does not have to be drained, but the LP fuel valve must closed and a label attached to the LP Cock handle, in the flightdeck, to prevent inadvertent operation. In addition, the aircraft should be made electrically safe which will entail isolation of the engine starting and ignition system. Planning is an essential part of any engine removal activity. The Supervisor and personnel involved, should ensure that all necessary resources, such as sufficient manpower, special tools, lifting equipment and an engine transit / storage stand, are available. The engine access doors and fairings will either have to be removed or supported clear of the engine. Due to restricted access of some engine accessories and components, it is, in some cases, much easier to remove these items with the engine installed in the aircraft. Once the engine has been initially prepared for removal (accessories removed etc) the procedure of disconnecting the engine systems, at the engine/ aircraft interface, can begin. Most engines employ quick release plugs and sockets for ease of disconnection of the electrical systems, however some electrical systems, with heavier duty cables, such as the starter and generator cables, may be bolted connections. Disconnect any cable cleats going across the engine / airframe interface. The hydraulic pipes are usually quick release/self-sealing connections at both the hydraulic pump and the engine / airframe interface. Air supply connections will generally interface with a ‘vee band’ type of clamp or a bolted connection. The engine LP fuel inlet pipe must be drained, before disconnection, into a suitable container and the waste fuel disposed off in an approved manner. With the exception of the main engine bearers, all mechanical links must be released and either removed or tied back to prevent fouling during the removal operation.


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Bae 146 Engine Lift Equipment. Note. The Nose Cowling is attached to the Engine and is Removed Later. Figure 19.29.


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If the engine is not being replaced or refitted immediately, all open pipes must be blanked off to prevent foreign particle ingress and all electrical plugs tied back and protected. Once satisfied that the engine is ready for removal the lifting equipment can be fitted in accordance with the AMM. Jet engines are installed and removed utilising gantry cranes, mobile cranes or in many cases by use of 2,3 or 4 mini hoists. Whatever method is used the lifting equipment must be inspected before use. Particular attention should be paid to ensuring that the equipment has approval documentation and is of the correct ‘safe working load’ for the task. Cables should not show evidence of twisting or fraying and end fittings should be free of damage, corrosion etc. When mini hoists are used, the brake and clutch mechanisms of each hoist should be functionally checked and that the correct hoist is being used as similar units are rated at different settings. Supervisors should double check that all the lifting equipment is serviceable and correctly fitted prior to commencing the removal process. The supervisor should also carry out a final check of the engine / airframe disconnect points to satisfy himself/herself that the engine and equipment is safe for removal. Each winch / hoist is to be manned at all times during the removal process and at least one person who can check the engine to ensure it remains in a safe condition during removal. The supervisor must ensure that all team members are fully aware of the process and briefed on what is required of each individual. All instructions should be given in a clear and unambiguous manner and where hand signals are required, all members can see the supervisor and are aware of their meaning. Only the supervisor of the task should issue instructions during the process and unnecessary talk and noise (i.e. riveting operations in vicinity) minimised or stopped. Immediately prior to removing the engine and finally releasing the engine mounts / attachments, the weight of the engine must be ‘taken’ by the lifting equipment. This will ensure that there is no unnecessary ‘jerking’ or ‘snatching’ of the cables. With mini hoists this is achieved by winching the cable in until the clutch in the handle breaks (Always re-engage the handle before progressing further). At this point the effectiveness of the brake unit in the mini hoist should be checked following the relevant manufacturers procedures. Once the supervisor is satisfied that all procedures have been followed correctly and that all resources are in place the engine mountings / bearers can be disconnected and the engine removed / lowered from its housing. At all stages of the removal procedure checks should be carried out to ensure that the engine does not become caught on the airframe structure or components. WARNING NEVER WALK UNDER A SUSPENDED LOAD. EVERY EFFORT SHOULD BE TAKEN TO MINIMISE THE TIME NECESSARY TO CARRY OUT ANY MAINTENANCE BENEATH A SUSPENDED LOAD Issue 2 – April 2003

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When lowering an engine using a mini hoist system, the weight of the engine should always be taken by the winding handle and the brake should be released and held off. An engine stand should be positioned ready to accept the engine and any pins or mounts, between the engine and its stand, connected prior to allowing the weight to be removed from the winching system. If the engine is to be replaced remove any further dress items that have not already been removed. Complete and attach an equipment label to the engine detailing its condition, life used, etc. To avoid or minimise deformation on the aircraft structure due to removal of the engine, it may be necessary to fit a component called a ‘jury strut’ This requirement will be clearly stated in the relevant procedure of the AMM. Once removed further inspections on the engine and the nacelle will be carried out. If the engine is to be returned to the manufacturer these will entail blanking of exposed pipes and protection of exposed cables and components. If the engine is to be refitted to the same aircraft then these checks, often referred to as ‘bay checks’ are more involved and are designed to ensure that the condition of the hard to see areas of the engine and engine bay are thoroughly checked.


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Interface Disconnect Points. Figure 19.30a.


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Interface Disconnect Points. Figure 19.30b. Issue 2 – April 2003

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Interface Disconnect Points. Figure 19.30C.


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Interface Disconnect Points. Figure 19.30D.


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19.7.2 FITTING

Prior to fit remove the label from the engine and attach it to the paperwork for safekeeping. Check the engine over to ensure it is complete and check the label for any tasks required before fit. Fit any dress items that need to be fitted prior to fit. Check round the bay to ensure it is clear to fit the engine and remove the jury strut if fitted. Check the lift gear is correctly installed and that it is serviceable. Position the engine and correctly attach it to the lift gear (double check this). Lifting the engine in follows the same basic rule as lowering. If using mini hoists there is no need to operate the brake when hoisting as it ratchets. When the engine nears the installed position the person in charge and his assistant will align the mounts and fit the pins or bolts, this is a critical time and may require very small movements on the lifting gear to allow the mounts to be connected. Great care and concentration is required to prevent damage or injury. Do not use your finger to check alignment as a very small movement of the engine could trap or sever it. Once the mounts are made, and locked the lifting gear can be removed and the engine systems and accessories can be reconnected which is the reverse of the removal. Remember to fit new seals to the components. After engine fit the electrical systems can be reset. The LP fuel valve opened and the engine fuel system bled to remove any air. The engine oil system is then checked and followed by an engine ground run. During the ground run leak and performance checks are carried out to ensure that the engine is satisfactory. After the run the chip detectors are checked and duplicate inspection is required on the engine controls. 19.7.3 TURBO PROP ENGINE REMOVAL/FIT.

With a turboprop engine the prop would have to be removed prior to removal and fitted after the engine is mounted. The prop would also have to be bled and functioned prior to running to prevent damage. 19.7.4 FLIGHT TRANSIT

To allow an aircraft to return to a suitable base for an engine change, some multi engine aircraft can be flown with one engine shut down. In the case of the BAE 146 it has sufficient power to take off and fly on 3 engines. To prevent damage to the engine rotor locks are fitted to the LP and HP systems to prevent rotation. The starting and ignition systems must be inhibited for that engine to prevent damage by inadvertent selection.


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An ALF 502 Engine in its Stand Figure 19.31.


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20 FIRE PROTECTION SYSTEMS 20.1 FIRE DETECTORS A complete fire warning system consists of a detection system, an extinguisher system and a method of detecting that the fire is out. There are specified areas that only have detection systems, these are parts of engine bays and hot air ducts. Detectors are mounted within the zones next to the components more prone to a fire or overheat condition, the choice of detection system, fire or overheat, depends on the contents of the zone. Zone 1 (UK) contains the fuel control system, in this zone a fire could develop therefore the detection system used is a FIRE WARNING SYSTEM. Zone 2 (UK) includes the rear section of the engine and the jet pipe, this zone is identified as an overheat area only and will have an OVERHEAT WARNING SYSTEM. A fire or heat detection system should: 

Give a rapid indication of condition with an audio warning for fire (bell), the audio should have a cancellation facility and should be auto resetting.



Provide location information concerning the fire or overheat condition.



Have a warning system that will continue during fire.



Continue to operate where the fire is located.



Provide an indication that the fire is out or that the overheat condition no longer exists.



Include an “in flight” test facility.



Not automatically shut down the main power unit or operate the engine fire extinguishers, it may however shut down the APU usually only when on the ground.



