Appendix D

AIRWORTHINESS QUALIFICATION REQUIREMENTS

ENGINE, AIRCRAFT, TURBOSHAFT

DISTRIBUTION STATEMENT A. Approved for public release. Distribution is unlimited.

Appendix D

2

FOREWORD

1. This document provides guidance for the qualification of engines, aircraft, turboshaft and isapproved for use by the U.S. Army Combat Capabilities Development Command, Aviation &Missile Center, Aviation Engineering Directorate and is available for use by all Departmentsand Agencies of the Department of Defense.

2. This document is intended for application to manned or unmanned fixed- or rotary-wingmilitary aircraft

3. The testing protocols and procedures defined herein are to be followed by enginemanufacturers in order for the engine and its components to receive an airworthiness release(AWR).

4. Additional platform-level or system integration requirements for an engine and/or itscomponents are not addressed herein.

5. Comments, suggestions, or questions on this document should be addressed to:CCDC Aviation & Missile Center Attn: Aviation Engineering Directorate (FCDD-AEP) Bldg 4488 Redstone Arsenal, AL 35898-5000

or e-mailed to: Curtis.J.Stevens2.civ@mail.mil

Appendix D

3

SUMMARY OF CHANGE MODIFICATIONS

1. The following changes have been made: Not applicable; initial issue.

PARAGRAPH MODIFICATION

UNITED STATES ARMY CCDC AVIATION & MISSILE CENTER,

AVIATION ENGINEERING DIRECTORATE REDSTONE ARSENAL, ALABAMA 35898

FUNCTIONAL DIVISION: ________________________ Curtis J. Stevens Chief, Propulsion Division

APPROVED BY: ________________________________ David G. Stephan Associate Director for Technology Aviation Engineering Directorate

DATE: _______________________

Appendix D

4

TABLE OF CONTENTS 1. SCOPE. ........................................................................................................... 24 1.1 Metric Requirements. ..................................................................................... 24 1.2 Deviations. ....................................................................................................... 24 2. APPLICABLE DOCUMENTS........................................................................... 24 2.1 General. ........................................................................................................... 24 2.2 Government Documents. .................................................................................. 24 2.2.1 Specifications, Standards, and Handbooks. ..................................................... 24 2.2.2 Other Government Documents, Drawings, and Publications. ........................... 25 2.3 Nongovernment Publications. ........................................................................... 26 3. REQUIREMENTS. ........................................................................................... 29 3.1 Item Definition. ................................................................................................. 29 3.1.1 Item Diagram. ................................................................................................... 29 3.1.2 Interface Definition ........................................................................................... 29 3.1.2.1 Drawings. ......................................................................................................... 29 3.1.2.2 Not used. .......................................................................................................... 30 3.1.2.3 Installation Interfaces. ...................................................................................... 30 3.1.2.4 Moments of Inertia. ........................................................................................... 30 3.1.2.5 Externally Applied Forces. ................................................................................ 30 3.1.2.5.1 Gyroscopic Moments. ....................................................................................... 31 3.1.2.6 Mounts. ........................................................................................................... 31 3.1.2.6.1 Main Mounts. .................................................................................................... 31 3.1.2.6.2 Ground Handling Mounts. ................................................................................. 32 3.1.2.6.3 Engine Stiffness and Modes. ............................................................................ 32 3.1.2.7 Pads and Drives. .............................................................................................. 32 3.1.2.8 Engine Surface Temperature and Heat Rejection............................................. 32 3.1.2.8.1 Engine Component Limiting Temperature. ....................................................... 33 3.1.2.8.2 Heat Rejection and Cooling Analysis. ............................................................... 33 3.1.2.9 Air and Gas Leakage. ...................................................................................... 33 3.1.2.10 Engine Air Inlet System. .................................................................................. 34 3.1.2.10.1 Air Inlet Design and Dimensions. ...................................................................... 34 3.1.2.10.2 Allowable Inlet Connection Loads. .................................................................... 34 3.1.2.10.3 Inlet Airflow Distortion Limits. ........................................................................... 34 3.1.2.10.4 Pressure and Temperature Rate of Change. .................................................... 35 3.1.2.11 Customer Bleed Air System. ........................................................................... 35 3.1.2.11.1 Allowable Customer Bleed Connection Loads. ................................................. 36 3.1.2.11.2 Start and Acceleration Bleed Air. ...................................................................... 36 3.1.2.11.3 Bleed Air Contamination. .................................................................................. 36 3.1.2.12 Not used. ......................................................................................................... 37 3.1.2.13 Connections. ................................................................................................... 37 3.1.2.14 Shaft Power Absorber. .................................................................................... 37 3.1.2.14.1 Power Absorber to Engine Interface Characteristics. ........................................ 37 3.1.2.14.2 Not used. .......................................................................................................... 38 3.1.2.14.3 Not used. .......................................................................................................... 38 3.1.2.14.4 Turboshaft Engine Output Drives. .................................................................... 38 3.1.3 Major Component List. ..................................................................................... 38 3.1.4 Reliability. ......................................................................................................... 39 3.1.4.1 Reliability Quantitative Requirements. .............................................................. 39 3.1.5 Maintainability. ................................................................................................. 39 3.1.5.1 Peculiar Support Equipment (PSE). ................................................................. 39

Appendix D

5

3.1.6 Engine Monitoring System (EMS) ..................................................................... 40 3.1.6.1 EMS Software Development. ........................................................................... 40 3.1.6.2 EMS Functionality. ........................................................................................... 40 3.1.6.3 EMS Required Data. ........................................................................................ 40 3.1.6.4 LRU Fault Isolation. .......................................................................................... 41 3.2 Characteristics. ................................................................................................ 41 3.2.1 Performance Characteristics. ........................................................................... 41 3.2.1.1 Performance Ratings. ....................................................................................... 41 3.2.1.2 Performance Presentation. ............................................................................... 41 3.2.1.2.1 Performance Presentation Digital Computer Program. ..................................... 42 3.2.1.2.2 Engine Transient Program. ............................................................................... 44 3.2.1.3 Performance Verification. ................................................................................. 45 3.2.1.4 Operating Limits. .............................................................................................. 46 3.2.1.4.1 Operating Envelope. ......................................................................................... 46 3.2.1.4.2 Sea Level Operating Limits. ............................................................................. 46 3.2.1.4.3 Absolute Altitude. ............................................................................................. 46 3.2.1.4.4 Starting Limits. ................................................................................................. 46 3.2.1.4.5 Engine Temperature Limits. ............................................................................. 46 3.2.1.4.6 Rotor Speed Limits. .......................................................................................... 47 3.2.1.4.7 Fuel Flow Limits. .............................................................................................. 47 3.2.1.4.8 Oil Pressure and Temperature Limits. .............................................................. 47 3.2.1.4.9 Oil Consumption Limits..................................................................................... 47 3.2.1.4.10 Vibration Limits. ................................................................................................ 47 3.2.1.4.11 Output Shaft Torque Limits. .............................................................................. 48 3.2.1.4.12 Output Shaft Speed Limits. ............................................................................... 48 3.2.1.4.13 Customer Bleed Air And Power Extraction Limits. ............................................ 48 3.2.1.4.14 Emergency Power. ........................................................................................... 48 3.2.1.5 Operating Characteristics. ................................................................................ 48 3.2.1.5.1 Operating Attitude and Conditions. ................................................................... 48 3.2.1.5.2 Starting ............................................................................................................. 48 3.2.1.5.3 Stopping. .......................................................................................................... 48 3.2.1.5.4 Low Power Conditions. ..................................................................................... 49 3.2.1.5.4.1 Idle. .................................................................................................................. 49 3.2.1.5.4.2 No Load Condition. ........................................................................................... 49 3.2.1.5.5 Stability. ........................................................................................................... 49 3.2.1.5.6 Engine Power Transients. ................................................................................ 49 3.2.1.5.6.1 Minimum Requirements. ................................................................................... 49 3.2.1.5.6.2 Estimates. ........................................................................................................ 50 3.2.1.5.7 Engine Windmilling Capability. ......................................................................... 50 3.2.1.5.8 Rotor Droop and Overshoot. ............................................................................ 51 3.2.2 Physical Characteristics. .................................................................................. 51 3.2.2.1 Dry Weight of Complete Engine. ...................................................................... 51 3.2.2.2 Weight of Residual Fluids. ................................................................................ 51 3.2.3 Not used. .......................................................................................................... 51 3.2.4 Not used. .......................................................................................................... 51 3.2.5 Environmental Conditions. ................................................................................ 51 3.2.5.1 Ambient Temperature Conditions. .................................................................... 51 3.2.5.2 Icing Conditions. ............................................................................................... 52 3.2.5.3 Not used. .......................................................................................................... 52 3.2.5.4 Not used. .......................................................................................................... 52 3.2.5.5 Corrosive Atmosphere Conditions. ................................................................... 52

Appendix D

6

3.2.5.6 Environmental Ingestion Capability. ................................................................. 52 3.2.5.6.1 Bird Ingestion. .................................................................................................. 52 3.2.5.6.2 Foreign Object Damage (FOD). ........................................................................ 54 3.2.5.6.3 Ice Ingestion. .................................................................................................... 54 3.2.5.6.4 Sand Ingestion. ................................................................................................ 54 3.2.5.6.4.1 Fine Sand Ingestion. ........................................................................................ 54 3.2.5.6.4.2 Coarse Sand Ingestion. .................................................................................... 55 3.2.5.6.5 Atmospheric Liquid Water Ingestion. ................................................................ 56 3.2.5.7 Noise Levels. .................................................................................................... 56 3.2.5.8 Exhaust Gas Contamination. ............................................................................ 56 3.2.5.8.1 Exhaust Smoke Emission. ................................................................................ 56 3.2.5.8.2 Invisible Exhaust Mass Emissions. ................................................................... 56 3.2.6 Transportability. ................................................................................................ 57 3.2.7 External Surfaces. ............................................................................................ 57 3.2.8 Survivability and Vulnerability. .......................................................................... 57 3.2.8.1 Ballistic Vulnerability. ........................................................................................ 57 3.2.8.2 Not used. .......................................................................................................... 57 3.2.8.3 Not used. .......................................................................................................... 57 3.2.8.4 Electromagnetic Environmental Effects (E3). .................................................... 57 3.2.8.4.1 Electromagnetic Interference (EMI). ................................................................. 57 3.2.8.4.2 Not used. .......................................................................................................... 57 3.2.8.4.3 Electromagnetic Pulse. ..................................................................................... 57 3.2.8.4.4 Lightning. ......................................................................................................... 58 3.2.8.4.5 Nuclear Weapons Effects. ................................................................................ 58 3.3 Design and Construction. ................................................................................. 58 3.3.1 Materials, Processes, and Fasteners................................................................ 58 3.3.1.1 Materials and Processes. ................................................................................. 58 3.3.1.1.1 Adhesives and Sealants. .................................................................................. 58 3.3.1.1.2 Elastomeric Materials. ...................................................................................... 58 3.3.1.1.3 “O” Ring Seals and Packing. ............................................................................ 58 3.3.1.1.4 Corrosion Protection. ........................................................................................ 58 3.3.1.2 Fasteners. ....................................................................................................... 59 3.3.1.2.1 Self-Retaining Bolts. ......................................................................................... 59 3.3.1.2.2 Securing of Fasteners. ..................................................................................... 59 3.3.2 Not used. .......................................................................................................... 59 3.3.3 Nameplate and Product Marking. ..................................................................... 59 3.3.3.1 Engine data plate marking. ............................................................................... 59 3.3.3.2 Critical parts/critical safety items identification and tracking. ............................. 59 3.3.3.3 LRU/WRA marking. .......................................................................................... 59 3.3.4 System Nuclear Survivability. ........................................................................... 59 3.3.5 Interchangeability. ............................................................................................ 59 3.3.6 Safety. .............................................................................................................. 60 3.3.6.1 Flammable Fluid Systems. ............................................................................... 60 3.3.6.2 Fire Prevention. ................................................................................................ 60 3.3.6.3 Explosion-Proof. ............................................................................................... 60 3.3.6.4 Fluid Leakage. .................................................................................................. 60 3.3.6.5 Combustible Fluid Drains. ................................................................................ 60 3.3.6.6 Ground Safety. ................................................................................................. 60 3.3.6.7 Engine Control System Safety. ......................................................................... 61 3.3.7 Not used. .......................................................................................................... 61 3.3.8 Structural Performance..................................................................................... 61

Appendix D

7

3.3.8.1 Design Service Life. ......................................................................................... 61 3.3.8.1.1 Expendables Design Service Life ..................................................................... 62 3.3.8.1.2 Parts Classification – Critical Safety Items........................................................ 62 3.3.8.2 High Cycle Fatigue Life (HCF). ......................................................................... 62 3.3.8.3 Low Cycle Fatigue (LCF) Life. .......................................................................... 62 3.3.8.4 Strength. .......................................................................................................... 63 3.3.8.4.1 Factors of Safety. ............................................................................................. 63 3.3.8.4.2 Blade and Disk Deflection. ............................................................................... 63 3.3.8.5 Material Properties. .......................................................................................... 64 3.3.8.6 Strength and Life Analysis. ............................................................................... 64 3.3.8.6.1 Points of Life. ................................................................................................... 65 3.3.8.7 Failsafe Design. .............................................................................................. 67 3.3.8.8 Creep. .............................................................................................................. 67 3.3.8.9 Containment and Rotor Structural Integrity. ...................................................... 67 3.3.8.9.1 Containment ..................................................................................................... 67 3.3.8.9.2 Rotor Integrity. .................................................................................................. 68 3.3.8.9.3 Blade Out.. ....................................................................................................... 68 3.3.8.9.4 Disk Burst Speed.............................................................................................. 69 3.3.8.10 Vibration and Dynamic Response. .................................................................. 69 3.3.8.10.1 Critical Speeds. ................................................................................................ 69 3.3.8.10.2 Vibration and Stress Analysis. .......................................................................... 70 3.3.8.11 Damage Tolerance. ......................................................................................... 70 3.3.8.11.1 Residual Strength. ............................................................................................ 71 3.3.8.11.2 Initial Production and In-Service Flaw Size. ...................................................... 71 3.3.9 Design Control. ................................................................................................ 71 3.3.9.1 Standardization. ............................................................................................... 71 3.3.9.1.1 Not used. .......................................................................................................... 72 3.3.9.1.2 Not used. .......................................................................................................... 72 3.3.9.1.3 Not used. .......................................................................................................... 72 3.3.9.2 Parts List. ......................................................................................................... 72 3.3.9.3 Assembly of Components and Parts................................................................. 72 3.3.9.4 Design and Construction Changes. .................................................................. 72 3.3.10 Bearings. .......................................................................................................... 73 3.4 Not used. .......................................................................................................... 73 3.5 Logistics. .......................................................................................................... 73 3.5.1 Maintenance. .................................................................................................... 73 3.5.1.1 Modular Concept. ............................................................................................. 73 3.5.1.2 Borescope. ....................................................................................................... 73 3.6 Not used. .......................................................................................................... 73 3.7 Major Component Characteristics. ................................................................... 73 3.7.1 Anti-icing System.............................................................................................. 73 3.7.1.1 Anti-Icing System Activation. ............................................................................ 74 3.7.1.2 Type of Anti-Icing. ............................................................................................ 74 3.7.2 Engine Control System. .................................................................................... 74 3.7.2.1 Control Signals. ................................................................................................ 74 3.7.2.1.1 Control Lever Torque. ...................................................................................... 74 3.7.2.1.2 Control Lever Rigging. ...................................................................................... 75 3.7.2.2 Engine Control System Performance. ............................................................... 75 3.7.2.2.1 Power Modulation. ............................................................................................ 75 3.7.2.2.2 Region of Control Limiting Functions. ............................................................... 75 3.7.2.2.3 Control System Reliability. ................................................................................ 75

Appendix D

8

3.7.2.2.3.1 Manual Mode Operation. .................................................................................. 76 3.7.2.2.3.2 Engine Operating Mode Selection. .................................................................. 76 3.7.2.2.3.3 Fail Fixed Mode. .............................................................................................. 76 3.7.2.2.3.4 Extended Redundancy ..................................................................................... 76 3.7.2.3 Engine Control System Design. ....................................................................... 77 3.7.2.3.1 Engine Control System Adjustment. ................................................................. 77 3.7.2.3.2 Not used. .......................................................................................................... 77 3.7.2.3.3 Overspeed Protection. ...................................................................................... 77 3.7.2.3.3.1 Power Turbine Overspeed Control System....................................................... 77 3.7.2.3.3.2 Gas Generator Overspeed Control System. ..................................................... 77 3.7.2.3.4 Engine Overtemperature Protection. ................................................................ 78 3.7.2.3.4.1 Engine Start Overtemperature System. ............................................................ 78 3.7.2.3.4.2 Steady-State Overtemperature. ........................................................................ 78 3.7.2.3.5. Engine Control System Reprogramming. ......................................................... 78 3.7.2.3.6 Software Growth Capability. ............................................................................. 78 3.7.2.3.7 Engine Control System Software Development. ............................................... 78 3.7.2.4 Incident Recorder ............................................................................................. 79 3.7.2.4.1 Incident Detection............................................................................................. 79 3.7.2.4.2 Incident Data .................................................................................................... 79 3.7.2.4.3 Incident Data Storage ....................................................................................... 80 3.7.2.4.4 Incident Data Retrieval ..................................................................................... 80 3.7.3 Fuel System. .................................................................................................... 80 3.7.3.1 Fuel System Interface. ..................................................................................... 80 3.7.3.1.1 Not used ........................................................................................................... 80 3.7.3.1.2 Maximum Fuel Flow. ........................................................................................ 80 3.7.3.1.3 Fuel Inlet. ......................................................................................................... 80 3.7.3.1.3.1 Fuel Inlet Dimensions. ...................................................................................... 80 3.7.3.1.3.2 Allowable Fuel Inlet Connection Loads. ............................................................ 80 3.7.3.1.3.3 Fuel Inlet Pressure and Temperature. .............................................................. 80 3.7.3.2 Fuels. ............................................................................................................... 80 3.7.3.2.1 Primary Fuel. .................................................................................................... 80 3.7.3.2.2 Restricted Fuel. ................................................................................................ 81 3.7.3.2.3 Emergency Fuel. .............................................................................................. 81 3.7.3.3 Fuel System Performance. ............................................................................... 81 3.7.3.3.1 Fuel System Calibration Limits. ........................................................................ 81 3.7.3.3.2 Fuel Contamination. ......................................................................................... 81 3.7.3.3.3 Fuel System Performance with External Assistance. ........................................ 81 3.7.3.3.4 Fuel System Performance with No External Assistance. .................................. 81 3.7.3.3.4.1 Pump Priming and Dry Lift. .............................................................................. 82 3.7.3.3.5 Fuel System Bubble Ingestion .......................................................................... 82 3.7.3.3.6 Fuel Resistance. .............................................................................................. 82 3.7.3.4 Fuel Filter. ........................................................................................................ 82 3.7.3.5 Not used. .......................................................................................................... 83 3.7.4 Electrical Systems. ........................................................................................... 83 3.7.4.1 Electrical Power. .............................................................................................. 83 3.7.4.1.1 Alternator. ........................................................................................................ 83 3.7.4.2 Alternate Electrical Power. ............................................................................... 83 3.7.4.3 Electrical Interface. ........................................................................................... 84 3.7.4.3.1 External Electrical Power. ................................................................................. 84 3.7.4.3.2 Electrical Connectors and Cables . ............................................................... 84 3.7.4.3.3 Digital Communication Signal Interface. ........................................................... 84

Appendix D

9

3.7.4.4 Electrical and Electronic Equipment. ................................................................ 84 3.7.4.5 Electrical Bonding. ............................................................................................ 84 3.7.4.6 Ground Isolation. .............................................................................................. 85 3.7.4.7 Potting Compounds. ......................................................................................... 85 3.7.4.8 Airframe Load Impedance Requirement. .......................................................... 85 3.7.4.9 Wire Faults ....................................................................................................... 85 3.7.4.9.1 Wiring Faults – Observed Short Circuit Conditions ........................................... 85 3.7.4.9.2 Control System Response to Wiring Faults (Afflicted Channel) ........................ 85 3.7.4.9.3 Control System Response to Wiring Faults (Non-Afflicted Channel) ................. 86 3.7.5 Ignition System. ................................................................................................ 86 3.7.5.1 Ignition System Interface. ................................................................................. 86 3.7.5.2 Ignition System Performance............................................................................ 86 3.7.5.3 Ignition System Fouling. ................................................................................... 86 3.7.5.3.1 Carbon Fouling. ................................................................................................ 86 3.7.5.3.2 Water Fouling. .................................................................................................. 86 3.7.6 Instrumentation Systems. ................................................................................. 86 3.7.6.1 Instrumentation Interface. ................................................................................ 87 3.7.6.1.1 Condition Indication. ......................................................................................... 87 3.7.6.1.2 Engine Condition Monitoring. ........................................................................... 87 3.7.6.3 Not used. ......................................................................................................... 87 3.7.6.4 Temperature Sensing Systems. ....................................................................... 87 3.7.6.5 Vibration Measurement. ................................................................................... 87 3.7.6.6 Torque Indication. ............................................................................................ 87 3.7.6.7 Engine Component Life Counter. .................................................................... 88 3.7.6.7.1 EMS Algorithms ............................................................................................... 88 3.7.6.8 Speed Indication.............................................................................................. 88 3.7.7 Engine Lubricating System. .............................................................................. 88 3.7.7.1 Lubricating System Interface. ........................................................................... 88 3.7.7.1.1 Oil System Installation and Servicing. ............................................................... 88 3.7.7.2 Lubricants......................................................................................................... 88 3.7.7.2.1 Lubricating Oil. ................................................................................................. 88 3.7.7.3 Lubricating System Performance. ..................................................................... 89 3.7.7.3.1 Oil Flow and Heat Rejection. ............................................................................ 89 3.7.7.3.2 Internal Oil Leakage. ........................................................................................ 90 3.7.7.3.3 Oil Flow Interruption. ........................................................................................ 90 3.7.7.3.4 Loss of Oil. ....................................................................................................... 90 3.7.7.4 Lubrication System Components and Features. ............................................... 90 3.7.7.4.1 Oil Reservoir. ................................................................................................... 90 3.7.7.4.1.1 Oil Reservoir Capacity. ..................................................................................... 90 3.7.7.4.1.2 Oil Reservoir External Features. ....................................................................... 91 3.7.7.4.2 Oil Drains. ........................................................................................................ 91 3.7.7.4.3 Oil Filter. ........................................................................................................... 91 3.7.7.4.4 Oil Debris and Condition Monitor. ..................................................................... 92 3.7.7.4.5 Oil Coolers. ...................................................................................................... 92 3.7.7.4.6 Not used. .......................................................................................................... 92 3.7.7.4.7 Oil Pressure Indication. .................................................................................... 92 3.7.8 Not used. .......................................................................................................... 92 3.7.9 Starting System. ............................................................................................... 92 3.7.9.1 Starting System Interface. ................................................................................ 92 3.7.9.1.1 Starting Torque and Speed Requirements. ...................................................... 92 3.7.9.1.2 Moment of Inertia of Rotating Parts. ................................................................. 93

Appendix D

10

3.7.9.1.3 Torsional Spring Constant. ............................................................................... 93 3.7.9.1.4 Starter Train Backlash. ..................................................................................... 93 3.7.9.2 Starting Requirements. .................................................................................... 93 3.7.9.2.1 Restart Time. .................................................................................................... 93 3.7.9.3 Starting Procedure. .......................................................................................... 93 3.7.9.4 Starting Drive Train. ......................................................................................... 94 3.7.10 Exhaust Nozzle System.................................................................................... 94 3.7.10.1 Allowable Exhaust System Connection Loads. ................................................ 94 3.7.10.2 Exhaust Nozzle. .............................................................................................. 94 3.7.10.3 Jet Wake Diagrams. ........................................................................................ 94 3.7.11 Not used. .......................................................................................................... 95 3.7.12 Not used. .......................................................................................................... 95 3.7.13 Wash System. .................................................................................................. 95 3.8 Not used. .......................................................................................................... 95 3.9 Quality Verification. .......................................................................................... 95 3.9.1 Engineering Evaluation Tests. .......................................................................... 95 3.9.2 Engine Integrity Testing (EIT). .......................................................................... 95 3.9.3 Preliminary Flight Rating (PFR). ....................................................................... 95 3.9.4 Qualification Test (QT) Rating. ......................................................................... 95 3.9.5 Operational Capability Release (OCR) Test. .................................................... 95 3.9.7 Engine Development Special Tests. ................................................................. 96 3.9.8 Acceptance Test (AT). ...................................................................................... 96 4 VERIFICATION. ............................................................................................... 97 4.1 General. ........................................................................................................... 97 4.1.1 Responsibility. .................................................................................................. 97 4.2 Quality Conformance Inspections. .................................................................... 97 4.2.1 Quality Evidence. ............................................................................................. 97 4.3 Manner of Test and Reporting. ......................................................................... 97 4.3.1 Test Surveillance. ............................................................................................. 97 4.3.2 Test Article Configuration. ................................................................................ 97 4.3.2.1 Test Engines. ................................................................................................... 97 4.3.2.2 Method of Qualification. .................................................................................... 98 4.3.3 Test Apparatus. ................................................................................................ 98 4.3.3.1 Automatic Recording Equipment. ..................................................................... 98 4.3.3.2 Vibration Measuring Equipment and Response Characteristics. ....................... 98 4.3.3.3 Test Stand Characteristics. .............................................................................. 98 4.3.3.4 Starter. ............................................................................................................. 99 4.3.4 Test Condition. ................................................................................................. 99 4.3.4.1 Servicing. ......................................................................................................... 99 4.3.4.1.1 Oil Servicing. .................................................................................................... 99 4.3.4.2 Inlet and Exhaust Duct Connections. ................................................................ 99 4.3.4.3 Bleed Air Connections. ..................................................................................... 99 4.3.4.4 Accessory Drive Gearboxes. ............................................................................ 99 4.3.4.5 Accreditable Test Time. .................................................................................... 99 4.3.4.6 Fuel Properties for Test. ................................................................................... 99 4.3.5 Data. ................................................................................................................ 99 4.3.5.1 Test Plans and Procedures. ........................................................................... 100 4.3.5.1.1 Engine and Engine Subassembles. ................................................................ 100 4.3.5.1.2 Components. .................................................................................................. 100 4.3.5.2 Preliminary Data. ............................................................................................ 100 4.3.5.3 Accuracy of Data. ........................................................................................... 100