Not produce false indication in event of failures or fault conditions.


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Engine Nacelle Fire Zones. Figure 20.1.


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20.2 FIRE WIRE SYSTEMS A fire wire system employs a continuous flexible sensing element that is wrapped around the potential fire or overheat areas within a fire zone. Three types of element are used; resistive, capacitive or gas pressure. The response to a temperature rise depends on the value of temperature applied and the length of sensing element to which it is applied. A high temperature over a short length or a low temperature over a long length will both operate the warning system. 20.2.1 RESISTANCE TYPE

Resistance type firewire consists of a stainless steel tube with a centre wire electrode, separating them is an insulating material of beads or powder. The resistance of the insulating material decreases with an increase in temperature until, at the warning temperature, sufficient current passes to operate the warning circuit. The element is fed with a current from a control box that also produces the output for the warning system.

Firewire sensing element. Figure 20.2. 20.2.2 CAPACITANCE TYPE

The capacitance firewire is similar in construction to the resistance type, the insulation has different characteristics. The element forms a capacitor, the capacitance of which varies with changes in temperature. The central electrode is fed with half wave alternating current which it stores and returns to a control unit during the second half of the cycle. The stored charge increases with the temperature and, when the warning temperature is reached, the back current is sufficient to operate a relay in the warning circuit. The main advantage of the capacitance system is that a short circuit grounding of the element or wiring does not result in a false fire warning.


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20.2.3 GAS OPERATION FIRE WIRE

The operating principle is the gas law i.e. pressure increases with temperature. As the helium in the sensor tube senses an overall temperature increase, its pressure is proportionately raised. Then a pressure switch operates to couple an electrical supply to the fire or overheat warning. The sensing element is pre-pressurised with helium and this lower pressure is monitored by another pressure switch that will if the base pressure is lost, indicate a failure of the sensing system. Should a localised temperature be experienced, which was of a value considerably above that needed to activate an overall temperature warning, a central core of titanium hydride will release hydrogen the tube. This action is sufficient to raise the pressure and initiate the fire warnings. As the temperature reduces the central core will re-absorb the hydrogen. Note: The detector is a hermetically sealed unit. Any attempt at disassemble it may cause serious damage and is likely to render the unit inoperative. By shutting the engine down, the fire or overheat warning should cancel as the temperature drops.

A Gas Filled Firewire System. Figure 20.3.


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20.2.4 SINGLE LOOP

One continuous loop clipped round the engine cowl in the most fire vulnerable areas. 20.2.5 DUAL LOOP

This is two independent systems usually running parallel round the engine cowl in the most fire vulnerable areas. Each fire zone has dual sensing loops. Each loop, A or B, is independent of the other. On some aircraft only one system is used at a time, the other being held as a spare. Some aircraft can use both loops at the same time, only giving a warning when both loops sense the overheat condition. (Figure 20.5) When the loop selector switch is selected to ‘BOTH’, loop A and loop B must detect a fire condition before the warning system will be activated. If only one loop detects a fire condition while the selector is at ‘BOTH’ a fire warning will not be given (some systems can give a lower grade indication of this happening). If the selector is switched to a single loop position (A or B), full fire warnings will be given if the selected loop senses fire conditions.

Loop Mounting. Figure 20.4.


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engineering

ENGINES

A Dual Loop Continuous Firewire System. Figure 20.5.

20.2.6 DUAL LOOP SYSTEMS

Dual loop fire warning systems are used to prevent spurious warnings, they consist of two identical systems. Both loops are required to detect the fire condition in order to initiate the fire warning, if only one loop detects the fire condition, only a “loop light” will illuminate. The following example shows the indications you would see on an electronic instrument system (Figure 20.6.)(E.I.C.A.S. engine indication crew alerting system), or as shown E.C.A.M. (electronic centralised monitoring system). In the example shown, the fire detection system provides the flight deck with nacelle temperature, loop faults, over-temperature and fire indication and warnings.


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Fire and Loop Fault Indication. Figure 20.6.


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20.3 FIRE AND LOOP FAULT INDICATION (E.C.A.M.) The fire detection control electronics module is made up of two circuits which process signals from fire detection loops. The loop fault circuit indicates a fire detector loop fault to the flight crew. The E.C.A.M. responds with a loop fault message on the warning display. E.C.A.M. illuminates the master caution light and sounds a single chime. The fire detection and protection panel illuminates the loop test lights. The over-temperature and fire circuit indicates a fire warning to the flight crew. The E.C.A.M. responds with an engine fire message and a corrective action procedure on the warning display. The E.C.A.M. also illuminates the master warning light and sounds a continuous chime. The fuel shut-off lever is illuminated on the pedestal and the engine fire pull handle is illuminated on the fire detection and protection panel.


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20.4 FIRE SUPPRESSION Typical fixed systems in the types of aircraft for which fixed fire extinguisher systems are specified, it is usual for the extinguishant to be stored in the containers under pressure and to be discharged by electrically firing cartridge units within the extinguisher discharge heads. The firing circuits are controlled by switches or fire control handles in the flight crew compartment; some types may also be automatic as in the case of an APU. The layout of a system and the number of components required, depend largely on the type of aircraft and number of power plants and also on whether fire protection is required for auxiliary power units, landing gear wheel bays and baggage compartments. A secondary function of the Engine fire handles is to isolate the engine from other aircraft systems to prevent them from making the fire worse, and also to stop the fire spreading. The systems usually affected are fuel, hydraulics, and air systems. See figure 20.9. There are two types of fixed systems used for power plant fire protection.


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Isolation Functions of Engine Fire Handles Figure 20.7 Issue 2 – April 2003

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20.4.1 TYPES OF FIRE SUPPRESSION SYSTEM ONE SHOT SYSTEM

In this system the extinguishant bottle has only one outlet from the neck and is connected to one engine only. If the operation of that cylinder fails to suppress the fire, nothing can be done unless another bottle is fitted as a back up.

A Single Shot Fire Extinguisher System. Figure 20.8.


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20.4.2 TWO SHOT SYSTEM (SHARED EXTINGUISHERS)

The extinguishant cylinder in a two shot system has two outlets from the neck and each outlet supplies extinguishant to a different engine. Each outlet is operated independently by a suitably marked firing button situated in the cockpit. When the “first shot” button is pressed, the relative extinguisher will discharge its contents via a Directional Flow valve to the required fire zone.


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A Two Shot Fire Extinguisher System. Figure 20.9.


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20.4.3 TWO SHOT SYSTEM (SINGLE HEAD EXTINGUISHERS)

In this type of system, there are two separate extinguisher bottles for each engine, each having a single outlet, to the same engine. The system operates in the same way as the two shot system.

A Two Shot System Using Single Head Bottles.

Figure 20.10


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A Two Shot System With Single Head Bottles. Figure 20.11. Issue 2 – April 2003

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20.5 EXTINGUISHERS Extinguishers vary in construction but are normally comprised of two main components: the steel or copper container and the discharge or operating head. CARTRIDGE

A Typical Two Head Fire Extinguisher Bottle. Figure 20.11. Issue 2 – April 2003

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20.5.1 OPERATING HEAD

A pressure gauge or operated indicator, discharge plug and safety discharge connection are provided for each container. The discharge plug is sealed with a breakable disk combined with an explosive charge that is electrically detonated to discharge the contents of the bottle.

A Twin Head Extinguisher. Figure 20.13. 20.5.2 SAFETY DISCHARGE

The safety discharge connection is capped at the inboard side of the structure with a red indicating disk. If the temperature rises beyond a predetermined safe value, the disk will rupture, dumping the agent overboard. A mechanical indicator is fitted to the outlet of the overboard vent.(Fig 20.14.) A pipe is connected between the indicator and the pressure relief outlet on the extinguisher. When discharge occurs, the extinguishant flows along the pipe and blows out the sealing plug and nylon disc revealing the bright red interior of the bowl. The sealing plug prevents the ingress of moisture that could corrode the rupture disk and cause premature leakage.


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Fire Bottle Discharge Indicator. Figure 20.14.

20.5.3 DISCHARGE TUBE CONFIGURATION

Very dependent upon the type and size of engine installation, typical system shown in figure 20.15.

Typical Discharge Tube Installation. Figure 20.15.


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20.5.4 OPERATING TIME

In most systems the extinguishant will dissipate in a few seconds. More recently a system has been developed which will discharge in 1 to 2 seconds. This system is known as HRD (high rate of discharge). 20.5.5 EXTINGUISHANT

Older aircraft use Methyl Bromide as the extinguishing agent, this has been replaced by BCF (Bromochlorodifluoromethane) Halon 1301. Both of these chemicals are CFC’s and are banned under the Montreal Protocol. A recent amendment to this document has allowed their continued use in aircraft until a suitable alternative is found or existing stocks run out. CO2 is sometimes used however it does form snow when released which can cause hot metal components to explode so its use is limited.


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20.6 INDICATIONS OF FIRE DETECTION When the fire detection system is exposed to an overheat condition or fire, the detector warning lights in the cockpit illuminate and the fire warning bell sounds. The warning light may be located in the fire-pull handle on the instrument panel, light shield, overhead panel or fire control panel. 20.6.1 FIRE T HANDLE

These fire switches are sometimes referred to as fire-pull T-handles. In some models of this fire-pull switch, pulling the T-handle exposes a previously inaccessible extinguishing agent switch and also actuates micro-switches that energise the emergency fuel shut-off valve and other pertinent shut-off valves.

Fire ‘T’ Handle. Figure 20.16.

20.6.2 FIRE BELL

An alarm bell control permits any one of the engine fire detection circuits to energise the common alarm bell. After the alarm bell sounds, it can be silenced by activating the audio cut-out switch or pressing either of the red alert flashers. The bell can still respond to a fire signal from any of the other circuits.


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Isolation Functions of Engine Fire Handles Figure 20.17 Issue 2 – April 2003

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20.6.3 FIRE DETECTION TEST

Most fire protection systems for turbine engine aircraft also include a test switches and circuitry that permit the entire detection system to be tested.

Fire Test. Figure 20.18.


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ENGINES

20.7 DISCHARGE INDICATORS In fire extinguisher systems of the fixed type, provision is made for positive indication of extinguisher discharge as a result of either (a) intentional firing, or (b) inadvertent loss of contents, ie. pressure relief overboard or leakage. The methods adopted are: 

Mechanical in operation.



Electrical in operation.

These devices are known as ‘bottle gone indicators’. 20.7.1 MECHANICAL INDICATORS

Mechanical indicators are, in many instances, fitted in the operating heads of extinguishers and take the form of a pin, which under normal conditions is flush with the cap of the operating head. When an extinguisher has been fired and after the charge plug has been forced down the operating head, the spigot of the plug strikes the indicator pin causing it to protrude from the cap.

Mechanical Bottle Fired Indicator. Figure 20.19.

20.7.2 ELECTRICAL INDICATORS

Electrical indicators are used in several types of aircraft and consist of indicating fuse indicators, magnetic indicators and warning lights. These are connected in the electrical circuits of each extinguisher so that when the circuits are energised, they provide indication that the appropriate cartridge units have been fired. In some aircraft, pressure switches are mounted on the extinguishers and are connected to indicator lights, which come on when the extinguisher pressure reduces to a predetermined value. Pressure switches may also be connected in the discharge lines to indicate actual discharge as opposed to discharge initiation at the extinguishers. Issue 2 – April 2003

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A fuse indicator has a pellet of coloured wax around heating element, when electrical power is applied to the element the wax vaporises and spreads itself all over the clear plastic indicator dome. 20.8 CARTRIDGES OR SQUIBS These devices are the electrical detonators that ‘fire’ the bottles. These detonators are explosive devices and special precautions apply when handling and transporting them. Prior to fit a ‘No Volts Test’ must be carried out to the fire system wiring to ensure that it will not go off when connected. When handling the cartridges do not touch the pins as a static discharge could fire it, ensure that you are earthed and are not wearing clothing that is generating large amounts of static. They should be transported and stored in steel boxes and in a secure manner. On some aircraft a ‘squib’ test is provided, when pressed provides a circuit through the cartridge with a current flow low enough to prevent firing the squib, but sufficient to illuminate a green light if the squib is serviceable. Do not press the fire button to do this test! 20.8.1 LIFE CONTROL OF SQUIBS

The service life of fire extinguisher discharge cartridges is calculated from the manufacturer’s date stamp, which is usually placed on the face of the cartridge. The cartridge service life recommended by the manufacturer is usually in terms of hours below a predetermined temperature limit. Cartridges are available with a service life of approximately 5,000 hours.


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Cartridge (Squib) Test. Figure 20.20


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21 ENGINE MONITORING AND GROUND OPERATIONS. 21.1 PROCEDURES FOR STARTING AND GROUND RUNNING. Before starting and ground running any gas turbine powered aircraft, several considerations must be taken into account. The first and most important is the safety of personnel, aircraft and equipment involved in the run. Secondly the safety of personnel, aircraft, equipment and buildings close to but not involved in the run, and thirdly the safety of the engine itself. The aircraft maintenance manual (AMM) will show the danger areas associated with the aircraft (fig 21.1 & 21.2) and these must be observed at all times. To alert other personnel of the need to take precautions safety signs should be positioned. Ideally engine ground runs should be carried out in a designated area which will to a large extent assist safety. There may however be occasions when these designated areas are not available. Precautions to protect personnel, aircraft and equipment must still be observed.

Diagram of Fokker 100 Aircraft showing the Engine running danger areas at idle and full power. Figure 21.1.


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The BAe 146 Danger Areas Showing Entry Corridors. Figure 21.2.


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Particular attention should be paid to the positioning of the aircraft and its ground support equipment (GSE). The aircraft should be facing into wind and securely chocked (possibly with the front and rear chocks tied together). The visual and free movement of both compressor and turbine should be checked, and the engine air intake examined for loose articles. The areas to the front and rear of the aircraft should be checked for loose articles and spilt fuel, which could cause a hazard to the aircraft during the run. The technical log must be checked to ensure that no outstanding entries will jeopardise the operation or function of other aircraft systems. Other entries may require functional checks to be carried during the ground run, which may also require involvement in the run of other tradesmen. Ground support equipment should be positioned to ensure their safe operation and movement, if required, during the start and run. 21.2 STARTING

Commonly Used Hand Signals for Ground Running. Figure 21.3. Prior to starting the engines all personnel involved must be made aware of their responsibilities and role during the run. If hand signals are to be used (fig. 21.3.) they should be agreed and understood by all concerned. All personnel outside the aircraft must wear ear-defenders, if possible one or more of the external team should have an intercom headset for direct communication with those inside.


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The person(s) operating the controls during starting and running must be familiar with the controls, instruments and limitations associated with the engines. In particular they should be aware of the limitations imposed upon the engines turbine temperature during start. If the start is to be made from the aircraft batteries, ensure they are fully charged. If a ground power unit is to be used, it must be appropriate for the aircraft and must be correctly connected. If the starter requires air, then the APU will be required or a suitable air-cart attached correctly to the aircraft.

Events in a Typical Gas Turbine Engine Start. Figure 21.4. Issue 2 – April 2003

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Starting procedures will vary depending on aircraft type and installation hence, the AMM must always be referred to. The example that follows (Fig. 21.4. refers) is however typical and will serve as a general guide: 1. Set all controls and switches etc. as per AMM. 2. Switch ‘on’ electrical power. 3. Carry out relevant flightdeck safety checks i.e. Brakes ‘on’, Engine fire warning tests etc. 4. Low pressure fuel valve (LP) [sometimes called the LP cock] check ‘open’. 5. Contact Air Traffic Control on the radio, giving location, type of run and number of people on board. 6. Switch ‘on the aircraft booster pumps. 7. Confirm ‘clear to start’ from safety man. 8. Select start master switch to ‘on’, the aircraft systems will be put into starting mode. 9. Select ‘start’ At this point the starting sequence becomes semi automatic. 10. The starter begins to rotate the compressor (HP if multi shaft) to provide a flow of air through the engine. 11. The engine ignitors are energised. Observing the engine’s RPM, when this reaches a speed of approximately 10 – 20%, advance the high pressure fuel valve to open either by moving the throttle or the HP cock lever (on aircraft with a separate lever) to the ‘fuel on’ or ground idle (GI) position. The engine speed will increase as the starter motor continues its acceleration; fuel will be supplied to the atomisers and will be burnt in the combustion chambers. ‘Light up’ will occur which will be indicated by a rapid rise in Exhaust Gas Temperature (EGT). 12. The rise in gas temperature will cause the air within the combustion chamber to expand which when passed through the turbine will assist the acceleration. 13. During this phase the oil pressure should start to rise. 14. As the engine accelerates it will reach a point called ‘the self-sustaining speed’; this is the minimum speed at which the engine can run unassisted. 15. Once above self-sustaining speed the starter and ignition will cut out automatically, and the engine will accelerate to ground idle under the control of the fuel system. It is during this phase of the acceleration when there is a great risk of exceeding the maximum starting temperature of the engine, so vigilance is required to monitor the EGT. Issue 2 – April 2003

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16. The engine should settle quickly at ground idle. At this point the other flight deck indications should be checked to ensure the start was successful, i.e. the starter and ignition should have cut out, oil pressure should be in range (fig. 21.5), check N1,or propeller, or rotor speed.