Appendix D

11

4.3.5.3.1 Measurement Uncertainty Analysis. ............................................................... 101 4.3.5.4 Steady-State Data. ......................................................................................... 101 4.3.5.5 Transient Data. ............................................................................................... 101 4.3.5.6 Starting Data. ................................................................................................. 101 4.3.5.7 Miscellaneous Data. ....................................................................................... 102 4.3.5.8 Test Notes. ..................................................................................................... 102 4.3.5.9 Barometer Reading. ....................................................................................... 102 4.3.5.10 Relative Humidity Data. ................................................................................. 102 4.3.5.11 Fuel and Oil Data. .......................................................................................... 102 4.3.6 Reports. ......................................................................................................... 102 4.3.6.1 Test Reports. .................................................................................................. 102 4.3.6.2 Summary Reports. ......................................................................................... 103 4.4 Engineering Evaluation Tests. ........................................................................ 104 4.4.1 Customer Bleed Air. ....................................................................................... 104 4.4.2 Engine Heat Rejection and Oil Cooling. .......................................................... 104 4.4.3 Oil Flow Interruption Test. .............................................................................. 104 4.4.4 Engine Electrical Power Failure Tests. ........................................................... 104 4.4.4.1 Engine System Electrical Power Tests. .......................................................... 105 4.4.4.2 Alternate Electrical Power Tests. .................................................................... 106 4.4.5 Engine Vibration Survey. ................................................................................ 107 4.4.6 Starting Torque. ............................................................................................. 107 4.4.7 Not used. ........................................................................................................ 108 4.4.8 Maintenance Test. .......................................................................................... 108 4.4.9 Not used. ........................................................................................................ 108 4.4.10 Verification of Correction Factors. .................................................................. 108 4.5 Engine Integrity Testing (EIT)/Preliminary Flight Rating (PFR). ...................... 108 4.5.1 Endurance Test .............................................................................................. 108 4.5.1.1 Pretest Verification. ........................................................................................ 108 4.5.1.1.1 Engine Dry Weight. ........................................................................................ 108 4.5.1.2 Calibration. ..................................................................................................... 108 4.5.1.2.1 Component Calibration. .................................................................................. 108 4.5.1.2.2 Engine Calibration. ......................................................................................... 108 4.5.1.2.2.1 Customer Bleed Air Analysis. ......................................................................... 109 4.5.1.3 Endurance Test Procedure. ............................................................................ 109 4.5.1.3.1 Starts. ............................................................................................................. 113 4.5.1.3.2 Contingency Power Rating. ............................................................................ 113 4.5.1.4 Recalibrations. ............................................................................................... 114 4.5.1.4.1 Engine Recalibration. ..................................................................................... 114 4.5.1.4.2 Component Recalibration. .............................................................................. 114 4.5.1.5 Engine Disassembly and Inspection. .............................................................. 114 4.5.1.6 Endurance Test Completion. .......................................................................... 114 4.5.2 Engine Component Tests. .............................................................................. 115 4.5.2.1 Previous Component Approval. ...................................................................... 115 4.5.2.2 Component Acceptance Test. ........................................................................ 115 4.5.2.3 Component Re-Test, Disassembly, and Inspection. ...................................... 115 4.5.2.4 Component Test Success Criteria. ................................................................ 115 4.5.2.5 Component Test Procedures. ........................................................................ 116 4.5.2.5.1 Explosive Atmosphere. ................................................................................... 116 4.5.2.5.2 Vibration – Airframe and Engine. .................................................................... 117 4.5.2.5.3 Fuel Pump Altitude Test. ................................................................................ 117 4.5.2.5.4 Oil Reservoir Pressure Test. .......................................................................... 118

Appendix D

12

4.5.2.5.5 Fire Test. ........................................................................................................ 118 4.5.2.5.6 Overheat Test. ............................................................................................... 118 4.5.2.5.7 Output Shaft Static Torque Test. .................................................................... 119 4.5.2.5.8 Impact (Shock) Test. ..................................................................................... 119 4.5.2.5.9 Software/Programmable Hardware Verification ............................................. 119 4.5.2.5.10 Electrical Power System Test ........................................................................ 119 4.5.2.5.10.1 Engine-Supplied Electrical Power ................................................................... 119 4.5.2.5.11 Helicopter Drive System Torsional Stability ................................................... 121 4.5.3 Altitude Tests. ................................................................................................ 121 4.5.3.1 Altitude Calibration. ........................................................................................ 122 4.5.3.2 Altitude Test Procedure. ................................................................................. 122 4.5.3.3 Altitude Test Completion. ............................................................................... 122 4.5.4 Structural Tests. ............................................................................................. 123 4.5.4.1 Component Vibration Characterization. ......................................................... 123 4.5.4.2 Rotor Structural Integrity. ................................................................................ 123 4.5.4.2.1 Overspeed...................................................................................................... 123 4.5.4.2.2 Overtemperature. ........................................................................................... 124 4.5.4.3 Engine Static Load Test. ................................................................................ 124 4.5.4.4 Attitude Test. .................................................................................................. 124 4.5.4.5 Loss of Oil Test. ............................................................................................. 124 4.5.4.6 Engine Crash Load Test. ................................................................................ 125 4.5.4.7 Not used. ........................................................................................................ 125 4.5.4.8 Low Cycle Fatigue Engine Test. .................................................................... 125 4.5.4.9 Main Shaft, Seals and Bearing Mechanical Tests. .......................................... 126 4.5.5 Environmental Tests ....................................................................................... 126 4.5.5.1 Electromagnetic Environmental Effects (E3). ................................................. 126 4.5.5.2 Electromagnetic Interference (EMI). .............................................................. 126 4.6 Qualification Test (QT) Rating. ....................................................................... 126 4.6.1 Endurance Test. ............................................................................................. 126 4.6.1.1 Pretest Verification. ....................................................................................... 126 4.6.1.1.1 Engine Dry Weight. ........................................................................................ 126 4.6.1.2 Calibrations (QT). .......................................................................................... 127 4.6.1.2.1 Component Calibration. .................................................................................. 127 4.6.1.2.2 Engine Calibration. ......................................................................................... 127 4.6.1.2.3 Customer Bleed Air Analysis. ......................................................................... 127 4.6.1.3 Endurance Test Procedure. ........................................................................... 127 4.6.1.3.1 Starts. ............................................................................................................. 132 4.6.1.3.2 Contingency Power Qualification Rating. ........................................................ 132 4.6.1.4 Recalibrations ................................................................................................ 132 4.6.1.4.1 Engine Recalibration. ..................................................................................... 132 4.6.1.4.2 Component Re-calibration. ............................................................................. 133 4.6.1.5 Engine Disassembly and Inspection. .............................................................. 133 4.6.1.6 Not used. ........................................................................................................ 133 4.6.1.7 Endurance Test Completion. .......................................................................... 133 4.6.2 Engine Component Tests. .............................................................................. 134 4.6.2.1 Previous Component Approval. ...................................................................... 134 4.6.2.2 Simulated Operational Component Tests. ...................................................... 134 4.6.2.2.1 Component Acceptance Test. ........................................................................ 134 4.6.2.2.2 Component Test Procedures. ......................................................................... 135 4.6.2.2.2.1 Component Test Cycles (excluding Ignition System). ..................................... 135 4.6.2.2.2.2 Ignition System Test Cycle. ............................................................................ 135

Appendix D

13

4.6.2.2.3 Order of Testing. ............................................................................................ 136 4.6.2.2.3.1 Low Lubricity Fuel Test................................................................................... 136 4.6.2.2.3.2 High Temperature. ......................................................................................... 136 4.6.2.2.3.3 RoomTemperature & Contamination. ............................................................. 137 4.6.2.2.3.4 Low Temperature. .......................................................................................... 138 4.6.2.2.3.5 Temperature & Vibration Cycling. ................................................................... 139 4.6.2.2.3.6 Engine Fuel System Cavitation Endurance..................................................... 140 4.6.2.2.3.7 Fuel System Bubble Ingestion. ....................................................................... 142 4.6.2.2.4 Component Acceptance Test or Recalibration. ............................................... 142 4.6.2.2.5 Component Test Completion. ......................................................................... 143 4.6.2.3 Environmental Component Tests. .................................................................. 143 4.6.2.3.1 Component Acceptance Test or Calibration. .................................................. 143 4.6.2.3.2 Order of Testing. ............................................................................................ 143 4.6.2.3.3 Component Test Success Criteria .................................................................. 143 4.6.2.3.4 Component Test Procedures. ......................................................................... 144 4.6.2.3.4.1 Temperature – High, Low, Transient. ............................................................. 144 4.6.2.3.4.2 Vibration – Airframe and Engine. ................................................................... 146 4.6.2.3.4.3 Impact (Shock). .............................................................................................. 149 4.6.2.3.4.4 Gunfire Shock. ............................................................................................... 149 4.6.2.3.4.5 Sustained Acceleration. .................................................................................. 149 4.6.2.3.4.6 Low Pressure (Altitude). ................................................................................. 150 4.6.2.3.4.7 Rain................................................................................................................ 150 4.6.2.3.4.8 Explosive Atmosphere. ................................................................................... 150 4.6.2.3.4.9 Fungus. .......................................................................................................... 151 4.6.2.3.4.10 Humidity. ........................................................................................................ 151 4.6.2.3.4.11 Salt Fog. ......................................................................................................... 151 4.6.2.3.4.12 Sand and Dust. .............................................................................................. 151 4.6.2.3.4.13 Contamination by Fluids. ................................................................................ 152 4.6.2.3.4.14 Proof Pressure. .............................................................................................. 152 4.6.2.4 Special Components Tests. ........................................................................... 153 4.6.2.4.1 Oil Reservoir. ................................................................................................. 153 4.6.2.4.2 Accessory Drive. ............................................................................................ 153 4.6.2.4.3 Alternator Test. ............................................................................................... 154 4.6.2.4.4 Heat Exchangers. ........................................................................................... 154 4.6.2.4.5 Fire Test. ........................................................................................................ 155 4.6.2.4.6 Output Shaft Static Torque Test ..................................................................... 155 4.6.2.4.7 Overheat Test. ............................................................................................... 156 4.6.2.4.8 PMA Neutral Short-to-Ground Test. ............................................................... 156 4.6.2.4.9 Overspeed and Containment. ......................................................................... 156 4.6.2.4.10 Burst Pressure. .............................................................................................. 157 4.6.2.4.11 Pressure Cycling. ........................................................................................... 157 4.6.2.4.12 Proof Pressure. .............................................................................................. 157 4.6.2.4.13 Pressure Wash Test. ...................................................................................... 158 4.6.2.4.14 Ignition System Fouling. ................................................................................. 159 4.6.2.4.14.1 Carbon Fouling. .............................................................................................. 159 4.6.2.4.14.2 Water Fouling. ................................................................................................ 159 4.6.2.4.15 Electrical Loads Analysis ................................................................................ 159 4.6.2.4.16 Short Circuit Protection.................................................................................. 159 4.6.2.4.17 Data Bus Specification Compliance ............................................................... 159 4.6.2.5 Software Verification. ..................................................................................... 160 4.6.2.5.1 Engine Testing (Software FQT) ...................................................................... 160

Appendix D

14

4.6.2.5.2 Flight Testing (Software FQT) ........................................................................ 160 4.6.2.6 Complex Power Interrupt Test (Closed Loop). ............................................... 160 4.6.2.7 EMS Lifing Algorithms. .................................................................................. 160 4.6.2.8 Common Mode Multiple Signal Failure Test. ................................................. 161 4.6.3 Altitude Tests. ................................................................................................ 161 4.6.3.1 Altitude Calibration. ........................................................................................ 162 4.6.3.2 Altitude Test Procedure. ................................................................................. 162 4.6.3.3 Altitude Test Completion. ............................................................................... 163 4.6.4 Engine Environmental and Ingestion Tests..................................................... 163 4.6.4.1 Low and High Temperature Starting and Acceleration Test. ........................... 163 4.6.4.2 Environmental Icing Test. ............................................................................... 164 4.6.4.3 Corrosion Susceptibility Test. ........................................................................ 165 4.6.4.4 Bird Ingestion Test. ....................................................................................... 166 4.6.4.5 Foreign Object Damage Test......................................................................... 166 4.6.4.6 Ice Ingestion Test. ......................................................................................... 167 4.6.4.7 Sand Ingestion Tests. .................................................................................... 167 4.6.4.7.1 Fine Sand Ingestion Test. ............................................................................... 167 4.6.4.7.2 Coarse Sand Ingestion Test. .......................................................................... 168 4.6.4.8 Atmospheric Water Ingestion Test. ................................................................ 168 4.6.4.9 Engine Component Limiting Temperature Test. ............................................. 168 4.6.4.10 Noise Survey. ................................................................................................ 169 4.6.4.11 Exhaust Gas Emission Test. ......................................................................... 172 4.6.4.11.1 Exhaust Smoke Emission. .............................................................................. 172 4.6.4.11.2 Invisible Exhaust Mass Emissions. ................................................................. 172 4.6.4.12 Attitude Test. ................................................................................................. 172 4.6.4.13 Loss of Oil Test. ............................................................................................ 173 4.6.4.14 Electromagnetic Environmental Effects (E3). ................................................. 173 4.6.4.14.1 Electromagnetic Interference (EMI). ............................................................... 173 4.6.4.14.2 Not used. ........................................................................................................ 173 4.6.4.14.3 Not used. ........................................................................................................ 173 4.6.4.14.4 Lightning. ....................................................................................................... 173 4.6.5 Engine Characteristics and Fuel Tests. .......................................................... 173 4.6.5.1 Starting Torque. ............................................................................................. 173 4.6.5.2 Not used. ........................................................................................................ 173 4.6.5.3 Not used. ........................................................................................................ 173 4.6.5.4 Not used. ........................................................................................................ 173 4.6.5.5 Emergency Fuel Test. .................................................................................... 173 4.6.6 Structural Tests. ............................................................................................. 174 4.6.6.1 Emergency Power Demonstration ................................................................. 174 4.6.6.2 Low Cycle Fatigue Tests. .............................................................................. 174 4.6.6.2.1 Low Cycle Fatigue Component Tests. ............................................................ 174 4.6.6.2.2 Low Cycle Fatigue Engine Test. ..................................................................... 175 4.6.6.3 Containment. .................................................................................................. 176 4.6.6.4 Rotor Integrity ................................................................................................ 177 4.6.6.4.1 Overspeed...................................................................................................... 177 4.6.6.4.2 Overtemperature. ........................................................................................... 177 4.6.6.4.3 Disk Burst ....................................................................................................... 178 4.6.6.5 Static Load Tests............................................................................................ 178 4.6.6.5.1 Engine Static Load Test. ................................................................................ 178 4.6.6.6 Vibration and Stress Tests. ............................................................................ 178 4.6.6.6.1 Compressor Strain Test. ................................................................................. 179

Appendix D

15

4.6.6.6.2 Gas Generator Turbine Strain Test. ............................................................... 179 4.6.6.6.3 Power Turbine Strain Test. ............................................................................. 179 4.6.6.6.4 Gas Generator Rotor Bearing Evaluation. ...................................................... 179 4.6.6.6.5 Power Turbine Rotor Bearing Evaluation. ....................................................... 179 4.6.6.6.6 Engine Vibration Survey. ................................................................................ 179 4.6.6.6.7 ”Stinger” Rig Vibration Test. ........................................................................... 180 4.6.6.8 Engine Overtemperature Control System Test. .............................................. 181 4.6.6.9 Engine Overspeed Control System Test. ........................................................ 181 4.6.6.10 Main Shaft Bearings Assurance Testing. ....................................................... 181 4.6.6.11 Gear Resonance test. ................................................................................... 182 4.7 Acceptance Test (AT). .................................................................................... 182 4.7.1 Test Apparatus. .............................................................................................. 182 4.7.1.1 Automatic Recording Equipment. ................................................................... 182 4.7.1.2 Vibration Measuring Equipment and Response Characteristics. ..................... 182 4.7.1.3 Test Stand and Test Equipment. .................................................................... 182 4.7.1.3.1 Dynamic Characteristics. ................................................................................ 182 4.7.1.3.2 Power Absorption Characteristics. .................................................................. 183 4.7.1.4 Starter. ........................................................................................................... 183 4.7.2 Test Conditions. ............................................................................................. 183 4.7.2.1 Servicing. ....................................................................................................... 183 4.7.2.1.1 Oil Servicing. .................................................................................................. 183 4.7.2.1.2 Fuel Servicing. ............................................................................................... 183 4.7.2.2 Electrical and Electronic Interference and Susceptibility Check. ..................... 183 4.7.2.3 Component Calibration. .................................................................................. 183 4.7.2.4 Not Used. ....................................................................................................... 183 4.7.2.5 Environmental Stress Screen. ........................................................................ 183 4.7.3 Test Records. ................................................................................................. 184 4.7.3.1 Acceptance Test Log Sheet. .......................................................................... 184 4.7.4 Test Data. ...................................................................................................... 184 4.7.4.1 Preliminary Data. ............................................................................................ 184 4.7.4.2 Steady-State Data. ......................................................................................... 184 4.7.4.3 Transient Data. ............................................................................................... 185 4.7.4.4 Starting Data. ................................................................................................. 185 4.7.4.5 Miscellaneous Data. ....................................................................................... 185 4.7.4.6 Accuracy of Data. ........................................................................................... 186 4.7.5 Test Procedure. .............................................................................................. 186 4.7.5.1 Initial Run. ...................................................................................................... 187 4.7.5.1.1 Inspection After Initial Run. ............................................................................. 188 4.7.5.1.2 Penalty Run. ................................................................................................... 189 4.7.5.1.3 Inspection After Penalty Run. ......................................................................... 189 4.7.5.2 Final Run. ....................................................................................................... 189 4.7.5.2.1 Special Control System Features. .................................................................. 189 4.7.5.3 Rejection and Retest. ..................................................................................... 189 4.7.5.3.1 Engine Vibration. ............................................................................................ 189 4.7.5.3.2 Overtemperature. ........................................................................................... 189 4.7.5.3.3 Stoppage. ....................................................................................................... 189 4.7.5.3.4 Fluid Leakage. ................................................................................................ 189 4.7.5.3.5 Maximum Hours of Running. .......................................................................... 189 4.7.6 Test Completion. ............................................................................................ 190 4.7.7 Sampling Plan. ............................................................................................... 190 4.8 Operational Capability Release (OCR) Test. .................................................. 190

Appendix D

16

4.8.1 Accelerated Endurance Test (AET). ............................................................... 190 4.8.1.1 Endurance Test. ............................................................................................. 190 4.8.1.1.1 Pretest Verification. ........................................................................................ 190 4.8.1.1.1.1 Engine Dry Mass. ........................................................................................... 190 4.8.1.1.2 Calibrations (AET). ......................................................................................... 191 4.8.1.1.2.1 Component Calibration. .................................................................................. 191 4.8.1.1.2.2 Engine Calibration. ......................................................................................... 191 4.8.1.1.3 Endurance Test Procedure. ............................................................................ 191 4.8.1.1.3.1 Starts. ............................................................................................................. 195 4.8.1.1.3.2 Contingency Power Run. ................................................................................ 196 4.8.1.1.4 Recalibrations. ............................................................................................... 196 4.8.1.1.4.1 Engine Recalibration. ..................................................................................... 196 4.8.1.1.4.2 Component Recalibration. .............................................................................. 196 4.8.1.1.5 Engine Disassembly and Inspection. .............................................................. 196 4.8.1.1.6 Endurance Test Completion. .......................................................................... 197 4.8.2 Accelerated Mission Test (AMT). .................................................................... 197 4.8.2.1 Accelerated Mission Test. .............................................................................. 197 4.8.2.1.1 Pretest Verification. ........................................................................................ 197 4.8.2.1.1.1 Engine Dry Weight. ........................................................................................ 197 4.8.2.1.2 Calibrations (AMT). ........................................................................................ 197 4.8.2.1.2.1 Component Calibration. .................................................................................. 197 4.8.2.1.2.2 Engine Calibration. ......................................................................................... 197 4.8.2.1.3 Accelerated Mission Test Procedure. ............................................................. 198 4.8.2.1.4 Recalibrations. ............................................................................................... 199 4.8.2.1.4.1 Engine Recalibration. ..................................................................................... 199 4.8.2.1.4.2 Component Recalibration. .............................................................................. 199 4.8.2.1.5 Engine Disassembly and Inspection. .............................................................. 199 4.8.2.1.6 Accelerated Mission Test Completion. ............................................................ 200 4.8.3 Reliability Evaluation Testing. ......................................................................... 200 4.9 Engine Development Special Tests. ............................................................... 200 4.9.1 PFR Phase. .................................................................................................... 200 4.9.1.1 1000 Cycle Engine Low Cycle Fatigue Test. .................................................. 200 4.9.1.2 Preliminary Icing Test. .................................................................................... 200 4.9.1.3 150 Hour Benchmark Endurance. .................................................................. 200 4.9.1.4 Fine Sand Ingestion Tests. ............................................................................. 201 4.9.1.5 Not Used. ....................................................................................................... 201 4.9.1.6 Highly Accelerated Life Test (HALT)............................................................... 201 4.9.1.7 Inlet Thermal Distortion Test. ......................................................................... 201 4.9.1.8 Inflow Bleed Test. ........................................................................................... 201 4.9.1.9 Engine Back Pressure Test. ........................................................................... 201 4.9.1.11 Not used. ....................................................................................................... 201 4.9.1.12 Not used. ....................................................................................................... 201 4.9.1.13 Fuel System Suction Test. ............................................................................. 201 4.9.2 QT Phase. ...................................................................................................... 202 4.9.2.1 300 Hr Preliminary Endurance Test. ............................................................... 202 4.9.2.2 Preliminary Low Cycle Fatigue (LCF) Test. ................................................... 202 5. PACKAGING. ................................................................................................ 203 5.1 Packaging. ..................................................................................................... 203 6. NOTES. .......................................................................................................... 2036.1 Intended Use. ................................................................................................. 203 6.2 Acquisition Requirements. .............................................................................. 203

Appendix D

17

6.3 Definitions. ..................................................................................................... 203 6.5 Material Safety Data Sheet. ............................................................................ 216 6.6 Subject Term (Key Word) Listing. ................................................................... 217 6.7 Metrication. Not used. ................................................................................... 217 6.8 Engine Specification Preparation. ................................................................... 217 6.8.1 Instruction for Preparation. ............................................................................. 217 APPENDIX E .......................................................................................................................... 275 E.1 SCOPE .......................................................................................................... 275 E.2 APPLICABLE DOCUMENTS.......................................................................... 275 E.3 DEFINITIONS ................................................................................................ 275 E.3.1 General. ......................................................................................................... 275 E.3.1.1 Acronyms used in this appendix. .................................................................... 275 E.3.1.2 Maximum Allowable Transient Rotor Speed. .................................................. 275 E.4 GENERAL REQUIREMENTS......................................................................... 276 E.4.1 General. ......................................................................................................... 276 E.4.2 Content. ......................................................................................................... 276 E.4.3 Material Property Curves. ............................................................................... 276 E.4.4 Analysis Models. ............................................................................................ 276 E.4.5 Points of Life. ................................................................................................. 276 E.4.6 Delivery. ......................................................................................................... 276 E.5 DETAILED REQUIREMENTS. ....................................................................... 277 E.5.1 Introduction. ................................................................................................... 277 E.5.1.1 Purpose and Scope. ....................................................................................... 277 E.5.1.2 Summary. ....................................................................................................... 277 E.5.2 Configuration. ................................................................................................. 277 E.5.2.1 Parts Listing. .................................................................................................. 277 E.5.2.2 Mission Profile. ............................................................................................... 277 E.5.2.3 Cycle Parameters. .......................................................................................... 277 E.5.2.4 Secondary Flow. ............................................................................................ 277 E.5.3 Analyses Methodology/Results. ...................................................................... 278 E.5.3.1 Thermal Analysis. ........................................................................................... 278 E.5.3.2 Stress Analysis. .............................................................................................. 278 E.5.3.3 Strength Analysis. .......................................................................................... 278 E.5.3.4 Low Cycle Fatigue Analysis. ........................................................................... 279 E.5.3.5 Fracture Mechanics Analysis. ......................................................................... 279 E.5.3.6 Creep/Stress Rupture Analysis. ...................................................................... 279 E.5.3.7 Burst Margin Analysis. .................................................................................... 279 E.5.3.8 Oxidation/Corrosion Analysis.......................................................................... 280 E.5.3.9 Containment Analysis. .................................................................................... 280 APPENDIX F .......................................................................................................................... 287 F.1 SCOPE .......................................................................................................... 287 F.1.1 Scope. ............................................................................................................ 287 F.2 APPLICABLE DOCUMENTS.......................................................................... 287 F.3 DEFINITIONS ................................................................................................ 287 F.4 GENERAL REQUIREMENTS......................................................................... 287 F.4.1 General. ......................................................................................................... 287 F.4.2.1 Stress/Thermal Analyses. ............................................................................... 287 F.4.3 Material Property Curves. ............................................................................... 287 F.4.4 Analysis Models. ............................................................................................ 288 F.4.5 Delivery. ......................................................................................................... 288 F.5 DETAILED REQUIREMENTS. ....................................................................... 288

Appendix D

18

F.5.1 Introduction. ................................................................................................... 288 F.5.1.1 Purpose and Scope. ....................................................................................... 288 F.5.1.2 Summary. ....................................................................................................... 288 F.5.2 General Component. ...................................................................................... 289 F.5.2.1 Component Description. ................................................................................. 289 F.5.2.2 Vibration and Stress Analysis. ........................................................................ 289 F.5.2.3 Component Testing. ....................................................................................... 289 F.5.2.4 Engine Testing. .............................................................................................. 289 F.5.2.5 High Cycle Fatigue Assessment. .................................................................... 290 Appendix G ............................................................................................................................ 292 G.1. Introduction. ................................................................................................... 292 G.2. Software Lifecycle Processes. ........................................................................ 294 G.2.1 System Process. ............................................................................................ 294 G.2.1.1 System Process Artifacts. ............................................................................... 296 G.2.1.1.1 Software Safety Program Plan (SwSPP). ....................................................... 296 G.2.1.1.2 System Safety Assessment (SSA). ................................................................ 297 G.2.1.1.2.1 SSA (Sys Arch). ............................................................................................. 297 G.2.1.1.2.2 SSA (SW Arch). ............................................................................................. 297 G.2.1.1.2.3 SSA (SW FMEA). ........................................................................................... 298 G.2.1.1.2.4 SSA (SW FTA). .............................................................................................. 298 G.2.1.1.2.5 SSA (Final). .................................................................................................... 299 G.2.1.1.3 Functional Hazard Assessment (FHA). ........................................................... 299 G.2.1.1.4 System/Subsystem Design Description (SSDD). ............................................ 300 G.2.1.1.5 Data Accession List (DAL). ............................................................................. 301 G.2.1.1.6 Engine Specification. ...................................................................................... 301 G.2.1.1.7 Integrated Master Schedule (IMS). ................................................................. 301 G.2.1.2 System Process Requirements. ..................................................................... 301 G.2.1.2.1 System Level Functional Hazard Assessment (FHA) Requirements............... 302 G.2.1.2.2 System Safety Assessment (SSA) – System Architecture (Sys Arch) Reqmt . 303 G.2.2 Software Planning Process. ........................................................................... 303 G.2.2.1 Software Planning Process Artifacts. .............................................................. 304 G.2.2.1.1 Plan for Software Aspects of Certification (PSAC). ......................................... 304 G.2.2.1.2 Software Development Plan (SDP). ............................................................... 304 G.2.2.1.3 Software Test Plan (STP). .............................................................................. 304 G.2.2.1.4 Software Configuration Management Plan (SCMP). ....................................... 304 G.2.2.1.5 Software Quality Assurance Plan (SQAP). ..................................................... 305 G.2.2.1.6 Report, Record of Meeting/Minutes (RRM). .................................................... 305 G.2.2.1.7 Data Accession List (DAL) .............................................................................. 305 G.2.2.2 Software Planning Process Requirements...................................................... 305 G.2.2.2.1 Plan for Software Aspects of Certification (PSAC) Requirements. .................. 305 G.2.2.2.2 Software Development Plan (SDP) Requirements. ......................................... 306 G.2.2.2.3 Software Test Plan (STP) Requirements. ....................................................... 306 G.2.2.2.4 Software Configuration Management Plan (SCMP) Requirements. ................ 307 G.2.2.2.5 Software Quality Assurance Plan (SQAP) Requirements. .............................. 307 G.2.3 Software Requirements Process. ................................................................... 308 G.2.3.1 Software Requirements Process Artifacts....................................................... 308 G.2.3.1.1 Software Requirements Specification (SRS). ................................................. 308 G.2.3.1.2 Interface Requirements Specification (IRS). ................................................... 308 G.2.3.1.3 System Safety Assessment (SSA) – Software Architecture (SwArch). ........... 308 G.2.3.1.4 Data Accession List (DAL). ............................................................................. 308 G.2.3.1.5 Report, Record of Meeting/Minute (RRM)....................................................... 309