Oil Pressure Limits at Ground Idle. Figure 21.5.


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17. If only one engine is to be started, the Start Master Switch should be switched ‘Off’ and electric fuel pumps switched ‘on’ to support the running engine. 21.3 UNSATISFACTORY STARTS Unsatisfactory starts can be broadly categorised in the following three areas: 1. Hot starts. These occur when the EGT exceeds the manufacturers specified limits. They normally result from too rich a fuel/air ratio. The engine should be shut down immediately. It is good practice to shut down before the limit is reached if possible to prevent overswing . Improper ratio of fuel/air may be caused by a malfunction in the fuel control unit (FCU), incorrect use of the throttle, or a restriction of the air flow into the intake, i.e. ice, snow, cross wind etc. Manufacturers will list the degrees of overtemperature limits in terms of time and temperature rather than stating a specific overtemperature (fig.21.6.).

Overtemperature Limits During Starting. Figure 21.6.


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2. Hung Start. After light up the engine RPM does not increase to ground idle, but remains at some lower value. The EGT may stabilise or continue to rise (sometimes rapidly). Again EGT must be monitored closely and the engine shut down if limits are exceeded. Hung starts are often caused by insufficient power to the starter motor, or the starter cutting out too soon. It could also be caused by rotational stiffness within the rotating system, which may be caused by the engine or one of its accessories. 3. No Start. The engine does not ‘light up’ as indicated by no increase in RPM or EGT. This could be the result of a faulty starter motor, insufficient power to the starter motor, faulty ignition system or even a problem with the FCU, engine fuel system or possibly the aircraft fuel system. For any of the above, the limitations laid down in the AMM and Company Procedures must be adhered to. 21.4 ENGINE STOPPING. Normal shut down of a gas turbine engine is accomplished simply by closing the throttle (and/or HP cock) to the ‘fuel off’ position. This should be followed by switching ‘off’ the aircraft fuel booster pumps. There are however other factors to consider which will depend upon the operation of the engine prior to shut down. If the engine has been operating at high power for any length of time a three to five minute cooling period at ground idle is usually recommended prior to shut down. The shroud casing and turbine rotors do not cool down at the same rate after shut down. The turbine shroud casing, cooling at a faster rate may shrink onto the still rotating rotor and cause damage. Run down time should be monitored in terms of the time taken to stop, the manufacturers will give a recommended time, also check for unusual noises; compressor rub, turbine rub and accessory drives. Assuming all is well, all controls and switches should be positioned in accordance with the AMM and electrical power selected ‘off’. Remember to inform Air Traffic that the run has been completed.


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21.5 ENGINE FIRES There are two main types of engine fire that can occur when running gas turbine engines and they are dealt with in different ways. 1. Fire in the engine nacelle. This type of fire will usually cause the engine fire warning system to function, although it may be spotted by the safety man outside. The engine should be shut down immediately. The engine fire handle should be pulled to isolate the nacelle. A fire extinguisher should be discharged into the nacelle, preferably the CO 2 extinguisher by the safety man, if not available then one of the aircraft extinguishers. Inform the control tower, then shut down any other running engines, switch off power and evacuate the aircraft. 2. Fire in the core engine or external to the engine nacelle. Fire can occur within the core engine especially after a ‘wet start’ (a start which fails after fuel has been selected on). If insufficient time is allowed for fuel to drain from the engine or there is a fault in the drain system, fuel can pool inside the turbine area. On the next start this fuel ignites and flame and black smoke are seen in the exhaust. This may then be pushed out of the jet pipe by the airflow and spread onto the ground as a burning pool. This type of fire is usually spotted by the safety man. He should inform the engine operator, who should then cut off the fuel by shutting the throttle and or hp cock. The starter motor should continue to run to cool the engine and to push the fire out of the engine. The safety man should attempt to put out the fire by discharging CO2 directly into the intake never up the exhaust (as CO 2 produces ice when discharged which can have an explosive reaction when directed into very hot metal). If the fire has spread out of the jet pipe this fire should also be tackled with the CO2 extinguisher. When the starter reaches its maximum running time select the start master switch to ‘off’ to cancel the start signal, pull the fire handle but do not fire any extinguishers. Make the aircraft safe i.e. shut down running engines and electrical power and evacuate. Beware!! sometimes burning fuel from this type of fire can run down inside the cowlings and cause damage to the engine.


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21.6 INTERPRETATION OF ENGINE POWER OUTPUTS AND PARAMETERS. Engine Ratings. An understanding of gas turbine ratings is necessary in order to be able to interpret graphs published by the manufacturer in the AMM. Gas turbine engines are rated by the number of pounds thrust they are designed to produce for: 

Take-off (T.O.)



Maximum Continuous Thrust (M.C.T.)



Maximum Climb (CLB.)



Maximum Cruise (CRZ)

Parameters Turbojet and turbofan engines can be measured via Engine Pressure Ratio (EPR) or Fan Speed (N1). Turboprop and turboshaft engine power is measured via Torque produced. In the majority of cases the Take-off (T.O.) rating will be a ‘part throttle’ rating. This means that T.O. thrust will be obtained at throttle settings below the full throttle position. The reason for establishing a rating for a particular engine is quite simply to accommodate the various atmospheric conditions under which the engine will be operating. Engine Pressure Ratio (EPR). Figure 21.7. shows the manufacturers published tables which must be used to establish the engine is producing its certified T.O. thrust under varying temperature and altitude conditions.


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Fokker 100 EPR Setting Chart. Figure 21.7.


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Effects of Temperature and Altitude. Cold and/or Low On cold days the air density will increase. If the engine has a ‘part throttle rating’ and the throttle is advanced to its maximum position, the thrust produced will be significantly higher, resulting in the engine exceeding its mechanical and/or thermal limits as set by the manufacturer. Hot and/or High. On hot days, when the air density is less, a significant reduction in thrust will result. Advancing the throttle to its maximum condition could again cause the engine to exceed its thermal and /or mechanical limitations. Density of air will also be affected by altitude, although the temperature of air drop (1.97°C per 1000ft Temperature lapse rate) which should cause an increase in thrust, the density of the air drops at a higher rate due to the drop in pressure so thrust decreases with altitude at a relatively slow rate. When the tropopause is reached at 36,000ft the temperature remains constant and the thrust drops off at a greater rate when climbing. The engine manufacturers graphs and tables will enable the operator to control the engine within safe thermodynamic and mechanical limits. By observing these limits the engine will be protected against unnecessary wear and tear as well as maintaining the recommended time between overhaul periodicity’s.

Effects of Air Temperature and Altitude on Thrust. Figure 21.8.


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Flat Rating and Full Rating. The engine manufacturer will establish the safe maximum T.O. Thrust to be used under ISA conditions i.e. 15°C at sea level, and at other altitudes. This will necessitate the adjustment of the throttle position to protect against exceeding the laid down limits. The terminology, and a brief explanation of terms in common usage is: Flat Rating

The term has the same meaning as ‘part throttle’. It is used in conjunction with the graphs/tables published in the AMM. It will determine the maximum thrust setting that must not be exceeded if operating below ISA conditions. (early engines used 15°C as the max flat rating point, newer engines are rated to 22.5°C or higher). See Figure 21.9.

Full rating

Again used in conjunction with the graphs/tables in the AMM, this will determine the maximum thrust available if operating above ISA. See Figure 21.9.

Charts of the ALF502 (top) and LF507 (lower) engines. The charts show flat rating and full rating, the ALF502 to 15°C (59°F) and the LF507 to 23.3°C (74°F). The lower chart also shows Take-off derating. Both charts show the max continuous thrust for each engine. Figure 21.9.


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N1


settings for Engine Ground Running – Anti –ice OFF. Figure 21.10.


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Engine Operating Limitations for the ALF502. Figure 21.11. Issue 2 – April 2003

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Reserve Take-off Power for PW125 Turboprop Engine. Figure 21.12.


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Take-off (wet)

This is the maximum thrust available certified for engine with thrust augmentation systems (water or water/methanol injection or reheat). The rating is time limited usually to around 5 minutes. And is only used during take-off phase of flight.

Take-off (dry)

This is the maximum thrust certified without thrust augmentation. The rating is time limited usually to around 5 minutes. And is only used during take-off phase of flight.

Maximum Continuous This rating is the maximum thrust certified for continuous Thrust (MCT) use. This rating is used at the pilots discretion, to ensure continued, safe operation of the aircraft. MCT is used as the maximum normal thrust available throughout the majority of the flight, and is used when a rapid climb rate is needed (see Figure 21.9.). Maximum Cruise (MCZ)

This is the maximum power certified for cruising.

Ground Idle (GI)/ Flight These are not rating as such, but throttle positions that Idle (FI) are suitable positions for minimum power operations on the ground or in flight. Ground idle which is usually a fixed stop, provides a core engine RPM which will ensure the driven accessories, electrical, hydraulic and pneumatic, as well as providing a comfortable taxi thrust. This applies to flight idle, but must also include the effects of ram air and altitude as well. On approach the engine must be capable of acceleration from flight idle to full power within a maximum time limit of 5 seconds without surging. The flight idle RPM is set to a value where this requirement can be met. This can seriously affect the airframe design, as there may be too much thrust on the approach, so high drag devices may be needed to keep the approach speed as low as possible. Flight idle is a moveable stop which is usually activated by the aircraft weight sensing system, it may also have more than one position if the air bleed loads affect the acceleration time.


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ALF 502 Power Assurance Check. Figure 21.13.


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Some aircraft are certified with two take-off power ratings. The lower rating is normally selected for take-off, if however an engine fails during take-off the remaining engine(s) are set at the higher rating. By operating the engine at the lower rating for the majority of its life, maintenance costs are reduced. This method is often used with turboshaft engines. A similar system of flexible take-off thrust (T flex) is used with turbofan engines where a de-rate is applied to take-off thrust which takes into account aircraft performance and airfield conditions, again reducing the wear and tear on the engine.

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Engine Trend Monitoring Sheet Filled Out on Each Flight by the Crew. Figure 21.14.


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Intentionally Blank


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21.7 TREND MONITORING. The internal operating condition of a gas turbine engine should be monitored both during flight and on the ground. Flight Monitoring can be done by observing the relevant engine instruments (See Fig. 21.14.): 

EPR

 N1  N2  TGT  Fuel Flow  Oil Pressure/Temperature  Vibration These figures can then be transferred onto a graph that will serve to identify the normal/abnormal trends the engine may be developing. By utilising this method of monitoring the operator will be better able to predict the rate of deterioration in engine performance and to instigate some form of maintenance to correct and reestablish normal performance. The graphic trend charts can of course be produced be produced by a computer, and most modern turbine engined aircraft’s engine performance is automatically recorded during flight. The recorded data is then downloaded and processed and then analysed either manually using charts or automatically by computer. The common term used for this type of monitoring system is ‘Engine Condition Monitoring (ECM). Some airlines use this system to monitor pilot performance when handling engines, as fuel burn and engine life are two major costs, inappropriate operation can lead to further training and/or loss of job! Figure 21.15. shows a trend monitoring graph for an ALF 502 engine using data collected from the forms (fig. 21.14.)


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------ = Reference Baseline (Based on first 10  of new engine)  = NH, MGT,W f = Actual deltas  = Average Deltas (Average of last 10 ) Trend Monitoring Graph. Figure 21.15. Issue 2 – April 2003

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21.7.1 ON GROUND MONITORING

Inspecting and monitoring the engine for deterioration or damage is a vital part of aircraft maintenance. The inspections can be broken down into two main areas, ‘Air washed’ and ‘Oil washed’. Many of the inspection techniques involved are ‘non destructive’ of a component/system in order to determine its serviceability. Techniques in common use include inspection and monitoring via: 

Visual inspection



Boroscope inspection



Magnetic chip detectors (MCD) debris analysis



Oil filter debris analysis



Spectrometric oil analysis programme (SOAP)



Vibration analysis



Noise analysis

21.7.2 AIR WASHED COMPONENTS

Visual inspection There are three basic routine inspections to which gas turbine engines are subjected: 

Pre flight inspection



Cold section inspection



Hot section inspection (HSI)

Pre Flight inspection Typical routine inspection before flight will include: 

External inspection of engine cowlings



Inspection of intake, IGV’s, Fan blades and First stage compressor for signs of damage.



Inspection of exhaust unit, rear turbine stage and thrust reversers (if fitted) for signs of damage, cracks, and discoloration etc.



Inspect inside and out of the cowlings for fuel, oil and air leakage from the engine and its accessories.

Oil level checks are carried out with defined times after shut down and form part of the daily inspection which also includes a more detailed inspection covering the pre flight inspection areas. Issue 2 – April 2003

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Cold Section Inspection. The ‘cold section’ of a gas turbine will consist of the compressor and diffuser, IGVs and fan if fitted. Any interference with the flow of air will seriously effect the performance of the engine. Inspection (often using a boroscope) will include: 

Inspection for damage



Inspection for cleanliness.

Hot Section Inspection (HSI) The ‘Hot Section’ of a gas turbine engine is the ‘workhorse’ section of the engine. It provides the power to drive the turbines, which in turn drive, the compressor, the fan, or the propeller to, and in its own right produce thrust. Inspections are normally carried out with the engine ‘in situ’ i.e. on the wing/fuselage and consist mainly of: 

A review of engine performance just before the inspection, noting any indications/ history of hot starts, hung starts, overtemperatures, overspeeds, oil pressure/temperature fluctuations, vibration figures etc.



Inspection of fuel nozzles, combustion chambers, ignitors, exhaust unit etc for signs of damage, cracks, leaks discoloration and burning etc.



Inspection of turbine blades for signs of damage, excessive creep, discoloration etc.



Inspect for buckling, twisting and damage to the jet pipe and reversers, incuding the correct functioning of moving parts.

Boroscope Inspection Boroscope inspections involve looking at components within an engine using an optical probe. The probes are inserted in to the engine through ports in the engine casings, and can be rigid or flexible, the choice being dependant on the difficulty at obtaining a satisfactory view of the required features. Some of these inspection ports are the attachment points of other functional devices that intrude into the engine (e.g. ignitor plugs or temperature probes) but on more modern engines there are usually several purpose made ports for boroscope inspections.

A Rigid Boroscope. Figure 21.16.


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In general the boroscope inspection technique saves many hours of work and can reduce the down time of the aircraft in many cases, disassembly and reassemble of the engine not being required. The boroscope is essentially an eyepiece connected to a rigid or flexible tube. The tube contains fibre optic cables that carry light and therefore visual images, even when the tube is made to bend through considerable angles. A second fibre optic cable within the tube carries light from a bright light source to illuminate the target. At the end of the tube there will be a viewing lens, with a light source lens nearby. Most flexible probes have a steerable tip which allows the operator to steer toward the target, and the lens is mounted in the tip to view straight ahead. Rigid probes may have prisms behind the lens to allow the probe to view at right angles or 45° to the probe.

A Rigid and a Flexible Boroscope With Their Accessories. Figure 21.17.