Appendix D

19

G.2.3.2 Software Requirement Process Requirements. .............................................. 309 G.2.3.2.1 Software/Interface Requirements Specification (SRS/IRS) Requirements. ..... 309 G.2.3.2.2 System Safety Assessment (SSA) – Software Architecture (SwArch) Reqmt . 310 G.2.4 Software Design Process. .............................................................................. 311 G.2.4.1 Software Design Process Artifacts. ................................................................ 311 G.2.4.1.1 Software Design Description (SDD)................................................................ 311 G.2.4.1.2 Interface Design Description (IDD). ................................................................ 312 G.2.4.1.3 Software Test Description (STD). ................................................................... 312 G.2.4.1.4 System Safety Assessment (SSA) – SW FMEA. ............................................ 312 G.2.4.1.5 Safety Assessment Report (SAR). ................................................................. 312 G.2.4.1.6 Report, Record of Meeting/Minutes (RRM). .................................................... 312 G.2.4.1.7 Data Accession List (DAL). ............................................................................. 312 G.2.4.2 Software Design Process Requirement. ......................................................... 313 G.2.4.2.1 Software/Interface Design Description (SDD/IDD) Requirements. .................. 313 G.2.4.2.2 Software Test Description (STD) Requirement. .............................................. 315 G.2.4.2.3 System Safety Assessment (SSA) – SW FMEA Requirements. ..................... 315 G.2.4.2.4 Safety Assessment Report (SAR) Requirements............................................ 316 G.2.5 Software Coding Process. .............................................................................. 316 G.2.5.1 Software Coding Process Artifacts ................................................................. 316 G.2.5.1.1 System Safety Assessment – Software Fault Tree Analysis (SW FTA). ......... 316 G.2.5.1.2 Report, Record of Meeting/Minutes (RRM). .................................................... 316 G.2.5.1.3 Data Accession List (DAL). ............................................................................. 316 G.2.5.1.4 Software Test Description (STD). ................................................................... 316 G.2.5.2 Software Coding Process Requirements. ....................................................... 317 G.2.5.2.1 Software Code Requirements......................................................................... 317 G.2.5.2.2 Software Test Description (STD) Requirements. ............................................ 318 G.2.5.2.3 System Safety Assessment (SSA) – Fault Tree Analysis (SwFTA) Reqmts. .. 319 G.2.6 Software Integration Process. ........................................................................ 319 G.2.6.1 Software Integration Process Artifacts ............................................................ 319 G.2.6.1.1 Software Version Description (SVD). .............................................................. 319 G.2.6.1.2 Data Accession List (DAL). ............................................................................. 320 G.2.6.1.3 Report, Record of Meeting/Minute (RRM)....................................................... 320 G.2.6.2 Software Integration Process Requirements. .................................................. 320 G.2.6.2.1 Executable Code Requirements. .................................................................... 320 G.2.6.2.2 Software Version Description (SVD) Requirements. ....................................... 320 G.2.7 Software Test Verification Process. ................................................................ 320 G.2.7.1 Software Test Verification Process Artifacts. .................................................. 320 G.2.7.1.1 Software Test Report (STR). .......................................................................... 320 G.2.7.1.2 Data Accession List (DAL). ............................................................................. 321 G.2.7.1.3 Report, Record of Meeting/Minute (RRM)....................................................... 321 G.2.7.2 Software Test Verification Process Requirements. ......................................... 321 G.2.7.2.1 Software Test Report (STR) Requirements. ................................................... 321 G.2.8 Software Approval Process ............................................................................ 322 G.2.8.1 Software Approval Process Artifacts............................................................... 322 G.2.8.1.1 Software Accomplishment Summary (SAS). ................................................... 322 G.2.8.1.2 Software Product Specification (SPS). ........................................................... 322 G.2.8.1.3 Safety Assessment Report (SAR). ................................................................. 322 G.2.8.1.4 System Safety Assessment (SSA) – Final. ..................................................... 322 G.2.8.1.5 Data Accession List (DAL). ............................................................................. 322 G.2.8.1.6 Report, Record of Meeting/Minute (RRM)....................................................... 322 G.2.8.2 Software Approval Process Requirements. .................................................... 323

Appendix D

20

G.2.8.2.1 Software Accomplishment Summary (SAS) Requirements. ............................ 323 G.2.8.2.2 Safety Assessment Report (SAR) Requirements............................................ 323 G.3. Software Hazard Criticality Index (SHCI). ....................................................... 324 G.3.1 Software Hazard Criticality Index (SHCI) Definitions. ..................................... 324 G.3.2 SHCI Assignment Method. ............................................................................. 324 G.3.3 Guidelines for Determining SHCI Assignment Levels. .................................... 325 G.4. Software Lifecycle Process Requirements...................................................... 326 G.5. Software Change Procedures......................................................................... 330 G.5.1 Post Baseline. ................................................................................................ 330 G.5.2 Developmental Flight Test. ............................................................................. 330 G.6. Programmable Hardware Lifecycle Processes ............................................... 331 G.6.1 Programmable Hardware Planning Process ................................................... 331 G.6.1.1 Programmable Hardware Planning Process Artifacts .................................... 331 G.6.1.1.1 Plan for Hardware Aspects of Certification (PHAC) ........................................ 331 G.6.1.1.2 Hardware Design Plan (HDP) ......................................................................... 331 G.6.1.1.3 Hardware Validation Plan ............................................................................... 331 G.6.1.1.4 Hardware Verification Plan ............................................................................. 331 G.6.1.1.5 Hardware Configuration Management Plan (HCMP) ...................................... 332 G.6.1.1.6 Hardware Process Assurance Plan (HPAP) ................................................... 332 G.6.1.1.7 Hardware Design Standards .......................................................................... 332 G.6.1.1.7.1 Requirements Standards ................................................................................ 332 G.6.1.1.7.2 Hardware Design Standards .......................................................................... 332 G.6.1.1.7.3 Validation and Verification Standards ............................................................. 332 G.6.1.1.7.4 Hardware Archive Standards .......................................................................... 332 G.6.2 Programmable Hardware Design Processes .................................................. 332 G.6.2.1 Programmable Hardware Design Process Artifacts ........................................ 333 G.6.2.1.1 Hardware Requirements Document (HRD)..................................................... 333 G.6.2.1.2 Hardware Design Representation Data (HDD) ............................................... 333 G.6.2.1.2.1 HDD - Conceptual Design Data ...................................................................... 333 G.6.2.1.2.2 HDD – Detail Design Data .............................................................................. 333 G.6.3 Programmable Hardware Supporting Processes ............................................ 334 G.6.3.1 Programmable Hardware Supporting Processes Artifacts .............................. 334 G.6.3.1.1 Hardware Validation and Verification Data ..................................................... 334 G.6.3.1.1.1 Hardware Traceability Data ............................................................................ 334 G.6.3.1.1.2 Hardware Review and Analysis Procedures ................................................... 334 G.6.3.1.1.3 Hardware Review and Analysis Results ......................................................... 334 G.6.3.1.1.4 Hardware Test Procedures ............................................................................. 335 G.6.3.1.1.5 Hardware Test Results ................................................................................... 335 G.6.3.1.2 Hardware Acceptance Test Criteria ................................................................ 335 G.6.3.1.3 Problem Reports ............................................................................................ 335 G.6.3.1.4 Hardware Configuration Management Records .............................................. 335 G.6.3.1.5 Hardware Process Assurance Records .......................................................... 335 G.6.3.1.6 Hardware Accomplishment Summary (HAS) .................................................. 335 G.7. Testing Requirements and Terminology. ........................................................ 336 G.7.1 Robustness Test Case. .................................................................................. 336 G.7.2 Software Test Environment Validation. ........................................................... 337 G.7.3 Tool Qualification............................................................................................ 337 G.7.4 Model Based Development. ........................................................................... 337 G.7.5 Dead and Deactivated Code. ......................................................................... 338 G.7.6 Software Load Integrity................................................................................... 338

Appendix D

21

TABLES

TABLE IA. Performance 15°C (59°F), sea level static conditions with nozzle. TABLE IB. Performance 32°C (90°F) sea level static conditions with nozzle. TABLE IC. Performance 50°C (122°F) sea level static conditions with nozzle. TABLE IIA. Performance 2000 ft, 21°C (70°F) static conditions with _____ nozzle. TABLE IIB. Performance 4000 ft, 35°C (95°F) static conditions with ________ nozzle. TABLE IIC. Performance 6000 ft, 35°C (95°F) static conditions with ________ nozzle. TABLE IID. Performance 20,014 ft. 0.4 Mach number standard day with __________nozzle. TABLE III. Low Cycle Fatigue Damage Fraction Table. TABLE IV. Component test report documentation requirements. TABLE V. Gearbox pads and drives. TABLE VI. Engine Temperature Limits. TABLE VII. Not used. TABLE VIII. Not used. TABLE IX. Critical safety items low cycle fatigue lives. TABLE X. Fuel contamination. TABLE XI. Fuel contamination. TABLE XII. Data recording requirements. TABLE XIII. Sea level anti-icing conditions. TABLE XIV. Schedule of salt spray injection endurance cycles. TABLE XV. Electromagnetic compatibility test methods. TABLE XVI. Provisions for engine condition indication. TABLE XVII. EMS BIT code data. TABLE XVIII. Components requiring functional bench calibration. TABLE XIX. Engine split-line flange leakage. TABLE XX-A. Component tests for EIT. TABLE XX-B. Component tests for QT. TABLE XXI. Test plan format. TABLE XXII. Maintainability Requirements. TABLE XXIII. Turboshaft engine 4000 ft, 35°C (95°F) output shaft speed droop requirements. TABLE XXIV Parts classification. TABLE XXV Component Test Plan Documentation Requirements

Appendix D

22

FIGURES

FIGURE 1. Externally applied forces. FIGURE 2 Engine surface temperature versus engine length. FIGURE 3. Engine emissivity versus engine length FIGURE 4. Not used. FIGURE 5-1. Idle gas generator speed. FIGURE 5-2. Idle output shaft torque, ______ RPM. FIGURE 6. No load gas generator speed ______ RPM. FIGURE 7. Starting and Operating Envelope. FIGURE 8. Engine operating limits. FIGURE 9. Engine attitude limits. FIGURE 10-1. Estimated acceleration time versus altitude. FIGURE 10-2. Estimated acceleration time versus power. FIGURE 10-3A. Power absorber loading for acceleration; ground idle to MRP, SL to 10,000 ft. FIGURE 10-3B. Power absorber loading for acceleration; NL to MRP, SL to 10,000 ft. FIGURE 10-3C. Power absorber loading for deceleration; MRP to NL, SL to 10,000 ft. FIGURE 10-3D. Power absorber loading for acceleration; NL to MRP, 10,000 ft to maximum

altitude. FIGURE 10-3E. Power absorber loading for deceleration; MRP to NL, 10,000 ft to maximum

altitude. FIGURE 11. Continuous maximum icing conditions. FIGURE 12. Intermittent maximum icing conditions. FIGURE 13. Exhaust invisibility limit. FIGURE 14-1. Oxides of nitrogen. FIGURE 14-2. Gaseous emissions. FIGURE 15. Control lever angles versus control lever torques. FIGURE 16. Control limiter regimes. FIGURE 17. Starting torque and speed. FIGURE 18. Engine ground starting time versus ambient air temperature - static, no ram,

sea level to 20,000 ft. FIGURE 19. Engine altitude and starting test points. FIGURE 20. Not applicable. FIGURE 21. Not applicable. FIGURE 22. Not applicable. FIGURE 23. Engine corrosion operating cycle. FIGURE 24. Ambient temperature extremes versus altitude. FIGURE 25. Jet wake. FIGURE 26. Reference exhaust nozzle. FIGURE 27. Not used. FIGURE 28. Component surface limit temperatures. FIGURE 29. Engine vibration spectrum. FIGURE 30-1. Near field octave band sound pressure level contours (dB) center frequency -

250 Hz. FIGURE 30-2. Far field overall sound pressure level contours (dB). FIGURE 30-3. Estimated overall sound pressure level contours at idle.

Appendix D

23

APPENDICES Appendix A Requirements for EIT Appendix B Requirements for PFR Appendix C Requirements for QT Appendix D Requirements for OCR Appendix E Requirements for Strength, Life and Creep Analysis Report Appendix F Requirements for Vibration and Stress Report Appendix G Requirements for Software/Programmable Hardware Development

Appendix D

24

1. SCOPE.This airworthiness qualification plan (AQP) establishes the requirements and interfaceconfiguration definitions for the Improved Turbine Engine. It also establishes enginerequirements and their satisfactory completion and acceptance, by the Using Service for EngineIntegrity Testing (EIT), Preliminary Flight Rating (PFR), Qualification Test (QT) Rating , andOperational Capability Release (OCR). Further, this AQP identifies the tests, procedures, anddata required for satisfactory completion of production engine Acceptance Tests (AT). This AQPalso establishes the content and format to be used by the engine contractor for the preparationof the engine specification.

1.1 Metric Requirements. Not Applicable.

1.2 Deviations. Any projected design for a given application which will result in improved system performance, reduced life cycle cost, or reduced development cost through deviation from this AQP, or where the requirements of this AQP result in compromise in operational capability, will not be considered for approval unless brought to the attention of the Using Service in writing.

2. APPLICABLE DOCUMENTS

2.1 General. The documents listed in this section are specified in sections 3, 4, or 5 of this AQP. This section does not include documents cited in other sections of this AQP or recommended for additional information or as examples. While every effort has been made to ensure the completeness of this list, document users are cautioned that they must meet all specified requirements of documents cited in sections 3, 4, or 5 of this AQP, whether or not they are listed.

2.2 Government Documents.

2.2.1 Specifications, Standards, and Handbooks. The following specifications, standards, and handbooks form a part of this document to the extent specified herein. Unless otherwise specified, the issues of these documents are those cited in the solicitation or contract.

COMMERCIAL ITEM DESCRIPTIONS Federal A-A-52557AFED-STD-313D(1)

Fuel Oil, Diesel; for Posts, Camps and Stations Material Safety Data, Transportation Data and Disposal Data for Hazardous Materials Furnished to Government Activities

DEPARTMENT OF DEFENSE SPECIFICATIONS MIL-DTL-5624U NOT 1 Turbine Fuel, Aviation, Grades JP-4 and JP-5 MIL-PRF-7808L Lubricating Oil, Aircraft Turbine Engine, Synthetic Base, NATO

Code Number 0-148 MIL-PRF-23699F Lubricating Oil, Aircraft Turbine Engine, Synthetic Base, NATO

Code Number 0-156

Appendix D

25

MIL-DTL-38999L SUP 1 Connector, Electrical Circular, Miniature, High Density Quick Disconnect (Bayonet, Threaded and Breech Coupling), Environment Resistant, Removable Crimp and Hermetic Solder Contacts, General Specification for

MIL-PRF-7024F Calibrating Fluids, Aircraft Fuel System Components MIL-DTL-83133H

MIL-DTL-85470B NOT 1

MIL-L-85762A

MIL-PRF-85704C

Turbine Fuels, Aviation, Kerosene Types, JP-8 (NATO F-34), NATO F-35 and JP-8+100 (NATO F-37) Inhibitor, Icing, Fuel System, High Flash, NATO Code Number S-1745Lighting, Aircraft, Interior, Night Vision Imaging System (NVIS)CompatibleCleaning Compound, Turbine Engine Gas Path

MS3336 Accessory Drives, Aircraft Engine, Reference Chart ForMS3459D Connector, Plug, Electrical, Self-Locking, Coupling Nut, Rear

Release, Crimp Contact, An Type

DEPARTMENT OF DEFENSE STANDARDS MIL-STD-130N Identification Marking Of U.S. Military Property MIL-STD-461F Control of Electromagnetic Interference Emissions and

Susceptibility, Requirements for the MIL-STD-704A thru F Aircraft Electric Power Characteristics MIL-STD-810C and G Environmental Test Methods and Engineering Guidelines MIL-STD-882E System Safety MIL-STD-1553B NOT 4 Digital Time Division Command/Response Multiplex Data Bus MIL-HDBK-454B General Guidelines for Electronic Equipment MIL-HDBK-704-8 NOT 1 Guidance for Test Procedures for Demonstration of Utilization

Equipment Compliance to Aircraft Electrical Power Characteristics - 28 Vdc

MIL-HDBK-781A Handbook for Reliability Test Methods, Plans and Environments for Engineering, Development Qualification and Production

MIL-HDBK-1559(2) NOT 1 Numbers, Serial, Aircraft Gas Turbine Engine and Engine Module, Assignment of

MIL-HDBK-1812 Type Designation, Assignment and Method for Obtaining (Unless otherwise indicated, copies of the above specifications, standards, and handbooks are available online at https://assist.daps.dla.mil/quicksearch/ or http://dodssp.daps.dla.mil/ or from the Standardization Documents Order Desk, 700 Robbins Avenue, Building 4D, Philadelphia, PA 19111-5094, phone (215) 697-2667.)

2.2.2 Other Government Documents, Drawings, and Publications. The following other Government documents, drawings, and publications form a part of this document to the extent specified herein. Unless otherwise specified, the issues of these documents are those cited in the solicitation or contract.

Department of Defense DODD 5000.01 The Defense Acquisition System DODI 5000.02 Operation of the Defense Acquisition System

26

Appendix D

(These documents are available online at http://www.dtic.mil/whs/directives/. Application for paper copies should be addressed to DAPS, ATTN: DODSSP, Building 4D, 700 Robbins Avenue, Philadelphia PA 19111-5094) Department of Transportation FAA Advisory Circular No: 20-135

FAA Advisory Circular No: 33-28-1

Powerplant Installation and Propulsion System Component Fire Protection Test Methods, Standards, and Criteria Compliance Criteria for 14 CFR §33.28, Aircraft Engines, Electrical and Electronic Engine Control Systems

(Application for copies should be addressed to Department of Transportation, FAA Aeronautical Center, AML-611, P O Box 25082, Oklahoma City OK 73125-5082.) Naval Air System Command NAVAIR 01-1A-509 and Air Force TO 1-1-689

Avionics Cleaning and Corrosion Prevention/Control Technical Manual

(Application for copies should be addressed to Naval Publications & Forms Directorate, Attn: Code 301, 5801 Tabor Ave., Philadelphia PA 19120-5099) United States Army ADS-9C Propulsion System Technical Data ADS-37A-PRF Electromagnetic Environmental Effects (E3)

Performance And Verification Requirements

(Application for copies should be addressed to USA CCDC Aviation & Missile Center, Aviation Engineering Directorate, FCDD-AE, Redstone Arsenal, AL 35898.)

2.3 Nongovernment Publications. The following documents form a part of this document to the extent specified herein. Unless otherwise specified, the issues of these documents are those cited in the solicitation or contract.

American Society for Testing and Materials (ASTM) ASTM D5001-10 Standard Test Method for Measurement of Lubricity of Aviation

Turbine Fuels by the Ball-on-Cylinder Lubricity Evaluator (BOCLE)

ASTM D 910-11 Stand Specification for Aviation Gasolines

(Application for copies should be addressed to the American Society for Testing and Materials, 100 Barr Harbor Drive, West Conshohocken PA 428-2959.) Institute of Environmental Sciences

IEST-RP-PR001.1 Management And Technical Guidelines For The ESS Process

(Application for copies should be addressed to Institute of Environmental Sciences and Technology, Arlington Place One, 2340 South Arlington Heights Road, Suite 100, Arlington Heights, IL 60005-4516; www.iest.org.)

Appendix D

27

International Electrotechnical Commission (IEC) IEC 61672-1 IEC 61672-2

Electroacoustics - Sound level meters - Part 1: Specifications Electroacoustics - Sound level meters - Part 2: Pattern evaluation tests

IEC 61260 Electroacoustics - Octave-band and fractional-octave-band filters

(Application for copies should be addressed to American National Standards Institute, 11 West 42nd Street, New York NY 10036.) International Organization for Standardization ISO 12103-1:1997 Road vehicles – Test Dust for Filter Evaluation-Part 1: Arizona

Test Dust First Edition (Application for copies should be addressed to International Organization for Standardization, Case Postale 56, CH-1211, Geneva 20, Switzerland; www.iso.org.) Radio Technical Commission for Aeronautics, (RTCA), Inc. DO-160G Environmental Conditions and Test Procedures for Airborne

Equipment DO-178C

DO-254

Software Considerations in Airborne Systems and Equipment Certification

Design Assurance Guidance for Airborne Electronic Hardware DO-331 Model-Based Development And Verification Supplement

(Applications for copies should be addressed to RTCA, Inc. 1150 18th Street; Suite 910, Washington, DC 20036; www.rtca.org) Society of Automotive Engineers (SAE) International Aerospace Material Specifications AMS 2518C Thread Compound, Anti-size, Graphite-Petrolatum Aerospace Recommended Practice ARP 492C Aircraft Engine Fuel Pump Cavitation Endurance Test ARP 704A Helicopter Engine - Rotor System Compatibility ARP 994 Recommended Practice for the Design of Tubing Installation for

Aerospace Fluid Power Systems ARP 1179D Aircraft Gas Turbine Engine Exhaust Smoke Measurement ARP 1256D Continuous Sampling and Measurement of Gaseous Emissions

from Aircraft Turbine Engines, Procedures for ARP 1420B Gas Turbine Engine Inlet Flow Distortion Guidelines ARP 1533A Procedure for the Calculation of Gaseous Emissions from

Aircraft Turbine Engines ARP 1587B Aircraft Gas Turbine Engine Monitoring System Guide ARP 1797A Aircraft and Aircraft Engine Fuel Pump Low Lubricity Fluid

Endurance Test ARP 4671

ARP 4754A

Guidelines and Methods for Conducting the Safety Assessment Process on Civil Airborne Systems and Equipment Guidelines for Development of Civil Aircraft and Systems

ARP 5580 Recommended Failure Modes and Effects Analysis (FMEA) Practices for Non-Automobile Applications

ARP 5757 Guidelines for Engine Component Tests

Appendix D

28

Aerospace Resource Document ARD50015 A Current Assessment of the Inlet/Engine Temperature Distortion

Problem Aerospace Standard AS 568C Aerospace Size Standard for O-Ring Sizes AS 681J Gas Turbine Engine Steady-State Performance Presentation

for Digital Computer Program AS 755D Aircraft Propulsion System Performance Station Designation

and Nomenclature AS 870D Wrenching Configuration, Double Hexagon (12 points), for

Threaded Fasteners AS 972B AS 1055D

Spline Details, Accessory Drives and Flanges Fire Testing of Flexible Hose, Tube Assemblies, Coils, Fittings and Similar System Components

AS 4273A Fire Testing of Fluid Handling Components for Aircraft Engines AS 4395B Fitting End-Flared Tube Connection, Design Standard

Aerospace Information Report AIR 1419A Inlet Total-Pressure-Distortion Considerations for Gas-Turbine

Engines (Application for copies should be addressed to the SAE World Headquarters, 400 Commonwealth Drive, Warrendale, PA 15096-0001; www.sae.org.)

American Society of Mechanical Engineers (ASME) Performance Test Codes PTC19.1-2005 Test Uncertainty

(Application for copies should be addressed to the American Society of Mechanical Engineers, PO Box 2300, , Fairfield, NJ, 07007-2300; www.asme.org.).) North Atlantic Treaty Organization (NATO) NATO Code F-34, 40, 44 and 54 NATO Logistics Handbook, Aviation Fuels

2.4 Order of precedence. Unless otherwise noted herein or in the contract, in the event of a conflict between the text of this document and the references cited herein (except for related specification sheets), the text of this document takes precedence. Nothing in this document, however, supersedes applicable laws and regulations unless a specific exemption has been obtained.

Appendix D

29

3. REQUIREMENTS.The paragraphs of this AQP and all references herein comprise the complete set of enginerequirements. All requirements that are classified shall be provided in a classified appendix, aspart of this AQP. Terms, symbols, acronyms, and definitions shall be in accordance with Section6.

3.1 Item Definition. A brief description of the salient features of the engine shall be provided in the engine specification. This shall include, if applicable, a description of the Inlet Particle Separator (IPS); compressor(s) including number of stages, variable geometry provisions, acceleration/customer bleeds; combustor type, including method of all injection; turbine components, including number of stages, cooling provisions; exhaust nozzle; engine control features; prognostics and health monitoring systems; necessary gearbox provisions; type of lubrication and scavenge system; starting ignition system; number and location of main bearings; instrumentation and performance indicating provisions. The performance ratings shall be as specified in Tables IA, IB, IC, IIA, IIB, IIC and IID. Terms, symbols, and their definitions shall be in accordance with (IAW) sections 6 and 7 and with AS 755D. The performance is based upon a referee nozzle which shall be provided by the engine contractor for development and acceptance engine testing only. The engine shall include an emergency power capability as defined in 6.3.

3.1.1 Item Diagram. Item diagrams for lubrication system, cooling and sealing air system, IPS, anti-icing system, fuel system, electrical system, and engine control system shall be included in the model specification.

3.1.2 Interface Definition The interface requirements include all physical installation and performance requirements necessary for engine installation. All interface definitions shall be as shown on the Engine Configuration and Envelope Computer-Aided Design (CAD) models or in the text describing the applicable functional system.

3.1.2.1 Drawings. The following drawings, as figures, shall form a part of the engine specification. Reduced size copies of these figures shall be included in this AQP. These drawings shall be updated as required to reflect the latest engine configuration

a. Engine Configuration and Envelope Figure. This drawing or CAD model shall bedetailed profiles in all planes to show and identify the physical inner features of the engine. The drawing shall show mounting details and tolerances for the engine and all installation items, clearances for installation and removal of accessories and components subject to separate removal, access for adjustments and other maintenance functions, and center of gravity of the complete bare engine. This drawing shall show the maximum space required by the engine including tolerances and dimensional changes due to manufacturing, thermal effects, vibration, and operating and externally applied loads. Vibration and temperature measurement locations shall be specified on the drawing.

b. Electrical Installation Connection Figure. This drawing shall show and identify allengine systems external electrical circuit requirements and installation interface connection details. If an Electrical Interface Control Document (EICD) is provided as a supplement, it shall be referenced on the connection figure.

30

Appendix D

3.1.2.2 Not used.

3.1.2.3 Installation Interfaces. Engine features affecting engine installation interfaces shall be identified and defined in the engine specification.

3.1.2.4 Moments of Inertia. The engine specification shall specify:

a. The maximum effective mass moment of inertia (slug-feet squared) of the complete dry engine about three mutually perpendicular axes with the origin at the center of gravity.

b. The maximum effective mass moment of inertia (slug-feet squared) of each engine rotor system about the resultant rotational axis, together with the effective direction of rotation of the inertia and the direction and location of the resultant rotational axis. For engines with geared rotor systems, the shaft to which all inertias of each rotor system have been algebraically referred shall be specified.

c. The maximum effective mass moment of inertia (slug-feet squared) of the complete power output system (including the reduction gear train) referred to the output shaft speed.

d. The maximum effective mass moment of inertia (slug-feet squared) of the rotating parts which are cranked by the engine starter.

3.1.2.5 Externally Applied Forces. The engine shall function satisfactorily under the conditions specified in 3.1.2.6 and Figure 1 (combined linear and angular effects) and shall meet the requirements of this AQP during and after exposure to those conditions without permanent deformation and shall operate satisfactorily thereafter. The engine shall also not fail catastrophically when subjected to static loads equivalent to 1.5 times those values but the engine need not operate satisfactorily thereafter. The engine shall not fail catastrophically when subjected to the crash load factors identified below:

Direction Applied Separately

Applied Simultaneously

Case 1 Case 2 Case 3

Axial +/-20 +/-20 +/-10 +/-10

Lateral +/-18 +/-9 +/-9 +/-18

Vertical +10/-20 +5/-10 +10/-20 +5/-10

NOTE: Values are in g’s. For the vertical direction, up is positive (+). Down loads (negative vertical load factors) occur during pullout. The limit loads shall be based on a mass factor consisting of the dry mass of the engine, increased by the specific mass allowed for all engine mounted accessories and operating fluids. If airframe components are supported by the engine, the mass of these components shall be included in the mass factor. Externally applied forces include: loads produced by takeoff, landing, in-flight maneuvers, gusts, vibration, installation, and crash conditions.

31

Appendix D

For control and external components, the contractor shall ensure the components can meet the requirements of this AQP during and after exposure to the external forces induced by the installation, based on the anticipated usage and handling. MIL-STD-810C and G shall be used to establish the design and performance requirements. An analysis shall be conducted of the entire engine with regard to its capability to withstand the loads specified in this paragraph, 3.1.2.5.1, and 3.1.2.6.1 and 3.1.2.14.1.