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The operator inserts the probe into the appropriate port to view the internal components. Some techniques require the use of guide tubes to ensure that a steerable probe is going in the right direction. Ports are usually designed into the compressor, turbine and combustion sections of the engine. On the viewing end of the boroscope there will be the controls for the steerable tip (flexible probe) and to allow the operator to focus the probe. It is more usual these days to find a video camera attached to the eyepiece so that a recording of the inspection can be made. The video is presented on a television screen that allows a much bigger picture and also more than one person to view the screen. The recording is useful as sometimes it is very difficult to find or reproduce a view that may fleetingly pass and which gives you concern, also should a problem be observed it can be dispatched to the manufacturer for analysis by their experts. When turning the engine careful counting of the blades or number of turns of the hand turning point is required to ensure that all of the blades have been viewed.

The Boroscope Ports on an ALF 502 Engine. Figure 21.18.


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Boroscope probes are very delicate and expensive pieces of equipment and great care is needed when using them. It is very easy to damage a probe if it is inserted between rotor and stator blades, even to the point of cutting the end off the probe! If this is the technique you are using you may need to lock the rotor to prevent the risk of damage. If the technique requires the engine to be rotated, i.e. to check the turbine blades, then a port and probe which does not go through the blades is required. Remember when outside very little wind can cause the rotor to move! Interpretation of boroscope images is not always as easy as it might sound. The viewer is very small which can make tiny cracks look like the Grand Canyon! Equally relatively small distances can appear distant when viewed. These make it difficult when assessing a component which is close to a limit, and may require you to look at a similar object with the naked eye to make a proportional judgement. Most companies require special approval for people to carry out boroscoping.

Fourth Nozzle Inspection. Figure 21.19.


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Third Nozzle Inspection Figure 21.20.

Fuel Injector Inspection. Figure 21.21. Issue 2 – April 2003

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Combustion Liner Inspection. Figure21.22.

First Nozzle Inspection. Figure 21.23. Issue 2 – April 2003

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NH Compressor Inspection. Figure 21.24.


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21.7.3 OIL WASHED COMPONENTS

Lubrication Systems With oil washed components, any mechanical wear from contacting surfaces, gears, bearings etc. will produce debris which will be carried within the oil circulating round the engine. Analysis of this debris can provide a very useful method of assessing any trends in wear from the internal engine components. Analysis can involve a number of different methods. Magnetic Detector Plug Debris Analysis The magnetic chip detectors (MCDs), are small, permanent magnets installed in the scavenge/return lines of the engine oil system. They will attract ferrous debris from the oil. At specified intervals they are removed and visually inspected. As a general rule, the presence of small, fuzzy particles or grey metallic paste is considered satisfactory and the result of normal wear. Metallic chips or flakes however are an indication of a more serious nature requiring more in depth investigation. Some organisations have specialised departments that, by examining debris under a microscope can, by virtue of shape, size, colour and marks determine quite accurately where the debris is from; ball bearing, roller bearing, gear teeth etc. They may also utilise a ‘Debris Tester’ which will provide a means of measuring (magnetically) the mass of the debris produced. The figure gleaned can then be transferred to a graph which will indicate the normal /abnormal amounts of debris the engine is generating. A sudden increase in the amount of debris observed either visually or by graphs generated from debris tester figures may result in more frequent inspections of MCDs, or , in extreme cases, engine removal for subsequent strip examination. An indicating type of chip detector may be used to give a warning in the flight deck if and when excessive debris is present. Basically the detector has two probes which if connected by the debris act as a switch to bring on a warning. A much newer type of chip detector is the electric pulsed chip detector, which can discriminate between wear debris particles considered non-failure related, and large wear debris particles, which could be an indication of a more serious nature. Operating in a similar way to the indicating type chip detector, if the warning light illuminates, an electrical charge can be instigated either manually or automatically across the gap. Small wear debris particles will be ‘burnt’ off and the light will extinguish. Large wear debris particles will however not burn off and the warning light will remain ‘on’.


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(A) In line type scavenge magnetic oil chip detector (non-indicating). (B) Chip accumulation of ferrous particles. (C) Comparison between standard, pulsed and detector showing detector auto indicating chip Magnetic Chip Detectors. s. Figure 21.25.


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Oil Filter Debris Analysis Oil filters serve an important function within the lubrication system of a gas turbine engine in that they remove foreign particles that collect in the oil system. Filters are removed at regular intervals for cleaning, any particles present can then be analysed visually. If visual inspection reveals evidence of excessive debris this can be more accurately analysed via ‘spectrometric analysis’. Spectrometric Oil Analysis Programme (SOAP) Under certain conditions and within certain limitations, the internal condition of any mechanical system can be evaluated by the spectrometric analysis of the lubricating oil. The components of mechanical systems contain aluminium, iron, chromium, silver, copper, tin magnesium, lead and nickel as the predominant alloying elements. The moving contact between metallic components will, despite lubrication create wear, the debris resulting from this wear being carried away by the lubricating oil. If the rate of wear of each kind of metal can be measured and be established as normal or abnormal, the rate of wear of the contacting surfaces will also be established as normal or abnormal. At specified intervals samples of oil are removed from the engine for analysis. Spectrometric analysis is possible because metallic ions emit characteristic light spectra when vaporised in an electric arc or spark. The spectrum produced by each metal is unique to that particular metal and, the intensity of the light can be used to measure the quantity of metal in the sample Again, information gained could be transferred onto a graph to show evidence of normal/abnormal trends. In this process the oil is burnt which will also show on the analysis, but is ignored as a known substance. If we suspect that some or all of our fleet may have been contaminated by an incorrect oil, it is possible to sample the fleet using spectrometric analysis, to determine which components have the wrong oil in.

Oil Spectrometer. Figure 21.26. Issue 2 – April 2003

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21.7.4 INSPECTIONS

Maintenance covers both the work that is required to maintain the engine and its systems in an airworthy condition while installed in the aircraft, and the work required to return the engine to an airworthy condition after removal from the aircraft for overhaul. Comprehensive instructions covering the actual work to be done are contained in the relevant sections of the aircraft’s maintenance manual (AMM) for installed engines, frequently referred to as ‘on wing maintenance and the component maintenance manual (CMM) for uninstalled engines. Both sources of maintenance information are based on the manufacturers recommendations, which in turn are approved by the appropriate airworthiness authority. The maximum time an engine can remain ‘on wing’ is limited to a fixed period agreed between the engine manufacturer and the airworthiness authority. This period is often referred to as the Time Between Overhaul period (TBO) and on reaching this limit the engine must be removed for overhaul. Because the TBO is actually determined by the life of a few major more critical assemblies within the engine this means that other assemblies can continue in service for much longer periods based on an ‘on condition’ monitoring process. Basically this means that a ‘life’ is not declared for a total engine, but only for the more critical assemblies. Less critical assemblies on reaching their ‘life’ limit are replaced ‘on wing’ or are inspected to ascertain that they are in a condition, which will allow them to continue in service. It is the ‘on condition ‘ items which concern the aircraft maintenance engineer (AME) being the checks, inspections, and examinations that are required on wing. On wing maintenance falls into two categories, scheduled maintenance and unscheduled maintenance. Scheduled Maintenance Checks. These embrace the periodic and recurring checks that have to be carried out in accordance with the maintenance schedule and an example is shown in figure 21.27 Unscheduled Maintenance Checks. These cover work not normally related to scheduled maintenance or time limits. Bird strikes, lightning strikes, heavy landings will result in unscheduled checks being carried out. Defects, trouble shooting and even manufacturers specific requirements regarding repair, and adjustments etc. will also require unscheduled maintenance.


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Section of Maintenance Programme for BAe 146 for Oil System Components. Figure 21.27. Issue 2 – April 2003

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AME’s will invariably find that for most inspections the engine is divided into two main sections, the cold section (compressor, diffuser, fan, IGVs etc.) and the hot section (combustion chambers, burners, turbines, NGVs, exhaust unit, etc.). Cold Section Inspections. Damage to fan blades, IGVs and compressor blades can cause engine failure and possible loss of the aircraft. Much of the damage to this section of the engine is brought about by the ingestion of Foreign Objects into the intakes, hence the term Foreign Object Damage (FOD). The quality of air close to ground level or sea level leaves a lot to be desired. It is filled with tiny particles of dirt, soot, sand salt, oil and other foreign matter. The large volume of air being drawn inwards, then centrifuged outwards can result in a coating forming on the compressor casing and stators as well as the fan and rotors. This accumulation of dirt reduces the aerodynamic efficiency of the compressor resulting in a deterioration of engine efficiency. Repeated ingestion can also result in erosion of the compressor blades. It can even cause erosion and damage to the hot section assemblies, NGVs, turbine blades, etc. If inspection reveals an accumulation of dirt on the compressor it must be cleared. Some maintenance schedules will schedule regular periodicity’s for cleaning. An example of this is shown in Figure 21.28. Operating Environment

Nature of Wash

Recommended Frequency

Recommended Method

Remarks

Continuously salt laden

Desalination

Daily

Motoring

Strongly recommended after last flight of day

Occassionally salt laden

Desalination

Weekly

Motoring

Strongly recommended. Adjust washing frequency to suit condition.