3.1.2.5.1 Gyroscopic Moments. At all speeds up to at least maximum allowable steady-state rotor speeds, the engine shall operate satisfactorily when exposed to gyroscopic moments defined by the operating conditions in paragraphs a through d. Loads shall include the worst case impact of all engine mounted airframe hardware. Engine mounted airframe hardware includes, but shall not be limited to, the inlet and exhaust components supported by the engine.

a. A steady angular velocity of 2.5 radians per second in any axis in a planeperpendicular to the rotor axis, combined with a vertical load factor of either +1g or -1g for a total time of 30 seconds. The limit of occurrences at this condition shall be two (2).

b. A steady angular velocity of 0.4 radians per second in any axis in a planeperpendicular to the rotor axis for 24% of duty cycle life.

c. A steady angular velocity of 0.9 radians per second in any axis in a planeperpendicular to the rotor axis for 20% of duty cycle life.

d. A steady angular velocity of 1.4 radians per second in any axis in a planeperpendicular to the rotor axis for 1% of duty cycle life. The definition of “the engine shall operate satisfactorily” is as follows:

a. Loss of engine power shall not prevent safe flight and return to base withoutexceeding of any engine operating or physical limits.

b. Engine operability margins are sufficient to avoid engine stall with normalacceleration and deceleration rates.

c. Damage to engine components is limited to tip and seal rubs.d. Remaining part life is not reduced relative to life prior to exposure to the gyroscopic

induced load

3.1.2.6 Mounts.

3.1.2.6.1 Main Mounts. The engine mounts shall retain the engine, including retained fluids and externals, at all flight, takeoff and landing, and ground conditions and shall meet the requirements of this AQP during and after exposure to limit loads of (contractor to insert) without permanent deformation and ultimate loads of (contractor to insert) without fracture. The main mounts shall meet the damage tolerance requirements of 3.3.8.11. Mounting provisions shall be of sufficient strength to prevent loss of engine retention when subject to a single attachment point failure at any location at the end of the engine mount service life and able to handle limit Figure 1 loads. The locations and descriptions of all engine mounts shall be as specified in the Engine Configuration and Envelope Figure. The engine mounts shall be consistent with the mounting approach in the UH-60 and AH-64 installations.

32

Appendix D

The maximum system limits shall be specified in units of force and in reference to the engine. The specified values shall include, but not be limited to, the effects of the following requirements and specific design characteristics: externally applied forces (e.g., accelerations) of 3.1.2.5, gyroscopic moments of 3.1.2.5.1, all air vehicle loads which are supported through the engine structure (if such loads exist) and safety factors of 3.3.8.4.1, cyclic fatigue, engine mass, material strength and mechanics and service life. The bending moment limits shall be specified in the axial, vertical, and lateral directions. The engine mounts shall meet the applicable requirements of 3.3.8 and subparagraphs.

3.1.2.6.2 Ground Handling Mounts. The ground handling mounts shall support the engine, including all engine mounted equipment and externals, components, and operating fluids, under the following maximum inertia load conditions, without deformation to the mounts or damage to the engine: 4g axial, 2g lateral, and 3g vertical acting in combination at the engine center of gravity. The locations and descriptions for the individual ground handling mounts shall be as specified in the Engine Configuration and Envelope Figure. The arrangement shall be compatible with Maintenance Trailer, National Stock Number 1730-01-026-1653, Data List 4920-EG-081.

3.1.2.6.3 Engine Stiffness and Modes. The estimated stiffness of the engine in resisting loads and moments applied at the outboard end of the output shaft, relative to the engine mounting points, shall be stated herein. Translational stiffnesses in the three orthogonal directions shall be presented relative to the engine mount point(s) where that direction is constrained. Rotational stiffnesses in the three orthogonal directions shall be presented with the engine mount points constrained as designed. The first lateral and vertical engine bending modes shall be specified with the mount points constrained as designed. The loads shall include, but not be limited to, the effects of externally applied forces and gyroscopic moments.

3.1.2.7 Pads and Drives. Pads and drives suitable for mounting and driving accessories shall be IAW the basic configuration and rating requirements specified in the engine specification and presented as shown in Table V. The engine component drive system and accessory drive system shall be capable of simultaneous operation of all the drives when each drive is subjected to 1.25 times the maximum permissible torque or power rating specified for the individual drive. All drive splines shall be positively lubricated by engine oil. Complete dimensions and details of the drive pads including the clearance envelopes and alignment requirements shall be shown on the Engine Configuration and Envelope Figure. No part of the gearbox shall prevent independent removal of any one accessory mounted on these drives . Pads and drives for accessories shall conform to the standards listed on MS 3336, except compliance with the military standard tabulated data need not be maintained.

3.1.2.8 Engine Surface Temperature and Heat Rejection. The maximum operating surface temperatures of the engine and heat rejection rates shall be specified as shown in Figure 2. Factors such as engine bay cooling flows, accessory pad loading, compressor bleed air conditions, anti-ice operation, oil system cooling requirements, engine skin temperature and radiation properties, and air and gas leakage from engine case flanges and split lines shall be taken into consideration in establishing engine heat rejection rates. The conditions surrounding the engine and the engine power condition shall be as stated in items a, b, c and d. The oil system temperatures for which the surface temperatures are applicable shall be provided in an appendix. For components and accessories on the surface of the engine, the specified

33

Appendix D

component and accessory surface temperature and heat generation curves may differ from the engine temperature as shown by the dashed lines on Figure 2. The engine surface emissivity shall be as presented on Figure 3. The heat rejection and surface temperature data shall be for the conditions listed below. The engine case and split line flange leakage(s) shall be as shown on Table XIX.

a. Maximum power, sea level, 55°C (131°F) ambient, static.b. Maximum power, 4000 ft, 45°C (113°F) ambient, maximum stagnation inlet air

temperature. c. Maximum power, absolute altitude, hot atmosphere, maximum stagnation inlet air

temperature.d. Contingency power, sea level static, 55°C (131°F) ambient.

3.1.2.8.1 Engine Component Limiting Temperature.Engine components when mounted on the engine shall not exceed their allowable temperature under the following conditions:

a. Continuous operation with ambient air at the maximum stagnation temperature or80°C (176°F), whichever is warmer.

b. Flight shutdown from the most adverse condition and continued soaking withambient air at maximum stagnation temperature or 80°C (176°F) whichever is warmer.

c. Ground shutdown with ambient air at 55°C (131°F) with no special cooling suchas forced ventilation, refrigeration, or rotation of rotors. Off-engine mounted components shall not exceed their component limiting temperature as shown on Figure 28 when surrounded by air at 90°C (194°F). No additional cooling beyond natural convection shall be required for operation at or below this temperature. A tabulation of the maximum limiting surface temperature for all components shall be as shown on Figure 28. The specific points of measurement, where the ambient air temperature and surface temperature with respect to the three coordinate axes of the component are maximum, shall be defined prior to the completion of EIT.

3.1.2.8.2 Heat Rejection and Cooling Analysis. A heat rejection and cooling analysis shall be conducted to verify the heat rejection and surface temperature data presented in 3.1.2.8. Engine test data shall be used to substantiate the analysis. An engine heat rejection and cooling schematic diagram showing component and station locations, including temperature limits, shall be included in the engine specification..

3.1.2.9 Air and Gas Leakage. The engine shall be designed to minimize air and gas leakage. There shall be no locations where leakage flow will be of sufficient temperature and concentrated impingement to present a safety hazard or affect installation requirements.

Appendix D

34

3.1.2.10 Engine Air Inlet System.

3.1.2.10.1 Air Inlet Design and Dimensions. The engine shall have an inlet attachment flange that shall be designed to accept a quick-disconnect clamp. Interface dimensions for the attachment of the inlet duct shall be as shown on the Engine Configuration and Envelope Figure.

3.1.2.10.2 Allowable Inlet Connection Loads. The maximum allowable shear, axial, and overhung moment loads at the engine inlet flange shall be specified in the engine specification for static (1g) conditions. The allowable loads at the engine inlet flange shall also be specified for the maximum allowable maneuver loads as defined in Figure 1 (combined linear and angular effects).

Static (1g) Maximum allowable Maneuver loads

Shear TBD TBD Axial TBD TBD Overhung Moment TBD TBD

Crash Load Factor 25 As a starting point the minimum static (1g) capability shall be as follows:

Axial Load 35 pounds Shear Load 35 pounds Overhung Moment 40 pound-feet

3.1.2.10.3 Inlet Airflow Distortion Limits. The engine shall not surge, stall, flameout, or incur any damage with the steady state or time variant inlet distortion (pressure, temperature, or any combination of both) shown herein (contractor to provide). The aerodynamic interface plane shall be identified on the Engine Configuration and Envelope Figure. Where exhaust nozzle backpressure effects on the engine affect tolerance of the engine to inlet air pressure variation, the effect shall be specified in the engine system specification. Engine stability and performance assessments shall use the methodology and inlet distortion descriptors defined in ARP 1420B and AIR 1419A for total pressure and temperature distortion. Concurrent inlet pressure and temperature distortion methodology shall be specified herein. Engine distortion limits at which the engine shall operate without stall, rotating stall, and surge throughout the complete environmental and operating envelope shall be defined in the qualification altitude test and shall be included as an appendix in the engine specification once they are defined. An engine stability and performance assessment for total pressure and total temperature distortion shall be performed using the methodology and inlet distortion descriptors defined below:

a. A distortion descriptor methodology, based on the guidance of ARP 1420B,indicating the maximum distortion limits at which the engine shall operate without stall, rotating stall, and surge, throughout the complete environmental and operating envelope, due to steady-state and time-variant inlet air total pressure and temperature variation, shall be defined in an appendix in the engine specification.

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Appendix D

b. Using the distortion descriptor methodology, approximately five sets of pressure distortion patterns (screens) shall be provided, that, with Using Service approval, reflect the most severe conditions permitted within the limits of the distortion descriptor specified for the engine. Approximately three of these distortion patterns (screens), exceeding the most severe inlet spatial and planar pressure conditions expected in the weapon system installation, shall be used for QT testing. All distortion patterns (screens) shall be used during QT and OCR engine testing. One of the patterns (screens), if required by the Using Service, shall be used for the Acceptance Test surge and stall margin test. For each set of inlet flow distortion data specified, the engine interface operating conditions shall be defined, as applicable, in terms of Mach number, altitude, power setting, customer bleed air, customer power extraction, etc. Measurements of the engine inlet total pressure, temperature, and flow variation shall be made at the engine inlet (aerodynamic interface plane). All inlet instrumentation utilized in measuring airflow, pressures and temperatures, the arrangement, location, response, and instrumentation accuracies shall be consistent with the guidelines of ARP 1420B and defined in the specification. For each set of specified inlet flow distortion data, the total airflow, average total pressure recovery and pressure and temperature for each individual probe shall be specified. The available surge margin, maximum distortion limits, and the differences between the contractor’s methodology and ARP 1420B shall be presented in the engine specification.

3.1.2.10.4 Pressure and Temperature Rate of Change. The engine shall withstand the following rate-of-change of pressure and temperature, separately or in combination, without stall, rotating stall, surge, flameout, or mechanical damage:

Parameters Rate-of-change Duration

Engine inlet pressure 7 psi/sec 0.4 sec

Engine inlet temperature 1800°F/sec 0.15 sec

3.1.2.11 Customer Bleed Air System. The engine shall provide for customer bleed air extraction for aircraft use from the compressor. The locations and interface dimensions of all customer bleed air ports shall be shown on the Engine Configuration and Envelope Figure. During customer bleed air extraction, control lever modulation shall not be required to maintain engine stability and limits within the environmental conditions and operating envelope of the engine. The pressure and temperature of customer bleed air extracted from idle to maximum power at all operating altitudes, air inlet temperatures, and flight speeds, and the effects upon engine performance when bleeding this air from the engine from any number of ports, shall be included in the performance computer programs of 3.2.1.2 and 3.2.1.2.2. The maximum bleed air temperature and pressure and the stage from which it is extracted shall be specified in the engine specification. The maximum continuous flow capability of each bleed air port in percent of total airflow shall be specified in the engine specification. All bleed air ports shall be sized to prevent engine failure in event of a single failure in the aircraft bleed system. Bleed ducts with external surfaces exceeding 370°C (698°F) shall be insulated to prevent hazards from combustible fluid leakage. The engine provided customer bleed air extraction system shall insure that no upstream malfunction of the engine will cause specified contamination limits to be exceeded. The bleed air energy required shall be defined in the RFP.

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Appendix D

3.1.2.11.1 Allowable Customer Bleed Connection Loads. The maximum allowable axial, shear, and overhung moment loads on the customer bleed air connections shall be specified in the engine specification for static (1g) conditions. The maximum allowable loads at these connections shall also be specified for the maximum allowable maneuver loads as defined in Figure 1 (combined linear and angular effects).

3.1.2.11.2 Start and Acceleration Bleed Air. Where acceleration bleed airflow is necessary, the airflow conditions for which venting provisions must be made shall be specified in the engine specification. The ducting attachment detail shall be shown on the Engine Configuration and Envelope Figure. Compressor bleed air required for compressor surge protection, which operates continuously during steady state engine operation in a surge sensitive regime, shall be defined in the engine specification as to the operating envelope involved.

3.1.2.11.3 Bleed Air Contamination. Engine generated substances contained in the customer bleed air shall be no greater than the threshold limit values specified below. The engine manufacturer shall demonstrate, by analyzing bleed air samples, that the specified threshold limits for the substances are not exceeded. Where substances other than those listed are contributed to the extracted air by engine operation, the engine manufacturer shall report the substances and the contamination in parts per million to the Using Service for determination of maximum limits. When two or more engine generated substances are present, their combined effect shall be determined and reported. In the absence of information to the contrary, the combined effects of the different substances shall be considered as additive. If cleaning fluids are specified for use during normal engine maintenance, consideration should be given to the effect on bleed air contamination.

Substance Parts per million Carbon dioxide 1000.0

Carbon monoxide 10.0 Ethanol 1000.0

Fluorine (as HF) 0.1 Hydrogen peroxide 1.0

Aviation fuels 25.0 Methyl alcohol 200.0 Methyl bromide 1.0 Nitrogen oxides 3.0

Acrolein 0.1 Oil breakdown products

(e.g., aldehydes) 1.0

Ozone 0.05

The bleed air should not contain a total of more than 5 milligrams per cubic meter of engine generated particles.

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Appendix D

3.1.2.12 Not used.

3.1.2.13 Connections. The engine shall be permanently marked to indicate all instrumentation, fuel, oil, and electrical connections. The connections shall be marked with interconnection diagram reference (e.g. J1, P2, etc.) and notated function (e.g. N2, Fuel, etc.) in a distinctive manner as a visual aid for maintenance and identification. Markings shall be of appropriate military or commercial standards of the vendor’s choice for human factors considerations and the environmental conditions described herein. Connections located in close proximity to each other, which are not functionally interchangeable, shall be made physically non-interchangeable.

3.1.2.14 Shaft Power Absorber. The shaft power absorber shall include any mechanism which absorbs the output shaft torque. The engine manufacturer shall be responsible for the satisfactory operation of the complete engine with its controls, when operated with the shaft power absorber and its controls having the characteristics as specified in 3.1.2.14.1.

3.1.2.14.1 Power Absorber to Engine Interface Characteristics. The allowable range of characteristics of the shaft power absorber at the power absorber to engine interface shall be as shown herein. No resonant frequency shall be transmitted to or from the power absorber through the engine. The output drive shaft shall be capable of continuously absorbing thrust in either direction of an amount not less than 20 percent of the circumferential force acting on the output drive shaft spline at the steady state torque limit as specified in the engine specification. The engine contractor shall provide characteristics including but not limited to: maximum and minimum polar moment of inertia, torsional spring constant, torsional damping coefficient as a function of torsional spring constant, the maximum allowable static and dynamic loads on the engine output drive shaft, direction of rotation of the output drive shaft as viewed from the engine inlet, design of the output drive shaft spline and the power absorber shaft maximum misalignment during steady state and transient operation.

a. The maximum allowable polar moment of inertia of the power absorber referred to engine output shaft speed is 1.2 slug-ft squared.

b. The minimum allowable polar moment of inertia of the power absorber referred to engine output shaft speed is 0.2 slug-ft squared.

c. The maximum allowable torsional spring constant of the power absorber drive system referred to engine output shaft speed is 1500 lbf-ft per radian.

d. The minimum allowable torsional spring constant of the power absorber drive system referred to engine output shaft speed is 20 lbf-ft per radian.

e. Specific rotor system definition or characteristics shall be coordinated with the airframe manufacturer in accordance with ARP 704A as required by the contract.

f. The allowable range of torsional damping coefficient as a function of torsional spring constant.

g. The maximum allowable misalignment of the power absorber drive shaft with respect to the centerline established by the pitch diameter of the engine output drive spline shall be specified on the engine configuration and envelope figure.

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Appendix D

h. The minimum allowable loads on the engine output drive shaft shall be; static loads 5 pounds; minimum allowable dynamic radial load (resulting from power absorber input shaft imbalance) shall be 50 pounds; minimum allowable overhung moment shall be 20 pound-inches (static or dynamic).

i. The minimum allowable static loads that may be imposed by mounting a nose gearbox on the optional main front mount shall be; 100 pounds shear and 82 pound-feet overhung moment.

j. The power absorber shaft configuration must provide a leak free cavity to permit engine supplied spline lubrication to return to the engine system.

k. The power absorber shaft maximum misalignment shall not exceed 0.25 degrees during steady-state operation and shall not exceed 0.50 degrees during power transients with respect to the centerline established by the pitch diameter of the engine output drive spline.

3.1.2.14.2 Not used.

3.1.2.14.3 Not used.

3.1.2.14.4 Turboshaft Engine Output Drives. The output shaft shall have an internal spline conforming to ANSI B92.2M unless otherwise specified which shall be positively lubricated by engine oil. No provisions shall be made to supply oil to the power absorber. The output shaft pad and the output shaft are shown on the Engine Configuration and Envelope Figure. The drawing also contains the following data:

a. The pitch diameter of the output shaft spline.b. The rated output shaft speed.c. The maximum radial load and the maximum overhung moment capabilities of the

output shaft. d. The maximum radial load and the maximum overhung moment capabilities of the

output shaft drive pad.e. The maximum allowable angular misalignment of the power absorber drive shaft

with respect to the centerline established by the pitch diameter of the engine output shaft spline.f. Any other output shaft pad and output shaft limitations.g. The direction of rotation as viewed from the aft looking forward.

3.1.3 Major Component List. Each component or functional subsystem of the engine which requires testing for engine integrity testing or qualification (see sections 4.5 and 4.6) shall be identified along with the specific tests required for engine integrity testing and for qualification in a tabular format in accordance with Tables XXA and XXB.

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Appendix D

3.1.4 Reliability.

3.1.4.1 Reliability Quantitative Requirements. The engine shall achieve the reliability requirements when operated within the requirements specified herein, and maintained in accordance with approved maintenance manuals. Damage to equipment attributable to battle damage or operation of equipment outside of prescribed limits shall be excluded.

3.1.4.1 Mean Engine Hours Between Mission Affecting Failure (MEHBMAF) The engine shall have a MEHBMAF of not less than 330 engine operating hours.

3.1.4.2 Mean Engine Hours Between Essential Maintenance Actions (MEHBEMA) The engine shall have a MEHBEMA of not less than 200 engine operating hours.

3.1.4.3 Mean Engine Hours Between Unscheduled Maintenance Actions (MEHBUMA) The engine shall have a MEHBUMA of not less than 140 engine operating hours.

3.1.4.4 Mean Engine Time Between Engine Removal (METBER) The engine shall have a METBER of not less than 1800 engine operating hours.

3.1.4.5 Direct Maintenance Man Hour per Engine Operating Hour (DMMH/EOH) The engine shall have a DMMH/EOP of less than 0.03.

3.1.5 Maintainability. The engine shall be designed for ease of servicing and maintenance. The engine Mean Time To Repair (MTTR), including all corrective maintenance tasks required to affect repair, shall not be greater than 1.50 hours. Corrective maintenance tasks included to compute MTTR shall include, but not be limited to:

a. On-aircraft troubleshooting, including the performance of applicable Built-In-Test (BIT)to detect and isolate the fault,

b. On-aircraft repair or removal and replacement of the engine system component(s), asapplicable,

c. On-aircraft checkout of the repair, including the performance of BIT to verify fault hasbeen corrected.

The engine shall be designed so that no more than two maintenance personnel are required to perform any corrective maintenance task at the unit/field maintenance level. The engine shall be designed to be maintained by at least the central ninety percent (90 percent) of the maintainer population (5th percentile female stature through 95th percentile male stature). The engine design shall meet the requirements specified in Table XXII.

3.1.5.1 Peculiar Support Equipment (PSE). The contractor shall identify and design all necessary items of PSE required to support, disassemble, assemble, inspect, repair, overhaul and test the engine (including all components and accessories) at all levels of maintenance. The engine shall not require peculiar equipment or special tools at the user level.

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Appendix D

3.1.6 Engine Monitoring System (EMS) The engine shall incorporate equipment or provisions for sensors as required to monitor engine performance and mechanical condition. Sufficient engine parameters shall be made available to detect early failure systems, performance degradation, permit fault isolation and provide long-term failure trending information. The EMS as defined herein shall provide, monitor, coordinate and record appropriate parameters for maintenance, performance trending, parameter limit exceedance and life usage. The EMS shall provide automatic in-flight data collection, engine diagnostic analysis to aid in the human decision process and provide for full parametric data retrieval by ground personnel. EMS data retrieval/diagnostics system(s) shall be used to interrogate and download engine parametric data for additional processing to build engine history files and provide the basis for the overall engine maintenance management system. As a minimum, the system shall provide the following features:

a. Engine and component life usage trackingb. Fault detection/isolation to module/LRU levelc. Performance monitoring/trendingd. Engine limit exceedances recording and maintenance notificatione. Oil system and fluid condition monitoring

The malfunction of any EMS hardware or software shall not affect engine performance or operability. Failure of the EMS system shall not cause a loss of more than the current flight of data. The engine monitoring system shall be completely functional following restoration from a single failure of any other subsystem of the engine, including engine control and electrical systems. EMS component lifing algorithms shall be approved by the Using Service as a requirement for EMS Qualification, per paragraph 4.6.2.7 and shall incorporate sufficient resolution to accurately define component life usage throughout the operational envelope and at the maximum transient capabilities of the engine. EMS validation shall require analysis of engine data acquired during engine qualification tests. This engine testing shall require engine operation at multiple rating points & various gas generator and power turbine speeds. EMS performance monitoring the algorithms shall also be approved by the Using Service as a requirement for EMS performance monitoring qualification, per paragraph 4.6.2.7 and must incorporate sufficient resolution to accurately define degradation of the engine on wing and determine the maximum power capability of the engine at any temperature/altitude combination.

3.1.6.1 EMS Software Development. EMS Software shall be designed, developed and tested in accordance with RTCA/DO-178C, level A. Additionally, the contractor shall provide documentation of the system and software design, requirements flowdown and traceability, and verification and validation in accordance with RTCA/DO-178C, level A. The design methodology shall be approved by the Using Service.

3.1.6.2 EMS Functionality. The EMS capability shall be fully described herein. The failure of the onboard engine diagnostics system shall not cause a failure of any other mission or safety critical system.

3.1.6.3 EMS Required Data. The specific minimum data which the EMS is required to record and process shall be defined by the contractor and provided to the Using Service for approval.

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Appendix D

3.1.6.4 LRU Fault Isolation. Built-in-Test (BIT) shall be implemented to accomplish failure detection and failure isolation to the LRU without the assistance of support equipment at the unit level.

3.2 Characteristics.

3.2.1 Performance Characteristics. Unless otherwise specified, the minimum engine performance characteristics (the poorest performing engine that would be submitted for acceptance) shall be based on:

a. A fuel having a lower heating value of 18300 BTU/lb and otherwise conforming to MIL-DTL-5624U fuel and oil specified in 3.7.7.2.1.

b. US Standard Atmosphere, 1976 (geopotential altitude).c. No inlet air distortion.d. An inlet pressure recovery of 100 percent.e. The referee exhaust pipe.f. No accessory power extraction, other than that required for continuous

engine operation. g. Specified performance predicated on the tolerance of the control system which

produces the poorest performance. h. A shaft power absorber with characteristics as specified in 3.1.2.14.i. No customer bleed.j. Zero humidity.k. Hot, tropical and cold atmospheric temperatures and pressures as defined in

Figure 24.

3.2.1.1 Performance Ratings. The performance ratings shall be in accordance with Tables IA, IB, IC, IIA, IIB, IIC and IID. The first stage turbine rotor inlet temperature for each rating established at 6000 ft, 35°C (95°F) conditions shall be constant, for all atmospheric and Mach number conditions except when limited by fuel flow, compressor aerodynamics, gas generator speed, or torque limits. The maximum specific fuel consumption requirements and the maximum gas generator speed requirements in the performance rating tables shall be at the corresponding minimum output shaft power specified in the performance rating tables. Computer Program and Rating Tables shall be updated after the QT Altitude Test if the Deck and engine do not lapse within ±2% of each other. The lapse rates shall include Ng, shaft horsepower and SFC lapse independently with both ambient temperature and pressure altitude for all the rating table conditions. If the Deck and Rating Tables are adjusted, it should be noted that this adjustment is based on a sample of one. If in the future, other samples show the lapse rate to be different, the Deck and Rating Tables shall be adjusted to reflect the actual lapse rate.

3.2.1.2 Performance Presentation. Engine performance data shall be provided in two forms: performance rating tables referenced in section 3.2.1.1, and a computer program suitable for use with an automatic digital computer. The number and date of the computer program and associated User’s Manual shall be specified herein. The performance data shall cover the operating envelope of the engine. The engine

42

Appendix D

rating points and the curves contained in this AQP shall agree with the computer program within the accuracy of AS 681J (±0.25%). If there is a discrepancy between the rating tables and computer program the rating tables shall take precedence. Points of rated performance shall be identified on the curves. The engine specification shall contain a list of the symbols and a diagrammatic figure in Section 7 defining station identification. These symbols and station identifications shall conform to AS 755D.

3.2.1.2.1 Performance Presentation Digital Computer Program. A steady-state performance computer program will be constructed and furnished to the Using Service. This computer program shall form a part of this AQP and shall carry a suitable identification and date, including the engine model designation. The computer program shall be compatible with a Windows 7 Professional or later-based operating system. The computer operating system, compiler and executing program shall be identified in the engine specification and is subject to Using Service approval. Compilation and execution of the official computer program shall be demonstrated on the specified computer prior to acceptance by the Using Service. The computer program shall be prepared IAW AS 681J, except as modified herein.

a. Program Requirements. The performance program shall be a thermodynamiccycle simulation in which component identity is maintained; e.g., the compressor, turbine, and combustor must each be identifiable as entities in the model logic as required to obtain and maintain an accurate simulation. The computer program shall be submitted in source language compatible with the computer specified. Nomenclature both internal and external to the program shall be IAW AS 755D.

b. Program Capabilities. The program shall be capable of operating throughout theengine operating envelope. The program shall also be capable of operating at ambient static pressures up to 16 psia. Compilation shall not be necessary for each different run. Capability for determination of installation effects shall be included in the computer program. Effects of distortion, relative humidity, ram recovery, customer bleed, customer power extraction, nozzle geometry, engine anti-icing, windmilling, module deterioration and variable geometry shall be included as applicable for the engine. The distortion limits of stall free operation and the effects of distortion on performance shall be included in the performance computer programs. The performance program shall contain the capability of modeling installation losses in the form of inlet and exhaust maps that can be placed in the input file. Compilation shall not be necessary to run various inlet and exhaust configurations. Capability for determination of any compressor inflow or vent installation effect shall be included in the computer program and updated after completion of the official EIT/PFR and QT altitude tests. Compilation shall not be necessary to run various customer bleed extraction or module deterioration.

c. Documentation Requirements. A user’s manual, source listing, and sourceprogram file distributed on disk or other medium determined suitable by the Using Service shall be provided.