All

Performance Recovery

100 to 200 hours

Motoring or Running

Strongly recommended. Performance recovery required less frequently. Adjust washing frequency to suit engine operating conditions as indicated by engine condition monitoring system. Motoring wash for light soil and multiple motoring or running wash for heavy soil is recommended.

An Example of a Typical Wash Schedule. Figure 21.28.


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Two Methods of combating the effect of dirty compressors are in use. The fluid cleaning process and the abrasive grit cleaning process. Fluid Cleaning. This procedure involves spraying an emulsive type surface cleaning fluid into the compressor whilst the engine is turning either on the starter motor or at low RPM. This is followed by a rinsing solution being applied. This process would be used to restore engine performance as is commonly referred to as a ‘performance recovery wash’. To remove salt deposits a water wash only may be required. This process is termed a ‘de-salination wash’. A schematic view of equipment that might be used is shown in figure 21.29.

Fluid Cleaning. Figure 21.29 Issue 2 – April 2003

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Abrasive Grit. This method of compressor cleaning involves injecting an abrasive grit into the engine at selected power settings ( Figure 21.30.)grit used may be ground walnut shell or apricot pits. The type and amount of material and the operational procedures will be described in the AMM. The main advantage of this procedure is that allows the time between cleaning to be extended because it produces a better result. However because the grit is mostly burned up in the combustion zone of the engine, it will not give an effective cleaning of the turbine blades and vanes as the fluid.

Abrasive Grit Compressor Cleaning. Figure 21.30.

Compressor Damage. Foreign objects often enter engine air intakes either accidentally or through carelessness. Items such as pens, pencils cigarette lighters etc. can be drawn out of pockets and ingested by the engine. The compressor could be damaged beyond repair. Likewise, tools left in engine intakes could be drawn in causing damage. Prior to starting an engine therefore, the AME should ensure that all tools used in the vicinity of the intakes are free of any foreign objects and the area in front of intakes should be cleared of any loose stones or rubbish . Examples of the typical types of damage to be found on compressor blades is shown in Figure 21.31. and possible causes of damage and the terminology used in Figure 21.32.


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Compressor Blade Damage. Figure 21.31.

Blade Maintenance Terms. Figure 21.32. Issue 2 – April 2003

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Damage Limits and Repair. Minor damage to compressor and fan blades may be repaired provided the damage is within the allowable limits established by the manufacturer in the AMM. Typical limits for fan blades are shown in Figure 21.33.. All repairs must be well blended so that the finished surfaces are smooth.

Typical Fan Blade Damage Limits. Figure 21.33.


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ENGINES

The majority of cold section inspections will require the use of a strong light source and sometimes a small mirror. If however doubt exists as regard the extent of damage, then a boroscope inspection would be instigated. Always observe the safety precautions associated with working in the intake. Ensure that the flightdeck is suitably placarded informing other personnel that you are in the intake. Tripping of C/Bs may be required by the manufacturer in order to isolate the starting and ignition circuits. A safety man may be required who’s job it will be to look after your interest. Don’t get sucked in!!! Hot Section Inspections (HSI’s) The hot section includes all components in the combustion and turbine sections of the engine. Scheduled inspections may involve visual inspection of hot section components, and limited dimensional checks and fits and clearances as called up in the maintenance schedule and described in the AMM. The term ‘hot section inspection’ is usually interpreted to indicate a time related inspection of the hot section components. It may also be required following an over-temperature condition or hot start. Some more in depth HSI’s will require the removal of major components of the hot section. The modular construction of most modern gas turbine engine (Figure 21.34) will enable this removal element of the task to be carried out on the wing, thus reducing the down time. To reduce this down time figure even more, some operators maintain a stock of ‘hot section’ modules that are ready for immediate replacement, the removed item being returned for inspection to the operators overhaul facility.

ALF 502 Modules. Figure 21.34. Issue 2 – April 2003

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Disassembly of Hot Section. The disassembly/reassembly process must ensure that component parts are reassembled in the same position they came apart from. This will require marking of components. A note of caution here. When marking any hot section component do not use a marker that will leave a carbon deposit. Hot metal will absorb carbon which can lead to intergranular stress and failure of the component. Line Inspection of Combustor Turbine Section. On wing inspection of the combustor turbine section can be done visually through the jet pipe using a strong light source and a mirror and if required a magnifying glass. Boroscope inspection is also used as is, on occasion, non destructive methods of inspection such as dye-penetrant. As in other hot section inspections, the AME is most likely to see small cracks caused by compression and tension loads during heating and cooling. Other than on turbine blades and discs this type of distress is normally acceptable because after initial cracks relieve the stress, no elongation of crack normally occurs. Erosion of blades and NGVs is also quite common, this brought about as a result of the wearing away of metal due to either the gas flow or impurities within the gas flow. Combustion Section. One of the most common faults found in the combustor section of a gas turbine engine is cracks. The combustion liner is made of a high temperature resistant steel that is subjected tom high concentrations of heat. The most common methods of checking for faults is by boroscope (Figure 21.35). With this tool the AME can easily view the internal combustion liner and fuel nozzles, and determine their airworthiness. During the inspection the AME is looking for signs of cracking, warping, burning, erosion and hot spots which may have developed possibly as a result of burner misalignment. What is observed is then compared with the manufacturers limitations.

Combustion Liner Inspection. Figure21.35.


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CFM 56 Combustion Liner Figure 21.36.


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Turbine Discs and Blades. The inspection for cracks is of the utmost importance, most inspections are visual, the dye penetrant method of inspection being too impractical. Cracks on discs however small will necessitate removal of the module or engine for overhaul. Blade cracking also will invariably require removal of the module or engine. Some manufacturers limitation allowance will permit repairs to be effected to damaged turbine blades. Figure 21.37. refers. Cracks however are not acceptable and will require blade replacement. In extreme cases part or whole blades may be missing due to severe overheating causing the blade to melt, on some engines this does not always show up on the vibration indicating system.


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Typical Turbine Blade Damage Limits. Figure 21.37. Issue 2 – April 2003

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Turbine Blade Clearance. Creep is term used to describe the continuous and permanent stretching of turbine blades due to high temperatures and centrifugal forces acting on the blades. Each time a turbine is heated, rotated then stopped (referred to as an engine cycle) each blade will be slightly longer. At regular interval, specified intervals the AME will carry out a turbine tip clearance check (Figure 21.38.). The AMM will stipulate what limitations must be observed and if these are exceeded then the engine or module will require replacing.

Turbine Tip Clearance Check. Figure 21.38.


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Turbine Blade Replacement. Some engine manufacturers will allow replacement of damaged turbine blades by an operators overhaul department. Blade replacement is generally accomplished by installing a new blade of equal moment weight. If the blade moment weight cannot be matched then the damaged blade ,and the blade 180° out may be replaced with blades of equal moment weight or the damaged blade and the blades 120° from it may be replaced with blades of equal moment weight. Code letters representing the moment weight are stamped onto the blade to enable correct balancing of the turbine assembly undergoing blade replacement. Figure 21.39 refers.

Typical Turbine Blade Moment Weight Coding and Change Methods. Figure 21.39.


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Nozzle Guide Vane Inspection. Inspection of the NGVs is possible using a strong light source and mirror, it is more probable however that a boroscope inspection will be required. The NGVs are examined for signs of damage and or bowing on their trailing edges. Bowing may be an indication of a faulty fuel nozzle. Again the engine manufacturer will detail the damage/bowing tolerances which, if exceeded will result in module or engine replacement (Figure 21.41.). Inspection of the exhaust section of the engine can be done visually using an appropriate light source. The exhaust cone and jet pipe are examined for signs of cracking, weeping, buckling or hot spots. Hot spots identified on the exhaust cone may be the result of a defective fuel nozzle or combustion chamber resulting in the requirement for further investigation.