1. User’s Manual. In addition to those items specified in AS 681J, the user’smanual shall contain a general description of the simulation techniques, general overall model flow chart and a clear explanation of the calculation process and related assumptions for all engine components. The user’s manual shall include a tabulation of all parameter limits and reference to all engine limits described in the engine specification, e.g., measured temperature, speed. All input parameters, output parameters, and parameters used in calculations shall be identified along with corresponding units for each parameter. Additionally, the user’s manual shall include graphs of all empirical functions used, reference values for normalized parameters,

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Appendix D

an index of subroutines including their inputs, outputs and functions, and a listing of all test case program inputs and the corresponding required outputs.

2. Source Program. A source program file distributed on disk or other medium determined suitable by the Using Service and program listing shall be provided. These shall include all program subroutines with comments to identify subroutines and their functions. Sufficient comments shall appear in the program code to elucidate the calculation process. The first lines of code shall be comments indicating the contractor, engine type designation, program number, and date. The format of the distributed medium shall be as specified by the Using Service.

d. Inputs and Outputs. The program output listing shall be so organized that the input data shall print separately from the output. The program output listing shall show the engine designation, program number, and date. Provision shall be made in the program for a line on the output sheet which can print declassification and authority. Program inputs shall be by namelist. Namelist nomenclature shall be IAW AS 755D. All program inputs must be independently variable and the programs shall be capable of sequentially accepting multiple changes to the computer inputs. In addition to the inputs listed in AS 681J, the program shall provide inputs of fuel heating value and measured temperature. The program shall be capable of calculating all the required output parameters of AS 681J, with input options of measured temperature with output shaft speed, and shaft power with output shaft speed, in addition to the options of rating code and power lever angle or power code, specified in AS 681J. In addition to the required output parameters of AS 681J, the program shall output gas temperature at the first stage turbine rotor inlet, gas temperature at the measurement plane (contractor shall parenthetically insert the measurement plane station identification here), exhaust gas velocity, and exhaust gas swirl angle. The program shall have features to print internal station properties such as physical flow, total pressure, and total temperature at all internal station locations. The program shall accept inputs and provide outputs in the U.S. Customary System of Units IAW ASTM E 380. SI units shall be selectively available. SI units shall be for information only and U.S. Customary System data shall be the sole specification data. All aerothermodynamic cycle parameters used in component performance evaluation and calculations must be available in labeled common block format as specified in AS 681J, are for information only, and will not be considered specification data. The method of selectively obtaining the cycle parameters by means of user input, shall be specified in the User’s Manual. The program output rating table parameters and all other program output parameters shall be characterized in the engine specification for both rating table conditions and all other operating envelope conditions in terms of whether they are estimated minimum, guaranteed minimum, nominal, or average performance values.

1. Inputs. All aerothermodynamic cycle parameters used in component performance evaluation calculations must be available in labeled common block format as specified in AS 681J. In addition to the inputs listed in AS 681J, free stream velocity (ZV0) in knots shall be provided. INLET MODE SELECTION shall be modified to SIM=0 (for altitude, free stream velocity, and ambient temperature). The following parameters shall be provided as Common Variable Inputs:

(a) ZPD1Q0 – Engine Inlet Duct Pressure Loss.(b) ZPD8Q7 – Engine Exit Pressure Loss.(c) ZA8 – Exhaust Nozzle Geometric Area(d) ZV0 – Free Stream Velocity.(e) SPART – Particle Separator Option.

Appendix D

44

(f) ZFLHV – Fuel Lower Heating Value.(g) ZTC (ZMGT) – Measurement Plane Control Temperature (if the

cockpit display temperature is the same as the measurement plan gas temperature or varies by a constant bias, only one is required – otherwise BOTH are required as separate inputs).

(h) The program shall have a Rating Code (ZRC = –55) that performsa power sweep consisting of 12 points ranging from maximum power to ground idle.

(i) All performance decks shall contain the input option to specify asecond, tabular output file when running multiple input points. The output parameters contained in this tabular output file shall be specified by the user within the input file. The tabular output file shall always contain the NSI codes as one (or more) of the data columns.

2. Outputs. In addition to the outputs listed in AS 681J the following outputsshall be provided:

(a) PD1Q0 – Engine Inlet Duct Pressure Loss.(b) PD8Q7 – Engine Exit Duct Pressure Loss.(c) A8 – Exit Nozzle Geometric Area.(d) V0 – Free Stream Velocity.(e) W01 – Engine Inlet Airflow.(f) P01 – Engine Inlet Pressure.(g) TC (MGT) - Measurement Plane Control Temperature (if the

cockpit display temperature is the same as the measurement plan gas temperature or varies by a constant bias, only one is required – otherwise BOTH are required as separate inputs).

(h) TRIT – First Stage Turbine Rotor Inlet Temperature.3. Engine Control Limits. Engine control limits which restrict the maximum output

shaft horsepower including, but not limited to, engine torque, corrected gas generator speed, fuel flow, measurement plane control temperature, and physical gas generator speed shall be set to the values corresponding to output values for a minimum guaranteed engine. A method shall be provided for changing these limits. If the measurement plane control temperature limit (or measurement plane gas temperature limit if applicable) is set below a value that would allow operation at the maximum allowable temperatures of section 3.2.1.4.5, the method for changing this limit shall not require compilation.

4. Precision. The program provided to the Using Service shall be the masterprogram. Any other programs of the same revision number shall match the output of the master program within ±0.250% on all parameters for all possible input conditions regardless of the hardware and software platforms used.

5. Verification. The test cases for verification will include all the engine ratingpoints, exercise all options, and include at least 100 cases in total. The input and output for each test case shall be included with the program source listing for verification.

3.2.1.2.2 Engine Transient Program. The transient program shall comply with all requirements of 3.2.1.2.1, except 3.2.1.2.1d(3) and 3.2.1.2.1d(4). The transient performance computer program shall be an average or status model, and the output is considered part of this AQP. Transient engine models used for the purpose of control system software development and qualification shall be capable of simulating engine

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Appendix D

starting, shutdown, and off-design operation of variable geometry, and shall be validated in accordance with 3.2.1.3. The transient program shall identify control system performance limiting conditions such as mechanical, thermal and aerodynamic limits and provide sufficient documentation as to the tolerances related to those limits. The program shall allow modification of limits set point values by the Using Service without compilation of the program. The transient program shall meet the consistency and precision requirements of AS681H. Component maps shall be sufficiently accurate in the idle and sub-idle range to characterize ground and airstart transient times to within 10 percent. Actual engine idle to intermediate and maximum transient times shall be accurate to within 5 percent of the transient performance computer program. Engine performance defined by the transient computer program shall be substantiated by sufficient test data to assure that theoretical assumptions used to develop the computer program are correct. This substantiation shall include submittal of a report comparing time rate of changes for torque, fuel flow, gas generator speed, power turbine speed, and measured gas temperature as calculated by the program to those measured in the test cell for the transients performed in the altitude test of Section 4.6.3.

3.2.1.3 Performance Verification. The performance presentation of 3.2.1 and 3.2.1.2 shall be used as the basis for verifying engine performance throughout the operating envelope. Engine performance defined by the computer program of 3.2.1.2 shall be substantiated by sufficient test data to assure that theoretical assumptions used to develop the computer program performance presentation are correct. The definition of ambient temperature and pressure lapse rates shall be a part of the substantiation test data. The substantiating data and analysis shall be submitted to the Using Service for approval prior to the start of the EIT and during the QT Phase. The following documentation shall also be required:

a. Component map information and thermodynamic cycle parameters shall be supplied. The supplied cycle parameters shall include cycle limits, leakages, variable geometry effects, control schedules and cooling air temperatures. A listing of the maps shall be provided to the Using Service in a suitable electronic format.

b. Factors used by the performance calculation program to simulate deterioration effects on the compressor, gas generator turbine and power turbine and their basis of derivation shall be submitted to the Using Service for approval as part of the performance verification process.

c. Factors used to refer and correct test data to performance values used as the basis for verifying the performance throughout the operating envelope shall be provided to the Using Service for approval as part of the performance verification process.

d. A schematic of the secondary airflow paths shall be presented. Each secondary path shall be listed as a percentage of the compressor inlet airflow.

e. When changes are made to component parameters, the new component map shall be submitted to the Using Service.

f. When the secondary airflow changes, a new secondary schematic diagram shall be submitted to the Using Service.

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Appendix D

3.2.1.4 Operating Limits. All engine steady-state and transient operating limits (maximum, minimum) shall be as specified. The specified limits shall be predicated on the most critical tolerances of the engine. The Using Service shall have the ability to adjust the tolerances in the computer program as described in section 3.2.1.2.1

3.2.1.4.1 Operating Envelope. The engine operating envelope limits as defined by aerothermodynamic and mechanical limitations shall be as specified IAW Figures 7 and 8. These operating envelope limits are the conditions within which the engine shall meet all specified steady-state and transient performance characteristics.

3.2.1.4.2 Sea Level Operating Limits. The engine Mach number limits for standard, cold and hot atmosphere sea level operation shall be as specified on Figure 7. The minimum Mach number shall be 0.35.

3.2.1.4.3 Absolute Altitude. The absolute altitude of the engine and the range of Mach numbers applicable at standard, hot and cold atmosphere conditions shall be as specified in Figure 7and and shall be a minimum of 20,000 feet and a Mach number of 0.35.

3.2.1.4.4 Starting Limits. The starting and operating limits shall be as specified in Figures 7 and 8. The engine shall start under ram and no-ram conditions from sea level up to at least the altitude of 20,000 feet. The maximum turbine temperature for ground and air starts shall be specified herein. Differences in the altitude starting limits for “hot” and “cold” engines shall be specified in the engine specification for ground elevations and flight altitudes. A “cold” engine shall be defined as one which has been allowed to windmill at the specified test condition until the engine combustor exit temperature is within 55°C (99°F) of the engine inlet temperature before a start is attempted. A “hot” start shall be defined as one where a start is attempted within 10 seconds after a flameout or shutdown. The contractor shall state the starting gas temperature limit (measured) at which point the start is aborted per paragraph 3.7.2.3.4.1. Limits with maximum customer bleed air and maximum power extraction shall be specified herein.

3.2.1.4.5 Engine Temperature Limits. a. First Stage Turbine Rotor Inlet Temperature. The maximum allowable steady-

state gas temperature at the first stage turbine rotor inlet (the Min Endurance Test Temperature given in Table VI) averaged over the gas path area, shall be specified herein. This temperature is at least 30°C (54°F) above the highest first stage turbine rotor inlet temperature that can occur within the engine operating envelope on Figure 7 and the 30°C (54°F) temperature difference is defined as the deterioration margin. The highest first stage turbine rotor inlet temperature that can occur within the engine operating envelope on Figure 7 corresponds to the first stage turbine rotor inlet temperature at the (contractor to specify) rating condition. The corresponding engine operating condition (including compressor inlet temperature and pressure) shall be specified. An analytical method for calculating the first stage turbine rotor inlet gas temperature value shall be provided by the engine manufacturer. This minimum of 30°C (54°F) deterioration margin shall be applied to each rating condition in the rating tables.

b. Measurement Plane Temperature. The maximum allowable steady-state gastemperature (the Min Endurance Test Temperature given in Table VI) averaged over the gas

47

Appendix D

path area at the measuring plane and corresponding to the maximum allowable steady-state gas temperature at the first stage turbine rotor inlet shall be specified herein (contractor shall parenthetically insert the measurement plane location and station identification here). The maximum allowable measured temperature at the measurement plane (based on the signal indication after the temperature sensor electrical harness plug) shall be specified herein. This temperature may be changed prior to the qualification endurance test based on correlation studies of the relationship between measured temperature and the gas temperature at the first stage turbine rotor inlet during the development program. The correlation studies shall be presented to the Using Service for approval prior to initiation of the engine integrity test endurance test and shall be revised and approved by the Using Service prior to the qualification endurance test. The maximum allowable transient (12 second) measured temperature during starting and for acceleration above idle shall be specified herein.

3.2.1.4.6 Rotor Speed Limits. The maximum allowable steady-state and transient rotor speed (mechanical) limits for each rotor system shall be specified herein. For the power turbine, 100% speed shall be 21,162 rpm.

3.2.1.4.7 Fuel Flow Limits. The maximum and minimum engine fuel flow shall be specified herein.

3.2.1.4.8 Oil Pressure and Temperature Limits. The maximum and minimum operating oil pressure limits for both steady and transient operation (starting and initial operation) and maximum allowable steady-state and transient oil temperature limits shall be as specified herein. Maximum oil pressures during starting and initial operation, predicated on a 13,000 centistokes oil viscosity, shall not persist for more than 2.5 minutes. Minimum oil pressure during starting and initial operation shall not persist for more than 30 seconds.

3.2.1.4.9 Oil Consumption Limits. The oil consumption, including all forms of oil loss, shall not exceed the amount as specified herein . If the average oil consumption rate during the qualification tests is less than one-third of the specified value, the specification oil consumption rate shall be adjusted to a value no greater than three times the qualification test average.

3.2.1.4.10 Vibration Limits. The maximum permissible engine vibration limits (overall velocity limit [true RMS]) at each sensor location on the engine compressor and turbine cases, and accessory gearbox case shall be as specified in Figure 29. The overall velocity limit specified shall be divided into two frequency bands with the lower frequency band covering airframe induced vibrations and the upper frequency band covering engine induced vibrations. The lower frequency band will have a lower cut off frequency of not more than 3 Hz and the upper frequency band will extend to 5 kHz. The bands will be continuous with the upper cut off frequency of the lower frequency band coincident with the lower cut off frequency of the upper band. The maximum permissible vibration levels shall be based on the structural strength of the engine and shall insure that wear caused by vibration to all engine parts and components, including bearings and seals, will not degrade the specified service life of the engine. In addition, engine operation throughout the entire load factor envelope of the aircraft shall be free from failure caused by rotor to case rubbing due to engine case deflection for the maximum permissible vibration levels. At the option of the Using Service, the true RMS velocity limits shall be provided prior to the initiation of the EIT for incorporation into the engine specification.

48

Appendix D

3.2.1.4.11 Output Shaft Torque Limits. The maximum allowable steady state torque (mechanical) limit shall be at least 20 percent greater than the highest rating value and shall be specified herein.

3.2.1.4.12 Output Shaft Speed Limits. The maximum allowable steady-state output shaft speed (mechanical) limit shall be at least 10 percent greater than the highest rating value and is as stated in 3.2.1.4.6. The maximum steady-state output shaft speed (governor) shall be at least 5 percent greater than the highest rating value. During non-governed (idle) conditions, the steady-state output shaft speed shall be set a minimum of 15 percent above the critical speed of the shafting system.

3.2.1.4.13 Customer Bleed Air And Power Extraction Limits. All limits on engine loading by means of accessory pads or customer bleed air ports shall be specified in the engine specification. Limits shall be specified for each port and pad, individually and for all possible combinations.

3.2.1.4.14 Emergency Power. The emergency power capability of the engine shall be defined. Included shall be a description of how the engine controls shall be activated for the emergency power application. The expected damage and potential failure modes resulting from the use of emergency power shall be defined for a new engine; engines at mid-life; and an engine nearing the end of its design life. After emergency power use, the engine shall require return to depot for repair. A method shall be proposed which indicates usage of emergency power to the operator.

3.2.1.5 Operating Characteristics.

3.2.1.5.1 Operating Attitude and Conditions. The engine shall be capable of continuous satisfactory operation in the clear area and at least 30 seconds operation in the shaded areas shown in Figure 9. Operation in the shaded area shall not degrade engine performance or cause any damage. The engine shall function satisfactorily for at least 60 seconds under conditions of negative “g” and for at least 30 seconds under zero “g” conditions. The engine shall be capable of being started, stopped, and stowed in any of the attitudes shown in the clear area on Figure 9.

3.2.1.5.2 Starting See 3.7.9.

3.2.1.5.3 Stopping. Stopping (termination of fuel flow) of the engine shall be accomplished by a single control command and it shall be possible to stop the engine by this means from any operating condition. Stopping of the engine from any power setting or at any rate shall not a) result in immediate or subsequent exceedance of any engine limits, b) adversely impact engine durability, structural integrity, or operational capability and c) cause rotor bowing or rotor seizure. The engine shall not experience any post-shutdown fires. No damage to the engine shall result from shutting off the fuel supply by the foregoing means or from shutting off the fuel supply to the engine inlet connection during any engine operating condition. Provisions for stopping the engine by means other than a completely mechanical system shall be subject to approval of the Using Service.

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Appendix D

In applications that require electrical power to accomplish the engine shut off function, additional independent means shall be provided to shut down the engine in an acceptable, safe time frame. It is acceptable for this independent means to be an aircraft provided fuel shutoff valve. Capability for Emergency stopping shall be provided by the airframe in the event of a complete control system failure. Fuel shut-off by means of an aircraft fuel shutoff of fuel can be assumed as part or all of such an emergency shutoff system. Adjusting the load of the shaft power absorber shall not be required when shutting down the engine.

3.2.1.5.4 Low Power Conditions.

3.2.1.5.4.1 Idle. Idle shall be the minimum self-sustaining condition of the engine and shall be set by the “idle power” command. Engine gas generator speed at idle power shall be as specified in Figure 5-1. Engine output shaft torque at idle shall be specified IAW Figure 5-2. Output shaft torque shall be as shown for output shaft speeds from zero to the maximum output shaft speed limit for altitudes of sea level, 2000 ft, 4000 ft, a n d t h e maximum altitude specified in 3.2.1.4.3 in approximately 20 percent speed increments. The engine shall be capable of operating satisfactorily at zero output shaft speed (locked PT rotor) for at least 30 minutes.

3.2.1.5.4.2 No Load Condition. No load condition shall be an engine operating condition and shall be established to maintain rated output shaft speed with a specified power absorber as defined within the limits of 3.1.2.14.1 operating at zero output shaft torque throughout the operating envelope. Engine gas generator speed(s) at no load shall be specified IAW Figure 6. The engine(s) and torque absorber system shall be capable of transitioning between load and no-load conditions without unacceptable mismatches in rotor speeds or torque defined in 3.2.1.5.6, 3.2.1.5.8, and without exceeding engine operating limits.

3.2.1.5.5 Stability. Under steady-state operating conditions, throughout the environmental conditions and operating envelope, output shaft power fluctuations (peak to peak) shall not exceed 1.0 percent of maximum continuous output shaft power between idle and maximum continuous power conditions. During operation above maximum continuous power, fluctuations shall not exceed 1.0 percent of the output shaft power available at that condition. During steady-state operating conditions, the period between frequency fluctuations greater than 0.5 percent of maximum continuous delivered shaft horsepower shall not be less than five seconds. Combustor stability margins shall be sufficient to protect against blowouts or flameouts during rapid throttle movements or load changes under worst case conditions. Additional requirements related to stability during operation of the bleed air system are stated in 3.1.2.11.

3.2.1.5.6 Engine Power Transients.

3.2.1.5.6.1 Minimum Requirements. The engine shall satisfactorily perform any power transient throughout the operating envelope while meeting the requirements of the specification. Power requests in any sequence and at any rate for both primary and backup control modes shall not result in exceeding any engine operating limit (including overspeed and overtemperature), result in unstable operation of the engine, or cause any mechanical damage. Power transient times in all backup control modes shall be specified herein. For a control power signal change, the times required to accomplish 95 percent

50

Appendix D

of the power change shall not exceed the values specified, when the power absorber load is applied/removed as specified, in “a” through “g” below. All transient times specified shall be based on standard conditions, with no anti-icing bleed air, but with all other engine system bleed air requirements, e.g., acceleration bleed air, cooling bleed air. The total time required to accomplish each specified power transient and reach stable operation shall be no more than that time noted plus ten seconds. Stable operation shall be as defined in 3.2.1.5.5.

a. Seven seconds from ground idle to maximum power, from Mach 0.0 to Mach 0.20 from sea level to 10000 ft altitude; power absorber load applied as specified in figure 10-3A.

b. Three seconds from no load to maximum power, from Mach 0.0 to Mach 0.20 from sea level to 10000 ft altitude; power absorber load applied as specified in figure 10-3B.

c. Three seconds from maximum power to no load, from Mach 0.0 to Mach 0.20 from sea level to 10000 ft altitude; power absorber load removed as specified in figure 10-3C.

d. Eight seconds from no load to maximum power, from Mach 0.0 to Mach 0.20 from 10000 ft to the maximum altitude; power absorber load applied as specified in figure 10-3D.

e. Eight seconds from maximum power to no load, from Mach 0.0 to Mach 0.20 from 10000 ft to the maximum altitude; power absorber load applied as specified in figure 10-3E.

f. One second from intermediate power to contingency power, from Mach 0.0 to Mach 0.20, from sea level to 10000 ft altitude; power absorber load applied as specified in figure 10-3F.

g. Two seconds from maximum continuous power to contingency power, from Mach 0.0 to Mach 0.20, from sea level to 10000 ft altitude; power absorber load applied as specified in figure 10-3G. Power transient times shall not exceed 125 percent of the times for conditions “a” through “g” above, under any of the following conditions, singly or in any combination: nonstandard conditions throughout the environmental conditions and operating envelope; customer power extraction; customer bleed air extraction; engine anti-icing bleed air extraction; inlet distortion; icing environment. Note that the transients to and from no-load are at constant commanded output shaft speed. Transients to and from idle use the minimum self-sustaining engine and output shaft speeds for idle. Continuous sinusoidal load transients with peak-to-peak power change demands of up to +5 percent with a frequency up to 1.0 cycle per second shall not result in a divergent sinusoidal power response. Power response phase lag shall be no more than 135 degrees and power peaks as a percent of demand shall be specified. Symmetrical throttle demands shall not result in divergent symmetrical throttle responses.

3.2.1.5.6.2 Estimates. Estimates of transient performance shall be as presented IAW Figures 10-1 and 10-2.

3.2.1.5.7 Engine Windmilling Capability. The engine shall be capable of continuous windmilling operation throughout its entire altitude/Mach number flight envelope and the clear area of Figure 9 (attitude envelope) without

51

Appendix D

damage, without excessive loss of lubricating oil, and without affecting capability of air restart and operation. Additional information is as follows:

a. Windmilling operation shall be allowable for at least eight hours and shall be specified herein.

b. The oil consumption rate during windmilling operation shall not exceed the maximum oil consumption limits of 3.2.1.4.9.

c. The minimum duration and limits of windmilling operation after depletion of oil supply shall be two hours.

d. No customer power and bleed air extraction shall be available during windmilling operation.

3.2.1.5.8 Rotor Droop and Overshoot. The engine shall not allow output shaft speed to deviate more than the limits specified in Table XXIII for the defined power transients at all specified ambient conditions. This shall be verified through computer modeling using either a full, linearized rotor model or a lumped inertia model (attached to the power turbine) derived from the full, linearized rotor model. The actual value for 100% aircraft torque shall be determined using 100% engine output shaft speed and the value of the power required for each engine as specified by the Using Service through coordination with the air vehicle manufacturer.

3.2.2 Physical Characteristics.

3.2.2.1 Dry Weight of Complete Engine. The dry weight of the complete engine shall not exceed that specified herein the engine specification. If the specification value exceeds the dry weight of the official qualification test engine by more than 2.0 percent, the specification value shall be revised downward to within 1.5 percent of the dry weight of the test engine. The weight of the engine components which are not mounted on the engine shall be listed and included in the dry weight of the engine. The dry weight of any electrical harness extensions for off-engine mounted components shall be a separate weight item added to the dry weight of the complete engine.

3.2.2.2 Weight of Residual Fluids. The weight of residual fluids remaining in the engine after operation and drainage, while the engine is in a horizontal attitude of the main rotor axis relative to the level plane, shall be specified herein. The weight of operating fluid shall also be specified herein. These values shall be verified or modified prior to EIT/PFR and QT and submitted to the Using Service for approval.

3.2.3 Not used.

3.2.4 Not used.

3.2.5 Environmental Conditions.

3.2.5.1 Ambient Temperature Conditions. The complete engine shall start and, thereafter, be capable of operating satisfactorily throughout the Mach number, temperature, and altitude conditions specified in Figure 7 using the fuels and

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Appendix D

oils specified in this AQP, and consistent with 3.7.3.3.3 and 3.7.3.3.4, under the following conditions:

a. From surface (ground) elevations of sea level to the maximum altitude specified in 3.2.1.4.3 after the engine has been soaked at a temperature of 16°C (29°F) warmer than hot atmospheric air temperature of Figure 24 for a period of at least 10 hours when supplied with fuel 16°C (29°F) warmer than hot atmospheric air temperature and engine inlet air at hot atmospheric conditions. Restart and operate satisfactorily from surface (ground) elevations of sea level to 20000 ft after a soak period of 15 minutes at an ambient air temperature 80°C (144°F) warmer than hot atmospheric air temperature when supplied with fuel and inlet air as specified above.

b. From surface (ground) elevations of sea level to the maximum altitude specified in 3.2.1.4.3, after the engine has been soaked for a period of at least 10 hours at an ambient temperature of -54°C (-65°F) or that temperature corresponding to a fuel viscosity of 12 centistokes, whichever is warmer, and when supplied with fuel and air at -54°C (-65°F) or that temperature corresponding to a fuel viscosity of 12 centistokes, whichever is warmer.

c. From surface (ground) elevations of sea level to the maximum altitude specified in 3.2.1.4.3 and throughout the starting envelope defined in Figure 7 for air starting when supplied with inlet air at all temperatures between hot atmospheric air temperature and cold atmospheric air temperature and when supplied with fuel at any temperatures between 16°C (29°F) hotter than hot atmospheric air temperature and cold atmospheric air temperature or that temperature corresponding to a fuel viscosity of 12 centistokes, whichever is warmer.

3.2.5.2 Icing Conditions. The engine shall operate satisfactorily under the meteorological conditions shown in Figures 11 and 12, and Table XIII with not more than 5.0 percent total loss in output shaft power at constant first stage turbine rotor inlet temperature at all operating conditions above maximum continuous power and 5.0 percent total increase in specific fuel consumption at constant power at all operating conditions above 50% maximum continuous power. Operation at less than maximum continuous power shall be such that 95 percent of the power desired above maximum continuous power can be obtained within the specified acceleration time. Upon termination of the icing conditions, the engine shall retain no performance deterioration. No damage to the engine shall be allowed as a result of operating in the icing environment.

3.2.5.3 Not used.

3.2.5.4 Not used.

3.2.5.5 Corrosive Atmosphere Conditions. The engine shall meet all requirements of this AQP during and after exposure to the corrosive atmosphere of Table XIV for 25 corrosion cycles. The engine shall not deteriorate more than 5 percent in power at constant first stage turbine rotor inlet temperature, or gain 5 percent in specific fuel consumption at constant power. Engine parts shall show no corrosion which impairs function, component integrity, or prescribed maintenance. Corrosion protected parts shall show no corrosion effects after cleaning and stripping of corrosion protection coatings.