First Nozzle Inspection. Figure 21.40.


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Nozzle Guide Vane Inspection. Figure 21.41. Issue 2 – April 2003

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Exhaust Section Inspection. Inspection of the exhaust section of the engine can be done visually using an appropriate light source. The exhaust cone and jet pipe are examined for signs of cracking, warping, buckling or hot spots. Hot spots identified on the exhaust cone may be the result of a defective fuel nozzle or combustion chamber resulting in the requirement for further investigation.

An Exhaust System. Figure 21.42.


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Vibration Analysis Gas turbine engines have extremely low levels of vibration compared to piston engines. Changes in vibration levels could occur therefore without being noticed. To assist the operator in identifying increasing vibration level, most engines are fitted with vibration indicators that continually monitor the vibration level of the engine. The indication is normally a milliammeter that receives its signals from an engine mounted transmitter via an amplifier. Analysis of engine vibration signals is an important tool for the detection of early failure in mechanical components. Engine Vibration Monitoring (EVM) System.

Vibration Transducer Schematic Figure 21.43.

This may take the form of a solid state circuit device utilising the piezoelectric effect. The device consists of quartz discs with a metallic pattern deposited on them and, arranged such that they serve as a flexible diaphragm. When subjected to pressure changes the resultant flexing sets up an electrical polarisation in the discs, so that electrical charges are produced relative to the amount of flexing. The electrical charges are routed, via an amplifier to the flightdeck indicator. This is calibrated in inches per second (IPS). On some engines there will be more than one sensor, enabling switching if one fails. Yet another useful variation is the wide and narrow band which means the readings can be either taken from over the whole range of vibrations from the engine or by one or two major rotational assemblies such as N 1 and N2 spools. An example of this type is fitted to the RR Tay engine.


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Engine Vibration Indication The engine vibration monitoring (EVM) system shows the out of balance force for the N1 and N2 shaft. High engine vibration shows engine damage or other deviations in the engine. Vibration also reduces the comfort level in the aft passenger compartment. Engine Vibration Monitoring System The EVM system shows vibration in inches/second (IPS) An amber limit shows the maximum vibration level. The EVM system has: 

a dual engine vibration transducer on each engine



an engine vibration signal conditioner



a pushswitch on the overhead panel.

The vibration transducer has two internal vibration pick-ups, a pick-up A and B. each pick-up gives a voltage proportional to the acceleration or deceleration of the vibration. The vibration transducer is on the IP compressor casing. This casing is the housing for bearings of the HP and LP shaft. The engine vibration signal conditioner is a single unit for both engines. It processes the output of the engine vibration transducers for indication. The engine vibration signal conditioner gives two modes of vibration indication, tracked and broadband. Tracked Indication The tracked mode shows vibration of the N 1 and N2 shaft. The engine vibration signal conditioner tunes two filters with an input of the N 1 and N2 RPM indicator generators. Both filters connect to one pick-up of the vibration transducer, the other is standby. The VIB pushswitch on the ENGINE panel controls the active pick-up of the vibration transducer. Broadband Indication The broadband mode is an alternative mode. Vibration of the total power plan t is shown. In this mode the output of both pick-ups in the vibration transducer goes through broadband filters. A semi-guarded switch selects the tracked or broadband mode.


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Fokker 100 Vibration Indication System Figure 21.44.


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22 ENGINE STORAGE AND PRESERVATION. 22.1 STORAGE AND TRANSIT All turbine engines which are to be either stored or shipped for overhaul should be packed in such a way as to prevent damage from corrosion or rough handling. The procedure to be followed is outlined below and should be observed irrespective of the condition of the engine. 22.1.1 FUEL SYSTEM INHIBITING.

The fuel used in turbine engines usually contains a small quantity of water which, if left in the system, could cause corrosion. All the fuel should therefore be removed and replaced with an approved inhibiting oil by one of the following methods: Motoring Method. This should be used on all installed engines where it is convenient to turn the engine using the normal starting system. A header tank is used to supply inhibiting oil through a suitable pipe to the engine. A filter and an on/off cock are incorporated in the supply pipe, which should be connected to the low pressure inlet to the engine fuel system and the aircraft LP cock closed. After draining the engine fuel filter a motoring run should be carried out bleeding the high pressure pump and fuel control unit, and operating the HP cock several times while the engine is turning. Neat inhibiting oil will eventually be discharged through the fuel system and combustion chamber drains. When the motoring run is complete the bleeds should be locked, the oil supply pipe disconnected and all apertures sealed or blanked off. Pressure Rig Method. This may be used on an engine which is installed either in the aircraft or in an engine stand. A special rig is used which circulates inhibiting oil through the engine fuel system at high pressure. The fuel filter should be drained and, where appropriate, the aircraft LP cock closed. The inlet and outlet pipes from the rig should be connected to the high pressure fuel pump pressure tapping and the system low pressure inlet respectively, and the rig pump turned on. While oil is flowing through the system the components should be bled and the HP cock operated several times. When neat inhibiting oil flows from the combustion chamber drains the rig should be switched off and disconnected, the bleed valves locked and all apertures sealed or blanked off. Gravity Method. This is used when the engine cannot be turned. A header tank similar to the one used in the motoring method is required but in this case the feed pipe is provided with the fittings necessary for connection at several positions in the engine fuel system. The fuel filter should first be drained then the oil supply pipe connected to each of the following positions in turn, inhibiting oil being allowed to flow through the adjacent pipes and components until all fuel is expelled: (a)

High pressure fuel pump pressure tapping.


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(b)

Fuel control unit pressure tapping.

(c)

Burner Manifold.

(d)

Low pressure inlet pipe.

Components should be bled at the appropriate time and the HP cock operated several times when inhibiting the fuel control unit. All bleeds and apertures should be secured when the system is full of inhibiting oil. 22.1.2 PACKING.

The engine should be securely attached to its transportation stand, all blanks fitted and apertures taped over to prevent the ingress of moisture. A compartment is usually provided on the stand for the documents relating to the engine, and any other information considered relevant should also be included. If the engine has been removed because of suspected internal failure, any metal found in the filters, broken blades or other evidence should also be packed for examination during overhaul. Engines are wrapped in a hermetically sealed moisture-proof bag which should be examined before covering the engine. Any large tears or holes should be repaired using the repair kit contained within the bag but small cuts may be repaired with adhesive PVC tape. Sponginess of the bag material is caused by contamination with oil or fluid and may sometimes be eliminated by washing with water. If the area remains tacky after washing the bag should be rejected. Some engines or components are packed into rigid containers of wood or metal these will have a mounting frame within them. Wooden containers will require the engine to be sealed in a moisture proof bag within the container however, metal containers are usually sealed and pressurised to approx. 5 PSI and do not require a bag. Bags containing silica gel desiccant should be placed in the air intake and exhaust unit and attached at convenient positions around the engine. Approximately 14 to 18 kg (30 to 40 lb) of desiccant will be required depending on the size of the engine and the manufacturer may specify the use of VPI paper in addition (see Leaflet BL/1-7). A humidity indicator should then be placed in the bag where it can be easily seen and the bag sealed up. Where possible the humidity indicator should be inspected at frequent intervals to ensure that the condition of the air inside the bag is still `safe' (i.e. the colour of the indicator is blue). If an `unsafe' condition is shown (i.e. the colour of the indicator is lilac or pink) the bag should be inspected and repaired as necessary, and the desiccant renewed.


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An Engine within a Moisture Proof Bag Mounted in a Transit Figure 22.1. Stand

22.1.3 STORAGE.

Complete engines and individual components should be kept in a clean, wellventilated store with an even temperature of 10 to 20°C. Components should be stored in open racks in their original packing and rubber items kept away from strong sunlight, oil, grease or heat sources. Any desiccant packs attached to stored components should be checked frequently for moisture contamination. With certain components (rubber seals, etc) the manufacturer may recommend that the number of components in a stack is limited to a specific number to prevent distortion. Components that have a shelf life should be used in sequence, any that become time expired being removed for overhaul, test and repacking.


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MODULE 15. GAS TURBINE ENGINE

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