3.2.5.6 Environmental Ingestion Capability.

3.2.5.6.1 Bird Ingestion. The engine shall continue to operate and perform during and after the ingestion of birds.

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Appendix D

The inlet area to be used shall be the engine inlet. The number of birds to be ingested shall be based on inlet area as follows: one 3.5 oz bird per 46.5 in

2 of inlet area plus any fraction larger than 50 percent thereof, up to a maximum of 16 birds;

one 2.2 lb. bird per 232.5 in2 of inlet area plus any fraction larger than 50 percent thereof; one 4.4

lb. bird regardless of the size of the inlet, provided the inlet is large enough to admit a 4.4 lb. bird. The contractor shall compare the size of a 4.4 lb bird to the engine inlet and provide this analysis to the Using Service for approval. The 3.5 oz birds shall be ingested at random intervals and be randomly dispersed over the inlet area. Birds 2.2 lb. and larger shall be directed at critical areas of the engine face. The bird velocity and engine power setting for each condition shall be as described below:

a. ___ bird(s) of ___ lb ingested at a bird velocity equal to Mach number of 0.1 and with the engine at maximum (takeoff) rated MGT.

b. ___ bird(s) of ___ lb ingested at a bird velocity equal to Mach number of 0.3 with the engine at maximum continuous rated MGT.

c. ___ bird(s) of ___ lb ingested at a bird velocity equal to a Mach number of 0.2 with the engine at 25 percent of the maximum continuous power obtained in part b. The bird ingestion tests shall be conducted on an engine mounted in a typical aircraft installation configuration. Under the conditions listed no failures shall result which cause shutdown of the engine although limited damage to engine components may occur. By definition a bird has been ingested if it has entered the engine inlet plane. Small Birds: For the test sequence in part a, no engine flameout shall occur and the engine shall recover to no less than 95 percent of the shaft power that existed prior to bird ingestion without exceeding any engine control limits. For the test sequence in parts b and c, no engine flameout shall occur and the engine shall recover to no less than 100 percent of the shaft power that existed prior to bird ingestion without exceeding any engine control limits (MGT and/or Ng physical/corrected speed). The control limits shall be set such that there is no more deterioration margin available than is allowed by the specification for Maximum (Takeoff) redline. The performance recovery time after ingestion of the bird(s) shall occur in 3 seconds or less after the final bird has been ingested. After the 3 second recovery time, the engine shall meet the stability requirement of section 3.2.1.5.5. Medium Birds: For the test sequence in part a, no engine flameout shall occur and the engine shall recover to no less than 75 percent of the shaft power that existed prior to bird ingestion without exceeding any engine control limits (MGT limits and/or Ng physical/corrected speed). For the test sequence in part b, no engine flameout shall occur and the engine shall recover to no less than 90 percent of the shaft power that existed prior to bird ingestion without exceeding any engine control limits (MGT and/or Ng physical/corrected speed limits). For the test sequence in part c, no engine flameout shall occur and the engine shall recover to no less than 100 percent of the shaft power that existed prior to bird ingestion without exceeding any engine control limits (MGT and/or Ng physical/corrected speed). The control limits shall be set such that there is no more deterioration margin available than is allowed by the specification for Maximum (Takeoff) redline. The performance recovery time after ingestion of the bird(s) shall be 3 seconds or less after the final bird has been ingested. After the 3 second recovery time, the engine shall meet the stability requirement of paragraph 3.2.1.5.5. Following each test condition identified in parts a through c above (for both small and medium birds) and prior to shutting down, the engine shall satisfactorily perform engine power transients

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Appendix D

a, b, and c of paragraph 3.2.1.5.6.1. Power transient times shall not exceed twice the value of those specified in 3.2.1.5.6.1.

Following each test condition identified in parts a through c and after the power transient demonstration, a performance evaluation shall be made to determine permanent power loss at constant rated MGT. The inlet may be cleaned and the engine water washed (per 3.7.13) prior to the evaluation. Permanent performance loss shall not exceed 5% for the small birds and 10% for the medium following each of the three test conditions. Large Birds: The 4.4 lb bird ingestion (if tested) shall not cause an engine failure which results in damage to the aircraft or adjacent engines.

3.2.5.6.2 Foreign Object Damage (FOD). The engine shall meet the requirements of the specification for the design service life of 3.3.8.1 utilizing the material properties identified in 3.3.8.5 without repair after ingestion of foreign objects which produce damage equivalent to a fatigue notch factor (Kf) of 3 at the most critical FOD location(s) of all compressor airfoils (blades, stators, and vanes of all compressor stages) as determined by the Contractor. A list of critical FOD location(s) and technical rationale for selection shall be approved by the Using Service. Unless approved by the Using Service, the fatigue notch factor (Kf) of 3 shall be applied to both the mean and alternating stresses.

3.2.5.6.3 Ice Ingestion. The engine shall be capable of ingesting hail and sheet ice as specified in 4.6.4.6 without flameout, lengthy power recovery, or major structural damage which could cause the engine to fail. Within 1 second after a hailstone ingestion or ice ingestion event, the engine power shall be at least 95 percent of the power immediately prior to the event at constant turbine rotor inlet temperature. Total power recovery time shall not exceed 3 seconds. For an engine inlet capture area of 100 in2 the engine shall be capable of ingesting one 1.0 in diameter hailstone. For each additional 100 in2 inlet area increase or fraction thereof, supplement the first hailstone with one 1.0 in and one 2.0 in diameter hailstone. Hailstones shall be ingested at typical takeoff (maximum power), cruise, and descent conditions. The engine shall be capable of ingesting one piece of sheet ice 3.0 in x 9.0 in x 0.25 in. Sheet ice should be ingested at typical takeoff and cruise conditions. The hailstones and sheet ice shall be between 50 lb/ft3 and 56 lb/ft3 density.

3.2.5.6.4 Sand Ingestion.

3.2.5.6.4.1 Fine Sand Ingestion. Unless otherwise specified, the engine shall meet all requirements of the specification during and after the sand and dust ingestion specified herein. The engine, including all components, shall operate satisfactorily throughout its operating range at ground environmental conditions with air containing 0-80 micron sand and dust in a concentration of 3.3 x 10-6 lb/ft3 of air for 54 hours, shall be able to conduct the accelerations of 3.2.1.5.6.1 without surge or stall and inspection shall reveal no impending failure. The engine and its components shall be capable of operating with air containing the specified concentration of sand and dust for a total period of at least 54 hours with not greater than 10 percent loss in output shaft power at constant first stage turbine rotor inlet temperature, 5 percent increase in specific fuel consumption at constant power, and no impairment of capability to execute satisfactory power transients as defined in 3.2.1.5.6.1 without surge or stall. Post test inspection

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Appendix D

shall reveal no impending failure of any components. The sand and dust constituents shall be IAW ISO 12103-1, A2 Fine Grade Test Dust which has the following size distribution:

Particle Size Microns

Quantity (by weight) Finer than Size Indicated

Percent of Weight (%) 1 2.5 – 3.5 2 10.5 – 12.5 3 18.5 – 22.0 4 15.5 – 29.5 5 31.0 – 36.0 7 41.0 – 46.0 10 50.0 – 54.0 20 70.0 – 74.0 40 88.0 – 91.0 80 99.5 - 100.0

3.2.5.6.4.2 Coarse Sand Ingestion. The engine, including all components, shall operate satisfactorily throughout its operating range at ground environmental conditions with air containing 0-1000 micron sand and dust particles in concentrations up to 3.3 x 10-6 lb/ft3 of air. The engine shall be calibrated prior to the test and recalibrated after the test. The calibration shall be conducted in accordance with paragraphs 4.6.1.2.2a except compliance with the starting torque requirement (paragraph 3.2.1.5.2) need not be shown. The recalibration shall be in accordance with paragraphs 4.6.1.2.2a except that compliance with the starting torque requirement need not be shown. The recalibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature condition, the values of output shaft power obtained during the initial calibration. The recalibration may be preceded by a specified run during which the cleaning procedure of paragraph 3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration. The engine and its components shall be capable of operating at maximum continuous rated measured gas temperature with the specified concentration of sand and dust for a total period of at least 50 hours with not greater than 10 percent loss in output shaft power at constant first stage turbine rotor inlet temperature, 5 percent increase in specific fuel consumption at constant power, and no impairment of capability to execute satisfactory power transients as defined in 3.2.1.5.6.1. The output shaft power deterioration shall not exceed 6.0 percent after 25 hours of sand ingestion. The final performance recalibrations shall be post test prior to any disassembly of the engine. Benching of the compressor blades is not allowed to regain performance loss. If an inlet air particle separator (IPS) is required to comply with the above, it shall be part of the engine and easily replaceable. If an IPS blower is required as part of the design it shall be capable of completing the 50 hour test without replacement. Based on operational experience, a representative sand contaminant shall consist of crushed quartz (SiO2) with the total particle size distribution as follows:

Particle Size Quantity (by weight) Finer than Size Indicated Microns Percent of Weight (%) 1,000 100 707 95-99500 89-93

Appendix D

56

354 77-81250 60-64177 38-42125 18-2288 6-1063 1-5

3.2.5.6.5 Atmospheric Liquid Water Ingestion. The engine shall operate satisfactorily throughout the operating envelope up to the absolute altitude at power settings from idle to maximum power with up to 5.0 percent of the total airflow mass in the form of water (liquid and vapor) and with 50 percent of the liquid water entering the engine inlet through a segment equivalent to one-third of the inlet area.

3.2.5.7 Noise Levels. The engine shall operate such that the total engine noise signature shall be minimized. The engine noise signatures shall be IAW Figures 30-1, 30-2, and 30-3.

3.2.5.8 Exhaust Gas Contamination.

3.2.5.8.1 Exhaust Smoke Emission. The engine shall not emit visible exhaust smoke at any power setting throughout the environmental conditions and the engine operating envelope when using any primary fuel specified in 3.7.3.2.1 or any restricted fuel specified in 3.7.3.2.2. The maximum allowable smoke emission level shall be determined from Figure 13 when using the exhaust nozzle of 3.7.10.2 and Figure 26 and when measured by the method of ARP 1179D and when using MIL-DTL-5624U grades JP-4, and JP-5, and MIL-DTL-83133H, grade JP-8 fuel with a minimum aromatic content of 25 percent by volume. Toluene may be added to the fuel to attain the required aromatic content.

3.2.5.8.2 Invisible Exhaust Mass Emissions. The engine shall not produce byproduct exhaust emissions greater than the levels specified on figures 14-1 and 14-2 when using any primary or alternate fuel. Emission limits shall be in compliance with the Federal Clean Air Act Amendments of 1990 as follows:

a. CO and HC levels must be below levels, which result in an idle combustionefficiency of 99.5 percent for engines with an idle pressure ratio above 3:1.

b. CO and HC levels must be below levels, which result in a combustion efficiencyof 99 percent for engines with an idle pressure ratio below 3:1.

c. NOX levels must be less than 50 percent of the uncontrolled level of figure 14-1.The contractor shall supply the emission limits in the format of figure 14-2 and tabulation of pollutant emission index vs. following power settings: idle, 25% maximum continuous, 75% maximum continuous, maximum continuous, intermediate, and maximum. The procedure for measuring the exhaust constituents emitted by the engine shall be as specified in ARP 1256D. The hydrocarbons are specified on the mass basis of methane and the oxides of nitrogen are specified on the basis of nitrogen dioxide. MIL-DTL-5624U, grades JP-4 and JP-5

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Appendix D

and MIL-DTL-83133H, grade JP-8 fuel with a minimum aromatic content of 25 percent by volume shall be used. Toluene may be added to the fuel to attain the required aromatic content. The emissions shall be determined from the average of the measurements obtained from at least four engine tests, one test with JP-4 and one tests with JP-5 and two tests with JP-8.

3.2.6 Transportability. The engine shall be suitable for movement on, and be compatible with, standard aircraft maintenance trailers. Adequate ground handling pads and other features shall be provided to permit installation on and use of appropriate static and mobile ground equipment. The Army standard aircraft maintenance trailer (NSN 1730-01-086-1653) conforms to Data List 4920-EG-081.

3.2.7 External Surfaces. The external surfaces of the engine, including its components, shall be designed to not collect dirt, dust, sand and liquids.

3.2.8 Survivability and Vulnerability.

3.2.8.1 Ballistic Vulnerability. The engine shall apply a combat survivable design concept that includes state of the art survivability technology (vulnerability and susceptibility) which insures survival in the intended threat environment. The engine shall not suffer an uncontained, instantaneous catastrophic failure after sustaining damage from a single 7.62mm Armor Piercing Incendiary (API) munition striking the engine in the lower hemisphere plus 15 degrees at 2500 feet per second (FPS). For larger munitions, up to 30mm high explosive (HE), the engine shall not suffer an uncontained, instantaneous catastrophic failure excluding a direct hit on a disk or main driveshaft. The engine shall be capable of being shutdown following ballistic damage.

3.2.8.2 Not used.

3.2.8.3 Not used.

3.2.8.4 Electromagnetic Environmental Effects (E3). The engine shall meet the electromagnetic environmental effects of ADS-37A-PRF with Table I, Part B.

3.2.8.4.1 Electromagnetic Interference (EMI). All engine electrical and electronic equipment and subsystems shall meet their performance requirements when subjected to the susceptibility requirements and emission requirements of ADS-37A-PRF as tailored by paragraph 3.2.8.4. These requirements shall be met for engine operation in all environmental conditions within the operating envelope for all control system operating modes.

3.2.8.4.2 Not used.

3.2.8.4.3 Electromagnetic Pulse. All engine electrical and electronic equipment and subsystems shall meet their performance requirements when subjected to the EMP requirements of ADS-37A-PRF.

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3.2.8.4.4 Lightning. All engine electrical and electronic equipment and subsystems shall meet their performance requirements when subjected to the lightning requirements of ADS-37A-PRF for indirect effects and tested IAW DO-160G, section 22. Applicable waveforms and voltage/current requirements for pin injection, Multiple-Stroke, and Multiple-Burst based on the location of their ECU in relation to the engine shall be provided.

3.2.8.4.5 Nuclear Weapons Effects. Shall be defined in the RFP.

3.3 Design and Construction.

3.3.1 Materials, Processes, and Fasteners.

3.3.1.1 Materials and Processes. The use of magnesium and cadmium is prohibited. Silver shall not be used without approval from the Using Service. Metals that are widely separated in the galvanic series shall not be used in direct contact with each other unless they are properly protected and approved by the Using Service. Any use of aluminum shall be insulated from contact with any other alloy in the engine. Copper shall be prevented from coming in contact with fuel or oil during engine operation. The use of brass and cadmium plating is not allowed in the fuel system due to possible corrosion of parts which contact fuel.

3.3.1.1.1 Adhesives and Sealants. The use, in any location, of adhesive or sealant compounds is permitted only upon specific application approval of the Using Service prior to its incorporation into the design of the engine and, as a minimum, shall be governed by the limitations listed in MS 18069.

3.3.1.1.2 Elastomeric Materials. Elastomeric material shall have unlimited shelf life and must not be allowed to deteriorate, to permit voids to form in the seals and, in turn, allow the fluid to leak past the seal. “Unlimited shelf life” for elastomers is equivalent to the term “non-age sensitive” elastomers. Elastomeric materials used in the engine shall be compatible with any fuel and fuel additives used by the Army (JP-8 +100).

3.3.1.1.3 “O” Ring Seals and Packing. Materials exposed to fuels and lubricants shall be compatible with such fluids throughout the entire fuel or lubrication system temperature cyclic envelope without experiencing swelling, shrinking, or other forms of material deterioration which would impair proper functioning or necessitate replacement to prevent impairment of function.

3.3.1.1.4 Corrosion Protection. The materials, coatings, and processes employed in the design and manufacture of the complete engine shall be corrosion and NBC resistant.

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3.3.1.2 Fasteners.

3.3.1.2.1 Self-Retaining Bolts. A self-retaining feature shall be used at joints where loss of the fastener could affect safety of flight or ability to control the engine.

3.3.1.2.2 Securing of Fasteners. Cotter pins, safety wire or safety cable shall not be used.

3.3.2 Not used.

3.3.3 Nameplate and Product Marking. Equipment, assemblies, modules, and parts shall be marked for identification with human and machine-readable unique identification (UID) in accordance with MIL-STD-130N. Component removal shall not be necessary in order to read the data plate. The type and length of product identification characters on the data plate shall be compatible with the Using Service’s maintenance tracking system. The color red shall not be used for marking unless adjacent non-red color marking is easily readable in red light conditions. See MIL-L-85762A for guidance.

3.3.3.1 Engine data plate marking. The engine data plate shall include: manufacturer’s identification, engine serial number (MIL-HDBK-1559), purchase order or contract number and engine model designation (MIL-HDBK-1812) with human and machine-readable UID.

3.3.3.2 Critical parts/critical safety items identification and tracking. The engine life critical parts/critical safety items and subassemblies shall be identified by serial numbers marked legibly on the part. The parts shall have multiple, non-wearing reference surfaces and have a space designated for marking the number of cycles and time accumulated between each overhaul period. The engine life critical parts/critical safety items and subassemblies shall be identified by serial numbers marked legibly with machine-readable UID on the part. All parts shall be marked with the drawing revision letter to which the part was made.

3.3.3.3 LRU/WRA marking. All LRU/WRAs shall have a barcode in accordance with ATA Spec 2000, Chapter 9. The barcode shall contain, at a minimum: (1) Applicable Enterprise Identifier (2) Serial Number (3) Part Identification Number.

3.3.4 System Nuclear Survivability. Shall be defined in the RFP.

3.3.5 Interchangeability. All parts having the same manufacturer’s part number shall be functionally and dimensionally interchangeable and replaceable with each other with respect to installation and performance, except that matched parts or selective fits will be permitted where required. If the use of matched parts and selective fits is required then a list shall be provided to the Using Service for approval at PDR, CDR, as well as EIT, and QT. Complete engines shall be interchangeable in multiengine aircraft (i.e., right side to left side).

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3.3.6 Safety.

3.3.6.1 Flammable Fluid Systems. All exterior lines, fittings and components which contain flammable fluid (fuel, lubrication oil and hydraulic fluid) shall be fireproof for at least 15-minutes at a minimum flame temperature of 1090°C (1994°F). The fuel pump, whether stand alone or integral to the FMU, shall be fire resistant for at least five minutes at a minimum flame temperature of 1090°C (1994°F). Other engine drive components, such as the lubrication system oil pump shall be fire proof for at least 15 minutes at a minimum flame temperature of 1090°C (1994°F). For an additional five minutes thereafter, the components shall not leak any flammable fluids in a quantity sufficient to intensify or re-ignite a fire, as determined by the Using Service. During exposure to the above conditions, the lines and components shall be conveying fluids under the worst possible operational combination of fluid parameters (i.e., flow rate, pressure and temperature) encountered throughout the complete environmental conditions and operating envelope of the engine.

3.3.6.2 Fire Prevention. Where engine surface area temperatures are sufficient to ignite flammable fluid that could impinge on the engine, these areas shall be specifically identified herein. If a fire shield is required, it shall be specified herein and the interface shall be shown on the Engine Configuration and Envelope Figure.

3.3.6.3 Explosion-Proof. Electrical components (except devices intended to ignite fuel air mixtures) shall not ignite any explosive mixture surrounding the electrical components.

3.3.6.4 Fluid Leakage. There shall be no leakage from any part of the engine except at the drains provided for this purpose. The total quantity of leakage from all drains, except that specified in 3.3.6.5, shall not exceed 0.169 fluid ounces per minute.

3.3.6.5 Combustible Fluid Drains. Provisions shall be made to prevent entrapment of fuel after each false start and for preventing excess combustible fluids from entering the combustion areas after shutdown with the engine in a level position, 15 degrees nose up, and 20 degrees nose down. No pilot or maintenance action shall be required to clear the engine of fuel after each false start or following engine shutdown. Provisions shall also be made for clearing all vent areas and other pockets or compartments where combustible fluids may collect during or subsequent to operation of the engine. The engine shall be designed so that combustible fluids cannot enter the combustion area when the engine is in, or turned to, a vertical position for maintenance. The maximum allowable quantity of combustible fluid which will discharge from the engine drains after shutdown from normal operation shall be specified herein. Requirements for the common drain shall be specified in the RFP.

3.3.6.6 Ground Safety. Warning notices shall be provided, where applicable, for high voltage electrical sources, radioactive devices, high powered optical sources, and explosive devices. For service while

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under red light conditions, all warning notices shall be capable of being read under red light conditions.

3.3.6.7 Engine Control System Safety. Engine control system components, when considered separately and in relation to other engine systems, shall have no catastrophic failure condition with a probability of occurrence greater than 10-6 and no critical failure condition with a probability of occurrence greater than 10-5 over the design service life specified in paragraph 3.3.8.1 below. Hazard severity categories catastrophic and critical are defined in MIL-STD-882E, Table I.

The engine control system shall be designed to fail safe. Fail safe is defined as the engine failing to a fixed fuel flow or power (depending on the engine control system architecture) or other condition approved by the Using Service. In all cases, the engine must be capable of being safely shutdown from the fail safe condition by the flight crew.

If the engine control system is located in a designated fire zone, the engine control system shall be fire resistant for at least five minutes at a minimum flame temperature of 1090°C (1994°F). During this period, and for five minutes thereafter, the engine control system shall continue to control the engine in accordance with the requirements of this AQP, or shall cause the engine to fail safe, as defined in the preceding paragraph. Additionally the fire shall not cause the engine to lose the capability to shut down. For system designs that depend on electric power to actuate the shutoff valve, high temperature wire or other protective means shall be used to ensure that the capability to shutdown the engine is maintained when the engine control system is exposed to fire.

3.3.7 Not used.

3.3.8 Structural Performance. All structural analysis methodologies used to substantiate structural performance shall require approval by the Using Service.

3.3.8.1 Design Service Life. Design usage shall not be less severe than the duty cycle shown below at the engine operating condition defined by 3.3.8.6.1.d Performance Lifing Model. The engine and all associated components shall have a minimum life of 6000 hours. Bearing B1 life shall be a minimum of 6000 hours. The 6000 hr life is based on the operational requirements below:

a. Power Condition Percent Engine Life at this Power (%)Contingency 0.1 Maximum 2 Intermediate 7 Maximum Continuous 45.9 60 percent MCP 20 40 percent MCP 20 Idle 5

b. The use of emergency power shall require removal of the engine from the airvehicle and return to the depot for inspection and refurbishment as required, per appropriate maintenance procedures.

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c. The off-engine mounted components shall have a minimum life of 6000 hours when exposed to the vibratory environment identified herein, at the ambient still air temperature limit of 3.1.2.8.1.

3.3.8.1.1 Expendables Design Service Life The minimum life without replacement of all expendable parts and components shall be equal to the minimum maintenance-free operating period.

3.3.8.1.2 Parts Classification – Critical Safety Items. All engine parts shall be classified for criticality per MIL-STD-882EE and listed in Table XXIV. The critical characteristics for all identified critical safety items shall be defined. The list of identified critical safety items and their critical characteristics shall be submitted to the Using Service for approval.

3.3.8.2 High Cycle Fatigue Life (HCF). Engine parts shall not fail when subject to the maximum attainable combined steady state and vibratory stresses. Parts that are subjected to LCF loads in addition to HCF loads shall be designed considering the effect of low cycle fatigue (LCF) damage on the material HCF life. The vibratory or HCF stress including any applicable stress concentration per 3.2.5.6.2 shall be restricted to 70 percent of the -3σ material HCF capability. Substantiation of this requirement shall be by Goodman diagrams presented for each component analyzed for HCF in the reporting requirements of 3.3.8.10.2.

For materials which have a discrete endurance limit knee on the stress versus cycles to failure curve, the contractor shall design the component to possess a minimum HCF life of 30,000,000 cycles.

For materials which do not have a discrete endurance limit knee on the stress versus cycles to failure curve, the contractor shall design the component to possess a minimum HCF life of 500,000,000 cycles or the actual cyclic life (frequency times 6000 hours).

3.3.8.3 Low Cycle Fatigue (LCF) Life. The -3σ LCF life of engine parts (except all turbine rotor blades and combustor liners) shall be 15,000 cycles. The minimum LCF life of turbine rotor blades and combustor liners shall be 7500 cycles. The LCF life and the fatigue crack growth life shall be based on the LCF analysis cycle as specified in the table below. All starts shall be made after the engine has cooled to ambient temperature. For the power turbine, the LCF analysis shall be conducted at 105% speed (22,220 rpm). The analysis shall be conducted IAW 3.3.8.6. The Contractor shall list in Table IX, the Critical Safety Items, as defined in 3.3.8.1.1, most susceptible to low cycle fatigue with their lives.

The LCF analysis cycle shall be as follows: Approximate Approximate total time schedule time (minute) (minute) Event

0.5 0.5 Start engine 2.5 2.0 Run at idle 2.6 0.1 Accelerate to maximum power 5.1 2.5 Run at maximum power 5.2 0.1 Decelerate to idle

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8.2 3.0 Run at idle 8.3 0.1 Accelerate to maximum continuous power 10.8 2.5 Run at maximum continuous power 10.9 0.1 Decelerate to idle 12.9 2.0 Run at idle 13.5 0.6 Shutdown TDB TDB Cool down

Note: Cool down shall be specified by the Contractor based on the analysis time required to return complete engine to ambient temperature.

3.3.8.4 Strength. The engine shall meet all the requirements of the specification during and after exposure to limit loads, singly and in combination where they occur naturally. The engine shall not experience catastrophic failure when subjected to ultimate loads, singly and in combination where the combinations occur naturally. The engine shall meet these strength requirements at the end of its life as defined by 3.3.8.1, 3.3.8.2 and 3.3.8.3.

3.3.8.4.1 Factors of Safety. The following factors of safety shall be applied to design usage induced loads to establish limit and ultimate conditions for all engine parts. All engine cases, mounts, and pressure loaded parts and components shall meet the requirements of this AQP during and after exposure to the ultimate loading conditions. The cases must remain intact, although permanent deformation and distress, requiring repair or replacement is permitted. Engine cases shall not fail due to combustion process burning or erosion.

LOAD TYPES LIMIT ULTIMATE Externally applied loads 1.0 1.5

Thermal loads 1.0 1.0 Mechanical loads 1.0 1.5

Internal pressures (air, fuel and oil fluid Lines)

1.5 2.0

Internal pressure (pressure vessels/structural case design)

1.33 2.0

Aircraft flow field loads 1.0 1.5 Crash loads N/A 1.0

NOTES: For all castings, a factor of safety of 1.33 shall be applied to the limit and ultimate load factors specified above, unless the castings have been fully characterized including fracture mechanics analysis of critical areas to address casting porosity and inclusions.

3.3.8.4.2 Blade and Disk Deflection. Blade and disk contact with any static parts of the engine during all phases of engine operation including surge and stall occurrences shall not adversely impact component capability to meet all requirements of this AQP.

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3.3.8.5 Material Properties. Material properties used in all analyses shall be based on minimum material capability, defined herein as -3σ. Therefore, all material properties shall be based on -3σ values with the exception noted below. These -3σ values shall be calculated using methodologies reviewed and approved by the Using Service. All relevant material properties (including property curves) used in any analysis shall be provided in the applicable report(s) to the Using Service. Use of Fatigue Crack Growth (FCG) to meet component life requirements shall be reviewed and approved by the Using Service. If FCG is used for any portion of component lives, the requirements of paragraphs 3.3.8.1, 3.3.8.2, 3.3.8.3 and 3.3.8.4 must be met, and the calculation of crack propagation life from initial flaw size to critical crack size shall be based on a validated crack growth analysis system. The crack growth analysis system shall be reviewed and approved by the Using Service. If FCG is used to meet the life requirements for cases or mounts, maneuver loads and static loads shall be included in the analysis. Additionally the consequence of failure shall be defined in the analysis and shall not violate any other design requirements (e.g., containment, overboard air leakage).

Physical properties, such as density, thermal expansion coefficient, thermal conductivity, specific heat, elastic modulus, and Poisson’s ratio shall reflect mean values.

3.3.8.6 Strength and Life Analysis. A strength and life analysis shall be performed on primary pressure vessels, disks, vanes, blades, mounts, combustion liners, bearing supports, gears, brackets, and tubing using structural analysis techniques approved by the Using Service. Controls and externals subject to thermal or fatigue loads shall be analyzed. The analysis shall be written in accordance with Appendix E and address the requirements of paragraphs 3.3.8.1, 3.3.8.3, 3.3.8.4, 3.3.8.5, 3.3.8.8, 3.3.8.9.1, 3.3.8.11, 3.3.8.11.1 and 3.3.8.11.2. Required elements of this analysis report are as follows:

a. The report shall thoroughly document the engine configuration, performance and operating conditions analyzed and provide a full description of the design analysis methods employed.

b. The component or engine module functional group discussion shall contain:1. A description of the component, material and component capability.2. A detailed discussion of all analytical and empirical design techniques,

methodologies, assumptions, boundary conditions, and other important analysis factors used to predict life of the component with associated results.

3. Temperatures used with assumed radial and circumferential turbine inlettemperature profiles, any temperature margins assumed to account for fuel control errors, secondary flows and/or disk pumping/recirculations assumed, running clearances, instrumentation errors, engine deterioration and other important performance variables.

4. Types and calculations of time varying loading considered, all assumptionsutilized to establish the applicable mission stresses and temperatures used in the lifing analysis, and a list of all engine operating variables versus time for the analyzed mission(s).

5. All component critical locations including associated failure modes, stresslevels, metal temperatures at critical mission time points, critical modes, excitation sources and other important feature characteristics.

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6. An assessment of blade tolerance to Foreign Object Damage (FOD)(reference 3.2.5.6.2) for all compressor stages and/or erosion. c. The results of component and engine tests used to validate or substantiate the

design relative to the requirements. Unless reviewed and approved by the Using Service, analytically calculated lives shall be presented in the report without adjustment or modification based on these test results.

d. A detailed description of the LCF lifing methodology shall be presented. Thedescription shall include all assumptions, material properties, sensitivity studies, LCF material curves and equations used in life the calculations. Sufficient detail shall be provided so that the lifing methodology can be fully understood by the Using Service. LCF life prediction shall be based on an allowable fatigue curve calculated from mean life to crack initiation reduced by a reliability factor of three standard deviations based on applicable material scatter (3.3.8.5) and shall account for stress concentrations as applicable to the design.

e. The LCF life and its sensitivities to increases in temperature and speed shall bepresented for all critical locations. Details regarding all critical locations and how they were determined shall be presented. Cumulative LCF damage shall be calculated using an appropriate damage theory. LCF component damage fraction tables as per Appendix C, Table C-IX illustrating damage due to major and minor cycles are a requirement and shall be createdand presented in the strength and life report and repeated herein in Table III.

f. Component life predictions required by this paragraph shall be presented for bothSea Level Static (SLS) standard day condition and 6000 feet, 95°F condition.

g. Engine specific points of life (3.3.8.6.1) shall be used in the life analysis. Beforethe Contractor begins any life analyses, the Contractor shall submit to the Using Service for written approval, how all life analyses will incorporate the points of life.

h. Life predictions shall be determined and presented for each potential failure modeto which the particular part is exposed, such as creep/stress rupture, oxidation, corrosion, erosion, HCF and LCF. For these analyses, engine operation shall be according to the design load spectrum and duty/test cycles described in all subordinate paragraphs of 3.3.8.

3.3.8.6.1 Points of Life. The following requirements will be incorporated into the engine lifing analysis:

a. Engine Rating - Engine rating approach will provide a minimum of 54 °F T4.1deterioration margin from the rating table (minimum ship) to field deterioration limit at each rated power except ground idle (GI).

b. Pattern Factor - For creep, LCF, and stress rupture life analyses of staticcomponents, -3σ material properties and a +2 sigma combustor pattern factor (PF) based on statistical analysis of data from combustor rig tests shall be used. For rotating components an average profile pattern factor shall be used.

c. Washout Factor – For washout factor by station, the methodology shall bepresented to and be approved by the Using Service. Mutual concurrence of the methodology is required before use.

d. Performance Lifing Model - For the performance lifing model (LCF, oxidation,stress rupture, creep rupture etc.) the following assumptions shall be used:

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1. Constant T4.1 deterioration model(s) at the agreed upon structural designpoint(s). Approval by the Using Service of the model(s) is required.

2. 2/3rds constant deteriorated T4.1 across power spectrum3. Rotor speeds per the thermodynamic model4. 100% cold starts for LCF life calculations.

e. Ground Idle - GI for life calculations shall be the lowest sustained gas generatorturbine (GGT) speed (reference 3.2.1.5.4.1) corrected to the structural design point(s) and shall accommodate no-load and rotor brake conditions.

f. Hold Time Effects – In the lifing analysis, hold time effects (creep/fatigueinteraction) shall be included on the appropriate components.

g. Temperature Stabilization - The component lives for the mission shall becalculated as defined in 3.3.8.3 and also provide life calculations incorporating thermal stabilization at both the Maximum Rated Power (MRP) and Maximum Continuous Power (MCP) conditions. Stabilization at the GI midpoint is not required. For components where temperature stabilization is not a lifing issue, the contractor shall identify them and not calculate the stabilized life debit. Both sets of lives (if calculated) shall be presented in the Strength and Life Analysis Report. The Using Service will use the approved Strength and Life Analysis Report to determine compliance with the engine specification requirements.

h. Combustor Plenum –Two 0-MRP-0 cycles maybe used to approximate the enginespecification LCF cycle. For limit loads, the maximum possible case pressure loading coupled with maximum maneuver loads shall be used.

i. Gas Path Components Other Than Structural Cases - The impact of pressureloading on gas path components shall be included in their LCF life analysis.

j. GGT and Power Turbine (PT) Vane Creep Life – In determining the creep life ofthe GGT and PT vanes, constant T4.1 deterioration model, shall be used as given in (d) above. The combustor hot streak factor shall be used in a 3D model to determine 1st and 2nd vane creep lives. Engine specification paragraph 3.3.8.1 time at power for creep life determination as well as pattern factor (b) and washout factor (c) above shall be used. A Monte Carlo simulation may be used if the method is reviewed and approved by the Using Service prior to beginning the analysis.

k. GGT and PT Vane Oxidation Life - For oxidation life of GGT vanes, the analysisshall use minimum print coating thickness and average combustor pattern factor. For the PT vanes, the analysis shall use minimum coating thickness and average washed out (station specific) pattern factor. See (c) above.

l. GGT and PT blades - Items (a), (d), (e), (f), (g), and (q) apply. For Oxidation - Theoxidation lives shall assume minimum coating thickness (determined from coating process specification) and average oxidation properties.

m. GGT rotating components - For GGT rotating components, items (a), (d), (e), (f),(g), and (q) apply.

n. PT shafts - For PT shafts, including all splines except the output spline, maximumallowable steady-state torque (mechanical limit per 3.2.1.4.11) shall be used for fatigue life calculations. For the PT shaft output spline the fatigue life calculations shall be done at 2125 SHP.

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o. Rotating components - For life analysis of rotating components, speed overshoots and undershoots shall be included. Overshoot and undershoot values and how they will be used in the life analysis, shall be reviewed and approved by the Using Service.

p. Combustor liner fatigue life analysis - For combustor liner fatigue life analysis, crack growth life may be utilized, but cracks must be limited to not cause detrimental changes in engine performance, pattern factor, radial profile or wall temperature. The LCF analysis-mission defined in specification paragraph 3.3.8.3 and the performance lifing model (item (d) above) shall be used

q. Hot section rotating components – An average radial profile based on available analysis results and/or test data shall be used for hot section rotating components.

r. Localized yielding - Localized yielding not affecting engine operation or the part service limits is permitted.

3.3.8.7 Failsafe Design. The engine shall incorporate a fail-safe design to eliminate catastrophic failure including the following:

a. Where possible, compressor and turbine disks shall be protected by having blades fail first under overspeed or overtemperature malfunctions.

b. A main rotor shaft bearing or lubrication system failure shall not cause parting or decoupling of the shafts.

c. In the event of a rotor bearing failure, the structures supporting the rotating masses shall be designed to minimize the probability of gross misalignment of the engine rotating parts.

d. In the event of shaft decoupling, the requirements of 3.3.8.9.4 shall be met.e. The rotor support system shall allow safe spool down of the rotors after exposure

to ultimate loads. f. A rotating subsystem (i.e. pumps, gears, etc.) failure shall not propagate failure of

attached rotating parts.

3.3.8.8 Creep. All engine parts (static and rotating) shall not creep to the extent that the engine does not meet the requirements of this AQP. Part creep shall not prevent disassembly and reassembly of the engine. Parts shall be able to be reused with used and new hardware. A creep analysis in accordance with 3.3.8.1, 3.3.8.5, 3.3.8.6 and 3.3.8.6.1 shall be written to demonstrate compliance with this paragraph (3.3.8.8). For the power turbine, the creep analysis shall be conducted at 105% speed (22,220 rpm).

3.3.8.9 Containment and Rotor Structural Integrity.

3.3.8.9.1 Containment The engine shall contain a compressor, IPS blower or turbine blade failure occurring at the blade root section in the fillet below the platform that is released within the engine. Blade loss for integrally bladed rotors shall be defined as liberation of the airfoil at the blade root fillet including material down to the rotor rim diameter. The engine shall contain all parts damaged and released by the failure of a single blade. Any post test damage shall not result in an uncontained engine fire at any engine operating speed up to the maximum transient speed limits specified in 3.2.1.4.6.

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Engine controls and externals which contain rotating parts shall contain any failed rotating part at any engine operating speed, including overspeed conditions. See specific alternator containment requirements in 3.7.4.1.1. -3σ material properties per 3.3.8.5 and minimum allowed case thickness above the released blade shall be used for the containment analysis. Spallation or the release of any exterior case or flange material by internal impact without case puncture shall constitute containment failure. External vane actuation rings, levers, tubing, heat shields, cables, wiring or other externally mounted devices shall not be utilized to meet the requirements of this paragraph. The engine design shall be such that failed turbine blades shall not penetrate the air vehicle provided exhaust pipe with a length equal to 1.5 times its diameter and made from AMS 5596 material with a wall thickness of 0.02 in.

3.3.8.9.2 Rotor Integrity. The rotors shall meet the requirements of this AQP and shall remain within functionally acceptable dimensions during and after exposure to the following abnormal conditions:

a. Overspeed1. Gas generator turbine and compressor rotors shall be subjected to 115

percent of the maximum allowable steady state speed limit given in 3.2.1.4.6 and at a maximum allowable gas temperature (given for turbine rotors in 3.2.1.4.5) for five minutes.

2. Power turbine rotors shall be subjected to the greater of the following twomethods for the speed determination and operated at a maximum allowable gas temperature (given for turbine rotors in 3.2.1.4.5) for five minutes:

(a) 115 percent of the transient speed limit given in 3.2.1.4.6.(b) 105 percent of the predicted peak speed attained in a loss of load analysis

per the requirements stated in 3.7.2.3.3.1, conducted throughout the engine operating envelope at worst case ambient temperature and altitude conditions. For calculation of this speed, the analysis shall assume loss of load at MRP redline conditions (min endurance test temperature per Table VI), 100% PT design speed per 3.7.6.8 and worst stackup of overspeed control system tolerances from the loss of load event until the point at which the overspeed device actuates.

b. Overtemperature. Gas temperature at least 45°C (81°F) in excess of themaximum allowable gas temperature given in 3.2.1.4.5 and at a maximum allowable steady state rotor speed (3.2.1.4.6) for five minutes.

3.3.8.9.3 Blade Out.. Subsequent to a single blade failure in any stage with resulting secondary loss of an adjacent blade in that same stage at maximum allowable transient speed, the engine shall meet the containment requirements of 3.3.8.9.1 and shall not cause an uncontained fire; catastrophic failure of the rotor, rotor bearing support, or propulsion system mounts; catastrophic rotor overspeed; leakage from flammable fluid lines, or loss of ability to shut down the engine.

Blade loss loads shall be based on the dynamic imbalance load equivalent to the higher of the load resulting from the loss of two full adjacent blades, or the load resulting from the loss of one full blade and the resulting secondary damage. Blade loss loads shall be combined with the worst case take off loads.

Blade out conditions shall also address the possibility of interactive blade and disk vibration modes resulting from imbalance or acoustics.

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3.3.8.9.4 Disk Burst Speed. The -3σ (per paragraph 3.3.8.5) disk burst speed of the complete disk assembly shall be:

a. For compressor and gas generator rotors, the disk burst speed shall be aminimum of 122 percent of the maximum allowable steady-state rotor speed (3.2.1.4.6) when the disk is subjected to the maximum temperature gradient and maximum material temperature that will occur for that part at any power condition and at the maximum allowable (Min Endurance) temperature listed in Table VI.

b. For power turbine rotors, the disk burst speed margins shall be:1. For engine power conditions of Maximum or below, the PT disk burst

speed using -3σ material properties and material temperatures which will occur for that part using Table VI maximum allowable (Minimum Endurance) temperatures at the maximum power condition shall be at least 5 percent greater than the maximum rated power peak power turbine speed resulting from the analysis prescribed in paragraph 3.7.2.3.3.1.

2. For engine power conditions greater than Maximum, the PT disk burstspeed using -3σ material properties and material temperatures which will occur for that part using Table VI maximum allowable (Minimum Endurance) temperatures at the highest power condition greater than maximum shall be at least 0 percent greater than the contingency rated power peak power turbine speed resulting from the analysis prescribed in paragraph 3.7.2.3.3.1.

3.3.8.10 Vibration and Dynamic Response. The engine shall be free of damaging vibration and dynamic response at all engine speeds and powers including steady state and transient operation throughout the environmental conditions and operating envelope of the engine. All gears shall be free from damaging resonance at all speeds up to 105 percent of maximum allowable steady state speed. The following shall be conducted:

a. A vibration and stress analysis shall be accomplished in accordance withparagraph 3.3.8.10.2.

b. Dynamic models shall be developed for major engine parts and components. Thevibratory stress distribution and the various modes of vibration including complex modes shall be obtained.

c. Acceleration spectrograms at the highest vibration level in the operating envelopeshall be provided at designated engine performance points. The spectrograms are to be generated from each sensor shown on the Engine Configuration and Envelope Figure for engine vibration monitoring. Critical components of the engine shall be identified on each spectrogram. Each spectrogram shall cover the frequency range of 3 Hz to 10 kHz.

d. Bench tests shall be conducted on various parts and components, as appropriate,to verify and correlate with analyses and models to establish material, part or component stress capabilities.

e. A vibration and stress test.

3.3.8.10.1 Critical Speeds. The engine shall be free of resonance conditions at all shaft rotational speeds that would cause the engine not to meet all the requirements of the specification. Critical speeds existing below the engine operating range shall be at least 15 percent below ground idle (reference 3.2.1.4.12

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for the output shaft). Critical shaft speeds existing above the maximum transient speed (3.2.1.4.6) shall be at least 15 percent above that speed. Critical shaft speeds between ground idle and maximum transient speeds shall avoid all steady-state engine operating speeds by at least 15 percent. The natural frequencies of the mounting system with the engine installed shall be at least 15 percent below ground idle shaft rotational speed(s) in all damaging modes of vibration that can be excited by rotor imbalances. Critical speeds between zero and maximum allowable transient speed shall be specified in the engine specification. Adequate damping shall be provided if an engine passes through a critical speed. Predicted critical speeds for the gas generator rotor system are (contractor to specify). Power turbine rotor system predicted critical speeds are (contractor to specify).

3.3.8.10.2 Vibration and Stress Analysis. A vibration and stress analysis shall be conducted on compressor and turbine blades, disks, and vane designs including compressor and turbine shafts and other components where high vibration and stress occur. The report shall be written per the requirements of Appendix F. The vibratory stress distribution and the various modes of vibration including complex modes shall be determined. The critical speeds, excitation frequencies, and stress values for the vibratory stress distributions and nodal patterns shall be determined. The report shall use either singly or jointly the following methods to demonstrate compliance to the requirements:

a. Test data of actual engine specific hardware being qualified.b. Validated analysis of actual engine hardwarec. Similarity analysis relating existing “similar” engine hardware to actual engine

hardware. If this method is selected, the following elements shall be required to show similarity basis:

1. Definition of subject engine hardware and usage.2. Definition of the similar engine hardware and usage.3. Determination and discussion of the differences between subject and

similar engine hardware as to usage, materials, temperatures, stress levels and other important factors.

4. Use validated quantitative methods to reconcile any differences anddemonstrate compliance. Analysis of the data shall include the measured and referred stress values at high stress areas on the cases, blades, vanes, disks, shafts, spacers, engine mounts, and other instrumented parts. Equations and sample calculations for all analytical methods used shall be included in the report. The data shall show the effects on stress levels due to vibration throughout the operating range of the engine. The report shall present modified Goodman and Campbell diagrams for each component analyzed and the Goodman diagrams will show compliance to the -3σ material HCF capability requirements of 3.3.8.2. Plots of excitation frequency versus rotor speed showing the primary orders of excitation and the modes of vibration shall be plotted with points noting stress, measured and referred. A summary of all critical speeds shall be presented in the report.

3.3.8.11 Damage Tolerance. Fracture safety and mission critical engine parts (per 3.3.8.1.2) shall be damage tolerant of defects resulting from material quality, manufacturing processing, and handling damage. The -3σ fatigue crack growth (FCG) life from initial flaw size, as specified in 3.3.8.11.2, to critical crack

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size shall be calculated per 3.3.8.5. The initial flaw shall not grow to critical size and cause failure of the part within the design service life and design usage specified in design service life and design usage specified in 3.3.8.3.

3.3.8.11.1 Residual Strength. The residual strength of all engine components shall be sufficient to withstand the maximum combined loads at the end of its design service life as defined by 3.3.8.1, 3.3.8.2 and 3.3.8.3. Residual strength requirements shall be established for all damage tolerant designed parts and components. Associated static and dynamic loading conditions for these parts and components shall include:

a. Ultimate loading per 3.1.2.5b. Maximum pressure loadingc. Maximum speed loadingd. Maximum temperature effects

Control system overshoot and engine deterioration shall be included in this evaluation.

3.3.8.11.2 Initial Production and In-Service Flaw Size. The assumed minimum initial flaw sizes that exist in a part as a result of material, manufacturing, and processing anomalies as well as handling damage which may have occurred after completion of depot, intermediate, or base level inspection shall be as specified below: (The contractor shall provide the information for the table.)

Inspection Method

Production or In-Service

Material Flaw Type

Flaw Size (depth X length)

Part Description

Location

Assumed initial flaw sizes based on production or in-service non-destructive inspection (NDI) methods shall have a probability of detection (POD) of at least 90 percent with 95 percent confidence level (CL). Assumed initial flaw sizes that are based on the intrinsic material defect distribution, shall encompass 99.9 percent of the defect population when a -3σ residual life is used.

3.3.9 Design Control.

3.3.9.1 Standardization. Standardization principles, standard parts, materials, processes, tools, subsystems, and components shall be used to the maximum extent possible without compromise in design, performance, operability, or economic life of the engine. All parts, materials, and processes, whether or not identified as a Government, industry, or contractor standard shall be qualified for the intended use as a part of the qualification specified herein. Items already in the Government inventory shall be used to the maximum extent possible where suitable for the intended purpose. Variation in similar components or parts shall be held to the absolute minimum. Proprietary designs shall be kept to a minimum. Under conditions wherein economics of production conflict with standardization objectives, the Using Service shall be requested to select the component

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desired for use. Military and industry standards and parts developed for engines in air vehicles shall be used unless they are unsuitable for the intended purpose.

3.3.9.1.1 Not used.

3.3.9.1.2 Not used.

3.3.9.1.3 Not used.

3.3.9.2 Parts List. The parts list for the engine including the software version that successfully completes the designated engine milestone verifications, shall constitute the approved parts list. Changes to the approved parts list shall require approval by the Using Service prior to parts incorporation. Prior to the initiation of Engine Integrity Testing (EIT), Preliminary Flight Rating (PFR), Qualification Test (QT), and Operational Capability Release (OCR), the contractor shall submit a proposed parts list, which shall include hardware and software configuration. Hardware configuration shall be identified by part number and associated drawings. Software configuration shall be identified by a specific source code listing and all associated logic diagrams. The parts list for the engine configuration that successfully completes the EIT, PFR and QT verifications shall be the approved parts list for the respective engine model.

3.3.9.3 Assembly of Components and Parts. Equipment, parts, and components which are not structurally or functionally interchangeable shall not be physically interchangeable. Parts and components shall be designed such that it is physically impossible to install them incorrectly, e.g. backwards, upside down, reversed in an assembly, or installed in the wrong location in an assembly. Connections located in close proximity to each other which are not functionally interchangeable, shall be made physically non-interchangeable. Nonmetallic “O” ring seals and packing shall remain on the part onto which they are installed when the associated mating part is not installed and the associated mating part shall be designed to prevent cutting or other damage to the “O” ring seals and packing.

3.3.9.4 Design and Construction Changes. Changes in any vendor, fabrication process, or fabrication source for any component or part shall be subject to all requirements specified herein. The contractor shall prepare and submit a list of those parts, components, and assemblies that require substantiation tests to qualify an alternative vendor source or process. The specific test(s) required to qualify parts as engine parts shall be defined and submitted with the list. The fabrication source and process of selected vendor components will be included in this list. The contractor shall be responsible for insuring that all parts, components, and assemblies on the substantiation list comply with the qualified fabrication source and process, and that any changes to those sources or processes are effectively controlled. The contractor shall be responsible for performance of the substantiation test to establish satisfactory alternate vendors or fabrication sources or processes. A fabrication source is defined as the prime physical source producing the part, component, or assembly. Changes of fabrication location, such as to another plant of an individual vendor, shall be construed as a change of fabrication source.

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Appendix D

3.3.10 Bearings. All bearing rings shall be positively prevented from rotation relative to the ring housing or shaft, as applicable.

3.4 Not used.

3.5 Logistics.

3.5.1 Maintenance.

3.5.1.1 Modular Concept. Requirement shall be specified in RFP.

3.5.1.2 Borescope. Provisions for 360 degree inspection of the uninstalled engine by borescope shall be made for the compressor, combustor, and turbine sections of the engine. A positive means of slowly rotating the rotor system shall be provided to facilitate inspection. The borescope access port size shall be a minimum diameter of 0.2 in. The same tool shall be used for removal and replacement of port covers, plugs, and associated fasteners. The location of access ports shall, as a minimum, permit inspection of the following locations: (to be contractor provided). Access ports shall be accessible without removing other components

3.6 Not used.

3.7 Major Component Characteristics.

3.7.1 Anti-icing System. The engine anti-icing system, when required, shall prevent the accumulation of ice on any engine part including all engine furnished shafts and housings which protrude into the inlet airstream while operating under the icing conditions specified in 3.2.5.2. The total loss in performance of 5.0 percent specified in 3.2.5.2 shall include the effects of operation in the icing environment plus the effects of operation of the anti-icing system. The effect of anti-icing system operation in a non-icing environment shall be specified in the engine specification. Operation within an icing environment or shedding of ice accretions shall not:

a. Cause damage to any engine components.b. Cause IPS scavenge flow degradation greater than 20% after a ten minute

exposure to the specified icing conditions. c. Cause airflow disturbances that excite harmonic compressor frequencies.d. Cause secondary damage due to reduced clearances between rotating and

stationary components. e. Result in compressor surge or stall.f. Adversely affect the power setting parametersg. Cause engine flame out

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3.7.1.1 Anti-Icing System Activation. Activation of the system shall be accomplished through an airframe supplied command. The anti-icing system shall provide a signal for indication that the anti-icing system is operating.

3.7.1.2 Type of Anti-Icing. The performance and operability effects of the anti-icing system operation in a non-icing environment shall be specified. If failure of the anti-icing system occurs, it shall revert to the anti-icing mode and provide a signal per 3.7.1.1 indicating the anti-icing system is operating. Continuous operation of the anti-icing system throughout the operating envelope shall not damage the engine. The acceleration performance of the engine when using anti-icing bleed air shall be as specified in 3.2.1.5.6. For electrical anti-icing systems, the engine shall be capable of simultaneous operation of the anti-icing system and all other engine electrical systems. If an inlet particle separator (IPS) is used, ice formation on the IPS shall not be permitted in the core engine flow path. Ice accumulation in the scavenge duct shall not affect IPS operation, to include IPS scavenge flow, separation efficiency, and excessive vibration levels.

3.7.2 Engine Control System. The engine control system shall provide complete automatic control of the engine. The operation of the engine control system in normal and emergency modes must minimize pilot workload. Engine control system failures shall not cause an unexpected engine transient or any urgent pilot actions. The control shall annunciate, as appropriate, to the pilot an indication of a malfunction. The automatic system shall be completely self-sufficient and shall require no external power from the airframe electrical system except as permitted by 3.7.4.1. The engine control system shall be fully described and each component and its function shall be listed in the engine specification.

3.7.2.1 Control Signals. The engine control system shall provide for input signals from external sources as required by the Using Service. The engine control system shall also provide for output signals for use external to the engine as required by the Using Service. Control input signal and output signal requirements, input signal and output signal parameters, and related functions shall be as specified below: (contractor to specify)

Input/Output Parameter Function/Purpose Digital transmissions shall be as listed in Appendix (contractor to specify).

3.7.2.1.1 Control Lever Torque. If levers are used, the torque required to rotate any engine control lever through its range of rotation shall not exceed 3 N-m (2.2 lb-ft). Control levers shall not rotate with the engine operating unless external torque is applied. The rotation of the control levers throughout their operating range shall be free of abrupt changes in actuating torque, and the maximum permissible variation shall not exceed 1N-m (0.74 lb-ft). If the torque is different when the engine is not running, this shall be specified in the engine specification. The relationship between control lever torque and control lever angles shall be shown as in figure 15. The maximum allowable static (1g) axial, shear, and overhung moment loads at the control lever connections shall be specified in the engine specification. The maximum allowable loads at these connections shall also be specified for the maximum allowable maneuver loads (combined linear and angular effects) as defined in figure 1. The control lever torsional loading limit between incremental stops and the maximum allowable loading when the control lever is against its travel stops shall be specified in Newton-meters in the engine specification.

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Appendix D

3.7.2.1.2 Control Lever Rigging. If levers are used, a readily visible indexing plate shall be provided at each control lever. Each control lever shall be provided with means to allow positive prevention of incorrect rigging.

3.7.2.2 Engine Control System Performance. The engine control system shall provide control of engine operation to obtain the steady-state and transient engine performance specified herein. The engine control system shall automatically prevent the engine from exceeding any of its limits throughout the environmental conditions and operating envelope of the engine when subjected to unrestricted power demand inputs. Loss of engine supplied electrical power shall not result in an output shaft power change and shall not affect acceleration or deceleration characteristics. The engine control system shall provide the proper relationship between control position and controlled engine variables. The engine control system shall:

a. Automatically match the output shaft torque of all engines in a multi-engine application to within two percent of maximum continuous power rated torque (i.e., the torque difference between engines is no more than 2% MCPQ during steady-state conditions).

b. Allow the selection of any output shaft speed from 90 percent rated output shaft speed to 110 percent rated output shaft speed, inclusive, at all powers equal to and greater than flight idle throughout the engine operating envelope.

c. Maintain actual output shaft speed within 0.5 percent of the selected output shaft speed during steady-state operation when driving a shaft power absorber with characteristics as specified in 3.1.2.14. Control output shaft speed IAW the limits specified in Table XXIII for the defined power transients at all ambient conditions. Required external control inputs shall be specified herein. Interface requirements for these inputs shall be defined in 3.7.4.3.3.

d. Sense (in a multi-engine application) engine power loss and automatically increase power on the other engine(s) to provide power required up to contingency power (one engine inoperative) capability. The control system shall automatically enable an emergency power capability when the engine is operating at defined temperature and/or gas generator speed limits, and excessive rotor droop or droop rate is detected.

3.7.2.2.1 Power Modulation. Power signal demand versus engine power output shall be defined herein, and shall be capable of being fully modulated, free of abrupt changes between idle and maximum power, and essentially linear.

3.7.2.2.2 Region of Control Limiting Functions. Regions of control limiting functions at intermediate and maximum power as a function of altitude and Mach number for hot, standard, and cold ambient conditions, shall be as shown in Figure 16. The limiting values used to establish the various regions shall be specified on the figure.Effects of control tolerance variability shall be specified with the appropriate ranges on the Figure16.

3.7.2.2.3 Control System Reliability. The engine control system shall be designed for inherent reliability in all modes of operation. The control system shall detect failures and accommodate them. For any single failure of input sensors or signals, output effectors or signals, electronic controls, or other electrical portions of

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Appendix D

the engine control system, the engine shall continue normal operation. However, any features or functions not associated with engine power delivery are not required to be operational (e.g., engine monitoring or event recording systems).The engine control system shall not propagate any failures to the aircraft. The system shall have the following capability after the failure of any single mechanical component (of the engine control system) which has probability of failure of 5x10-6 or greater per one hour flight.

a. It shall be possible to modulate the engine’s power between 5.0 percent and 100 percent of the maximum power normally available from sea level to 20000 ft altitude.

b. It shall be possible to start the engine up to the altitude limit specified in 3.2.1.4.4 and operate it thereafter in accordance with 3.7.2.2.3a.

c. It shall be possible to prevent engine operation beyond established limits. Preventative action shall be possible in a period of time such that damage to the engine will not occur.

d. It shall automatically switch to the back-up control mode and shall provide a signal for remote indication of primary control failure.

e. There shall be no overspeed or overtemperature beyond stated transient limits and no burner or compressor instability during or following a control power signal change, or a change of the output shaft load IAW 3.2.1.5.6.

3.7.2.2.3.1 Manual Mode Operation. If the control system has a Manual Mode, the system capability and any resulting limitations on engine steady state or transient performance shall be specified herein.

3.7.2.2.3.2 Engine Operating Mode Selection. Any failure resulting in loss of airframe command inputs (3.7.4.3.3) to the engine shall result in the engine remaining in the mode selected prior to the failure (Stop, Ground, Fly). The engine shall continue to meet all performance requirements, including load sharing, with the exception that rotor transient droop requirements do not have to be met if the control system relies on airframe signals for load anticipation. Any transient droop performance differences shall be quantified in the engine model specification.

3.7.2.2.3.3 Fail Fixed Mode. In the event that the engine control system is unable to control fuel flow, fuel flow shall remain fixed at the last control system commanded value. Design of the fail fixed mode shall ensure that the engine will not experience shutdown, transients, stall, instability, or operation beyond engine limits, at the flight condition (power, velocity, pressure-altitude, and temperature) where fail fixed fuel flow was initiated. If the control system has a manual mode, then the ability to switch fail fixed mode to manual mode shall be available in the cockpit. In all cases, it shall be possible for the pilot to initiate a safe engine shutdown from the fail fixed mode.

3.7.2.2.3.4 Extended Redundancy a. The control shall utilize all available data channels and sensor readings to

minimize the occurrence and severity of degraded operations, to include substituting redundant sensor readings from the engine, from another engine, or from the aircraft, as applicable.

b. The control shall govern, loadshare, and limit the engine to the extent possiblewith the available sensors and actuators before choosing fail fixed operation. Transient schedules and governor gains may be reduced as conditions dictate, while fail fixed is chosen only as a last resort.

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3.7.2.3 Engine Control System Design. All components of the engine control system shall be designed to be accessed for removal and replacement without necessitating removal or displacement of any other component, unless otherwise approved by the Using Service. All control components shall be replaceable without calibration or matching with any other component, and without the need for rigging adjustment (if any). If embedded software models are included in the control system design they shall be subject to approval of the Using Service.

3.7.2.3.1 Engine Control System Adjustment. External adjustments to the controls shall not be required.

3.7.2.3.2 Not used.

3.7.2.3.3 Overspeed Protection.

3.7.2.3.3.1 Power Turbine Overspeed Control System. An engine overspeed detection system shall be provided which takes corrective action when the power turbine rotor system exceeds its maximum allowable transient rotational speed to prevent a destructive overspeed condition. The measurement of the rotor speeds used in this system shall be provided by a device which directly senses rotational speed with no intermediate mechanical devices such a gears, flexible shafts, or clutches. The system shall provide for a ground test procedure to insure proper operation of the system. The system shall perform such that no single failure shall cause inadvertent activation and no single failure shall cause both the loss of power turbine speed governing and the loss of overspeed protection. The speed at which the overspeed device actuates shall be (Contractor to specify)____% (______ rpm) of power turbine speed. It shall be possible to shut down the engine from any power condition even with the loss of the power turbine overspeed protection system. A power turbine overspeed system analysis shall be conducted throughout the envelope to identify the peak power turbine rpm which could be reached following loss of load. This value shall be compared to minimum power turbine disk burst rpm. The analysis shall be conducted at worst case ambient temperature, Mach number and altitude conditions. It shall assume loss of load at maximum rated power redline conditions (min endurance test temperature per Table VI), 100% power turbine design speed per 3.7.6.8, engine performance model approved by the Using Service and worst stackup of overspeed control system tolerances from the loss of load event until the point at which the overspeed device actuates. This peak power turbine speed shall be compared to the power turbine disk burst speed calculated per paragraph 3.3.8.9.4.b.1. The analysis shall be repeated a second time for all the above same conditions except that MRP is replaced with at contingency rated power (one engine inoperative) redline conditions (min endurance test temperature per Table VI). This peak power turbine speed shall be compared to the power turbine disk burst speed calculated per paragraph 3.3.8.9.4.b.2. The shaft horsepower and power turbine acceleration rates, used in the analysis, shall be verified during engine altitude and overspeed control system testing.

3.7.2.3.3.2 Gas Generator Overspeed Control System. If a multi spool gas generator design is used, then a gas generator overspeed system shall be provided which shall take corrective action when the gas generator exceeds its maximum allowable transient rotational speed specified in 3.2.1.4.6 and shall be described herein.

Appendix D

78

3.7.2.3.4 Engine Overtemperature Protection.

3.7.2.3.4.1 Engine Start Overtemperature System. An engine overtemperature feature shall be provided which shall reduce engine fuel flow as the measured gas temperature approaches and exceeds the starting gas temperature limit specified in 3.2.1.4.4. Start temperature limiting shall be deactivated by providing an airframe signal to the electronic control. An output indication shall be provided to the airframe when the overtemperature feature is deactivated.

3.7.2.3.4.2 Steady-State Overtemperature. An engine overtemperature feature shall be provided which shall reduce engine fuel flow to limit the steady-state measured gas temperature to the values specified in paragraph 3.2.1.4.5.b. Steady state temperature limiting shall be capable of being overridden when operating conditions demand power beyond contingency. Each occurrence shall be recorded by the control system.

3.7.2.3.5. Engine Control System Reprogramming. The engine control shall be designed to allow software changes to be made in the field without the need to remove the unit from the engine or airframe. Safety design provisions shall be provided to prevent unauthorized reprogramming. Each software version shall be identified by a unique software version number that shall be verified during the loading process. Notwithstanding the above requirement for re-programmability, neither the unit serial number nor the full authority digital engine control (FADEC) accumulated operating hours shall be changeable once the unit leaves the OEM facility.

3.7.2.3.6 Software Growth Capability. The engine control system shall have software growth capability not less than 25% for each of the benchmarks defined below:

a. Real-time remaining divided by real-time used, considering only foreground tasks.However, in calculating the real-time remaining for foreground tasks, background task completion time shall not be allowed to increase by more than a factor of 4 from the baseline.

b. Memory (program, execution, and non-volatile storage) remaining divided bymemory used for each major function of the engine control system (e.g., engine control, engine monitoring, engine diagnostics, engine lifing).

3.7.2.3.7 Engine Control System Software Development. Software shall be designed, developed and tested in accordance with the general requirements of RTCA/DO-178C Level A. For engine controls, all software must be developed, verified and validated in accordance with the most stringent software level of RTCA/DO-178C determined for the aircraft; typically Level A. If a lower software level (often called development assurance level, or DAL) is proposed, the contractor shall provide rationale to the Using Service, in accordance with SAE ARP4754A, for concurrence. Additionally, IAW the practices herein, the contractor shall provide documentation of:

a. System requirements and validation thereofb. Software requirements and verification thereofc. Requirements flow down and bi-directional traceabilityd. Verification and validation test case coverage analysis

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Appendix D

Note that ‘firmware’ or software that is contained with programmable integrated circuits is covered by this document. The design and development standards for such devices, and their firmware or software, shall be in accordance with RTCA/DO-254.

The contractor shall ensure that its software development processes and procedures meet (at a minimum) the requirements of Level III of the Software Engineering Institute (SEI) Capability Maturity Model (CMM). If SEI Level III CMM equivalence cannot be substantiated, the contractor shall provide a rationale to the Using Service for concurrence.

The requirements for software qualification for preliminary flight rating (4.5.2.5.9) and full qualification (4.6.2.5) are in Appendix G of this AQP.

3.7.2.4 Incident Recorder The Engine Control System shall have an incident recorder to store the values of specified parameters prior to and subsequent to an incident. Incident Detection, Incident Data, and Incident Data Storage and Retrieval are defined below.

3.7.2.4.1 Incident Detection After a successful engine start, an incident is declared if any specified incident conditions are detected by the Engine Control System. The specified conditions include (but are not limited to):

a. Engine flameoutb. Un-commanded accelerationc. Un-commanded decelerationd. Sustained acceleration at the maximum fuel flow limiter settinge. System hard faultf. Rotor droop of more than 5% from the power turbine speed set-pointg. Torque increasing and a torque rate limit exceeded (contractor to define)h. Any parameter value exceeding its limit (Gas Generator Speed, Torque, Turbine

Gas Temperature, Power Turbine Speed, Main Rotor Speed.)

3.7.2.4.2 Incident Data The Engine Control System parameters are to be stored in non-volatile memory prior to and subsequent to an incident. Peak values and duration of exceedances shall be recorded. The engine control system parameters include but are not limited to the following:

a. Gas Generator Speedb. Main Rotor Speedc. Turbine Gas Temperatured. Torquee. Power Turbine Speedf. Calculated Engine Fuel Flowg. Inlet Guide Vane Positionh. Power Lever Anglei. Collective Pitch Positionj. Fuel Control Loopk. Compressor Inlet Temperaturel. Bleed Valve Positionm. Anti-ice Activatedn. Ambient Temperature

Appendix D

80

o. Engine Status Words

3.7.2.4.3 Incident Data Storage The Engine Control System must allocate sufficient non-volatile memory to completely record the data specified in 3.7.2.4.2 for no less than nine (9) separate incident-detections. Incident data shall be recorded in a ring-buffer fashion. 30 seconds of data before and after the incident shall be retained.

3.7.2.4.4 Incident Data Retrieval Incident data in non-volatile memory shall be retrievable in a manner that allows reconstruction of the data relative to the time and type of each incident-detection.

3.7.3 Fuel System.

3.7.3.1 Fuel System Interface.

3.7.3.1.1 Not used

3.7.3.1.2 Maximum Fuel Flow. The maximum flow, including that required for acceleration overshoot requirements throughout the operating envelope, shall be specified herein.

3.7.3.1.3 Fuel Inlet.

3.7.3.1.3.1 Fuel Inlet Dimensions. The location and interface dimensions at the engine fuel inlet connection shall be as shown on the Engine Configuration and Envelope Figure.

3.7.3.1.3.2 Allowable Fuel Inlet Connection Loads. The maximum allowable static (1g) axial, shear, and overhung moment loads on the fuel inlet connection shall be specified in the engine specification. The maximum allowable loads at this connection shall also be specified for the maximum allowable maneuver load (combined linear and angular effects) as defined in Figure 1. Connection shall meet the requirements of this AQP during and after exposure to maximum ultimate loads without fuel leakage.

3.7.3.1.3.3 Fuel Inlet Pressure and Temperature. The maximum and minimum fuel inlet pressures shall be specified in the engine specification. The maximum fuel inlet temperature allowable for continuous operation shall not be less than 71°C (160°F) and shall be specified in the engine specification. Minimum fuel inlet temperature shall be in accordance with 3.2.5.1.

3.7.3.2 Fuels.

3.7.3.2.1 Primary Fuel. The engine shall function as specified throughout its complete environmental and operating envelope for all steady-state and transient operation when using fuel conforming to and having any of the variations in characteristics permitted MIL-DTL-5624U grade JP-5 or NATO Code F-44 and MIL-DTL-83133H grade JP-8/JP-8+100 or NATO Code F-34/F-37/F-35 and ASTM-D1655, Jet A, and Jet A-1 and their international equivalents when properly mixed with fuel

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Appendix D

system icing inhibitor conforming to MIL-DTL-85470B. The engine shall meet all requirements of the specification after transition between any primary fuels.

3.7.3.2.2 Restricted Fuel. The engine shall function as specified from Sea Level to maximum altitude and from -54°C (-65°F) to 29°C (84°F) for all steady-state and transient operation conditions when using fuel conforming to and having any of the variations in characteristics permitted by MIL-DTL-5624U grade JP-4 (or NATO Code F-40) and Jet B and its international equivalent when properly mixed with fuel system icing inhibitor conforming to MIL-DTL-85470B. The engine shall meet all requirements of the specification after transition between restricted and any primary fuel.

3.7.3.2.3 Emergency Fuel. The engine shall function satisfactorily for a period of at least six hours from sea level to 20,000 ft altitude, throughout a range from idle to 90 percent of maximum power when using all the diesel fuels of A-A-52557A and NATO code F-54(OCONUS DF-2) providing the fuel remains below 12 centistokes viscosity and all grades of gasoline specified in ASTM-D-439 and ASTM-D-910. Diesel fuel NATO code F-54(OCONUS DF-2) shall have a sulfur mass of 0.30%, and a kinematic viscosity of at least 6.0 centistokes (9.5 centistokes desired) at 20°C (68°F). Tertiary butyl disulfide will be added to the diesel fuel, if necessary, to obtain the required sulfur content. Any limitations, restrictions, special instructions or maintenance inspections required by emergency fuel operation shall be specified herein.

3.7.3.3 Fuel System Performance.

3.7.3.3.1 Fuel System Calibration Limits. Whenever fuel flow calibrations are required for fuel system components, the applicable limits for this calibration, using test fluid IAW MIL-PRF-7024F, Type II, shall be furnished to the Using Service.

3.7.3.3.2 Fuel Contamination. The engine shall function satisfactorily when using fuel contaminated in any amount up to the extent specified in Table X.

3.7.3.3.3 Fuel System Performance with External Assistance. The engine fuel system shall supply the required amount of fuel at the required pressures for operation of the engine throughout its complete operating envelope including starting with the following conditions at the fuel inlet connection of the engine:

a. Fuel temperature: From a minimum equal to the cold atmospherictemperature of Figure 24, or that temperature corresponding to a fuel viscosity of 12 centistokes, whichever is warmer, to a maximum as specified in 3.7.3.1.3.3.

b. Fuel pressure: From a minimum of 5.1 psi above the true vapor pressure of thefuel to a maximum of 50.8 psi gage (relative to the atmosphere) with a vapor-to-liquid ratio of zero.

3.7.3.3.4 Fuel System Performance with No External Assistance. The engine fuel system shall supply the required amount of fuel at the required pressures for operation of the engine throughout its operating envelope and for starting throughout the starting envelope when the fuel temperature and fuel pressure at the fuel inlet connection are within the following ranges:

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a. Fuel temperature from a minimum equal to the cold atmospheric temperature of Figure 24 or that temperature corresponding to a viscosity of 12 centistokes, whichever is warmer, to a maximum as specified in 3.7.3.1.3.3.

b. Fuel pressure from a maximum of 50.8 psi gage (relative to the atmosphere) to a minimum of:

1. 1.0 psi above the true vapor pressure of the fuel.2. That pressure corresponding to a vapor-to-liquid (V/L) ratio of 1.0.3. Thirty-five percent of atmospheric pressure, or4. 2.2 psi absolute

whichever is the higher for steady-state fuel flows corresponding to engine operation at sea level standard day deteriorated contingency rating and below; or for transient fuel flows required to satisfy minimum engine power transient requirements as specified in 3.2.1.5.6.

3.7.3.3.4.1 Pump Priming and Dry Lift. The fuel pump shall be self priming when subjected to a “dry lift” of at least 6.0 feet using the fuels of 3.7.3.2.1 with a fuel temperature of up to at least 54°C (130°F), using a minimum inside line diameter of 0.625 inches. The initial state is air in line between engine fuel inlet and fuel tank at test onset. Time to reach full prime is 30 seconds (max) at minimum cranking speed. The requirements of this paragraph apply to operation with wetted fuel pump internal surfaces. Wetted is defined as a pump which has been completely filled with fuel and then rotated once in the direction of each fuel inlet and discharge interface such that each uncapped interface is pointed down, allowing all residual fuel to drain from that interface.

Vapor relief provisions shall be incorporated to vent excess vapor during pump priming.

3.7.3.3.5 Fuel System Bubble Ingestion The engine fuel system shall be capable of continuously ingesting slugging flow at its inlet without performance degradation throughout the engine operating envelope. Slugging flow is defined as alternating liquid fuel and vapor (dissolved air and fuel vapor released from the fuel) with a 12-inch length of liquid phase fuel followed by a 12-inch length of vapor only phase fuel, both of the same cross sectional area as the engine fuel inlet.

3.7.3.3.6 Fuel Resistance. The materials and designs used in the engine fuel handling components shall be satisfactory when tested with any of the fuels specified in 3.7.3.2 or MIL-PRF-7024F, TT-S-735A Type I, and TT-S-735A Type III test fluids, when used in any sequence.

3.7.3.4 Fuel Filter. A fuel filter assembly, incorporating a disposable filter element shall be provided as part of the engine. The element shall be made of a material that does not cause fuel contamination due to media migration. The absolute filtration ratings of the filter elements shall be specified in the engine model specification. The filter shall have sufficient capacity to permit a cumulative fuel flow equivalent to a minimum of 12 hours of continuous engine operation at maximum continuous rated power at sea level with fuel at a viscosity of at least three centistokes and contaminated as specified in Table X without being cleaned. The filter assembly shall incorporate a pressure relief bypass and be of a design which will prevent the discharge of filter contaminant through the bypass. The filter assembly shall

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provide electrical indications for remote readout of both actual and impending filter bypass conditions. The electrical indications shall be provided to the engine monitoring system. Filter capacity from impending bypass to actual bypass shall be sufficient to allow at least four hours of operation at maximum continuous rated power at sea level with fuel at a viscosity of at least three centistokes and contaminated as specified in Table X. Provisions to purge the fuel system of air following filter element replacement shall be provided. The filter shall be removable without spilling fuel. Fuel pressure surges during operation shall not cause false activation of the impending or actual bypass indicators.

3.7.3.5 Not used.

3.7.4 Electrical Systems.

3.7.4.1 Electrical Power. The engine shall be capable of providing all electrical power necessary for a successful engine start and continued operation without aircraft power. The compressor speed at which uninterruptable internal electrical power is available shall be specified in the engine specification.. Engine electrical equipment shall be capable of accepting externally supplied airframe electrical power as an alternate power source as defined in MIL-STD-704, A thru F. Transfer between primary and alternate power sources, and vice versa, shall be automatic and shall be completely transparent to engine operation.

3.7.4.1.1 Alternator. The alternator shall not be adversely affected by continuous operation at an alternator speed equivalent to 115 percent maximum allowable engine steady speed per 3.2.1.4.6 under full electrical load, and shall withstand five minutes of operation at an overspeed equivalent to 122 percent maximum allowable engine speed without electrical load. The alternator housing shall completely contain all damage if a mechanical failure should occur when operating at or below maximum transient rotor speed.

3.7.4.2 Alternate Electrical Power. In the event of a failure of the self-contained power system, the engine shall automatically switch to an external power source and shall retain normal operation. Engine electrical equipment, when utilizing external power, shall use power conforming to 3.7.4.3.1. Engine performance and operability shall not be degraded with electrical power conforming to MIL-STD-704, A thru F, with tailoring as follows:

a. Steady state voltages between 12 Vdc and 31.5Vdc conforming to normal,abnormal, emergency and start limits.

b. Normal transients and abnormal overvoltages.c. Interrupts occurring on a single channel shall have no discernible effect on engine

control. d. Interrupts occurring on both channels may result in depowering of the engine

control for the duration of the interrupt. Engine control shall be reestablished within 1second upon reapplication of power.

d. Abnormal undervoltage (below emergency limits) occurring on both channels areallowed to degrade engine performance to failed-fixed requirements. Restoration of electrical

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power (to emergency, normal, or abnormal overvoltage) shall not result in engine transients beyond transitions to new steady-state operating conditions.

3.7.4.3 Electrical Interface.

3.7.4.3.1 External Electrical Power. Engine electrical equipment normally utilizing external power shall operate with 28 volt, direct current power defined in MIL-STD-704, A thru F. Engine electrical power requirements are shown on the engine electrical installation drawing. In the event of loss of externally supplied electrical power, the engine shall accomplish satisfactory air starts and shall operate safely at all engine speeds at or above idle and throughout the complete power range. There shall be no limitation or loss of function caused by the loss of externally supplied electrical power for engine speeds at or above (the compressor speed of 3.7.4.1).

3.7.4.3.2 Electrical Connectors and Cables .Electrical connectors shall meet the requirements of MIL-DTL-38999, Series III or MS3459D, Series III as a minimum. At a temperature of -65°F it shall be possible to flex electrical cable and conductors during routine maintenance without damage to these items, and to connect or disconnect electrical connectors using normal maintenance procedures. Electrical connectors and cables shall have sufficient spare contacts in the connectors and wires in the cables to permit future growth. Electrical connectors and cables shall be repairable.

3.7.4.3.3 Digital Communication Signal Interface. All engine digital communication interface requirements shall be specified herein. The agreed upon data format and timing requirements shall be identified and controlled through the use of an Interface Control Document. This information shall be included as an Appendix (contractor to specify).

3.7.4.4 Electrical and Electronic Equipment. All electrical and electronic equipment requiring periodic or routine checkout shall be located on the engine (or in the nacelle) for easy access and shall be provided with test connections to facilitate checkout without removal from the engine or airframe.

3.7.4.5 Electrical Bonding. The engine shall have a clearly defined electrical bonding interface with the airframe that shall be identified on the engine electrical installation drawing. All electrical and electronic components on the engine shall have a clearly defined electrical bonding interface to the engine structure that shall be identified in the design documentation including drawings. Electrical components shall be grounded or bonded such that a personnel shock hazard does not exist during maintenance activities. Electrical components located in flammable leakage zones shall be grounded or bonded such that an ignition source does not exist under normal operation, shutdown, or failure conditions. Engine supplied electrical components shall be grounded or bonded such that the Electromagnetic Environmental Effects (E3) requirements of this AQP (3.2.8.4) are satisfied throughout the engine service life. The airframe manufacturer shall provide electrical grounding and bonding provisions for any engine-supplied equipment mounted on the aircraft structure (e.g., ECU) and for any airframe supplied equipment mounted on the engine structure (e.g., electric starter).

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3.7.4.6 Ground Isolation. a. The engine engine to aircraft interface shall not require that the aircraft structure be used

as a power return. Power returns shall be assigned to pins in the connector assigned to the aircraft interface. The structure is optionally allowed to be used as a backup power return, in parallel with the connector pins.

b. The structure shall not be used as a signal return. Signal returns are optionally allowedto be galvanically connected to aircraft structure at (at most) one location to create a single-point grounding configuration. The design shall ensure that currents flowing in the aircraft structure (possibly up to 500 amps at 400Hz AC superimposed on 500 Amps DC) won’t find a sneak path onto signal returns sufficiently to cause signal corruption.

3.7.4.7 Potting Compounds. Potting materials in repairable components shall be prohibited unless approved by the Using Service.

3.7.4.8 Airframe Load Impedance Requirement. a. For analog signals, the Engine Control System shall be immune to resistances

greater than 10k Ohm between the connector pins supporting a common function and 40k Ohm between connector pins and airframe structural ground, with pin-to-structure resistance measured separately for each pin. The system shall be tested for airframe resistive loading prior to the endurance test of 4.5.1 and 4.6.1

b. For thermocouple signals, the engine control system shall be immune to 1.5megohms between connector pins supporting the same function, and 1.5 megohms between pins and airframe structural ground, with each pin-to-structure resistance measured separately.

c. For analog signals, the Engine Control System shall be immune to capacitanceless than 0.2 µF between the connector pins and connector shell, measured separately. The system shall be tested for airframe capacitive loading prior to the endurance test of 4.5.1 and 4.6.1.

d. For thermocouple signals, the engine control system shall be immune tocapacitance less than 0.001 µF between each pin and connector shell, measured separately.

e. For digital data links, the control system shall meet the electrical impedancerequirements of the data link specification.

3.7.4.9 Wire Faults

3.7.4.9.1 Wiring Faults – Observed Short Circuit Conditions The following wiring short circuit conditions are considered faults:

a. All permutations of conductor shorts under a common metallic shieldb. Any adjacent connector pinsc. Any two conductors from the same sensor/device connectord. Ground to any single conductore. Any 28 VDC pin and component casef. Any 28 VDC pin to any other conductor pin in the same connector

3.7.4.9.2 Control System Response to Wiring Faults (Afflicted Channel) The electronic engine control system shall:

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a. Not experience permanent damage after any wiring fault as defined in 3.7.4.9.1(a) thru (d)

b. Return to normal operation after the removal of the wiring faults as defined in3.7.4.9.1 (a) thru (d).

c. Not propagate any loss or disruption of any function to any other function withinthe same channel due to a wiring fault as defined in 3.7.4.9.1 (a) thru (d).

3.7.4.9.3 Control System Response to Wiring Faults (Non-Afflicted Channel) The electronic engine control system shall maintain normal operation on opposite channel during and after experiencing any wiring fault defined in 3.7.4.9.1 (a) thru (f).

3.7.5 Ignition System.

3.7.5.1 Ignition System Interface. The interface requirements shall be shown on the engine electrical installation drawing.

3.7.5.2 Ignition System Performance. The ignition system shall function satisfactorily throughout the complete environmental conditions and operating envelope of the engine. The engine ignition system(s) shall ignite the combustor(s) under carbon and water fouling conditions. An external means shall be provided to deactivate the ignition system, and it shall be described in the engine specification. The ignition system and its power source shall be described herein including the ratings, in terms of stored energy level and delivered energy level to each igniter, in joules per spark and frequency of spark. The points in the starting envelope where the minimum stored and delivered energy level to each igniter occurs shall also be specified. If the ignition system is designed for less than continuous duty, any limitations shall be described herein, including maximum allowed continuous ON time for the exciter or igniter life reduction.

3.7.5.3 Ignition System Fouling.

3.7.5.3.1 Carbon Fouling. The spark igniters of the ignition system shall provide sparking performance with spark gaps completely covered or bridged with a material simulating carbon fouling. With the minimum power input of 4.6.2.2.3.3, and the carbon fouling conditions, the sparking rate and energy delivery shall not be less than the minimum design value. An amorphous carbon/oil mixture shall be used for the material simulating carbon fouling.

3.7.5.3.2 Water Fouling. The spark igniters of the ignition system shall provide sparking performance when thoroughly drenched with water. With the minimum power input of 4.6.2.2.3.3, and the water fouling conditions, the sparking rate and energy delivery shall not be less than the minimum design value.

3.7.6 Instrumentation Systems. The range, accuracy, time response and transmission rate, and electrical characteristics for each parameter defined in the following paragraphs shall be as listed in either 3.7.4.3.3 or Table XVI. (contractor to specify)

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3.7.6.1 Instrumentation Interface.

3.7.6.1.1 Condition Indication. The engine shall incorporate sensors (pressure, temperature, speed, torque, quantity, and warning) to provide information to permit safe engine operation within established operating limits.

3.7.6.1.2 Engine Condition Monitoring. The engine shall incorporate equipment or provisions for sensors as required to monitor engine performance and mechanical condition per 3.1.6. 3.7.6.2 Not used.

3.7.6.3 Not used.

3.7.6.4 Temperature Sensing Systems. The temperature sensing systems of the engine shall provide a signal for the aircraft instrumentation system. The output of the sensing devices as a function of temperatures and range of normal operation shall be shown herein. If thermocouples are used, the relationship between temperatures and output signal shall be in accordance with the applicable calibration of National Institute of Standards and Technology Monograph 125. The accuracy of the signals in relation to the actual measured temperature and transient time response characteristics shall be specified. The engine specification shall contain a brief description of each sensing system including circuitry, construction, number of thermocouple or measuring devices, and their locations. Optical or radiation pyrometer temperature sensing systems shall not require removal for re-calibration. The cleaning interval shall be specified herein and shall be no less than 300 hours.

3.7.6.5 Vibration Measurement. The engine shall incorporate provisions (brackets, mountings) for determination of vibration in three mutually perpendicular planes at appropriate locations on the engine cases and accessory gearbox. Additional external or internal locations (e.g., main bearing locations, control system components, accessories) shall be specified and approved by the Using Service for the particular engine design. The points of attachment for the vibration sensors shall be shown on the Engine Configuration and Envelope Figure. Any required signal processing and the output signal characteristics of any vibration sensors that are part of the engine shall be specified.

3.7.6.6 Torque Indication. The engine shall provide a signal for operation of a torque indicator throughout the operating envelope of the engine. The type and range of engine furnished signal and the relationship of torque to signal shall be as specified. The specified accuracy shall be maintained without re-calibration after the torque sensor is replaced. The accuracy of the torque signal, from 50 percent output shaft speed to 110 percent rated output shaft speed, shall be:

a. Within plus or minus two percent of the torque at maximum continuous ratedpower from zero output shaft torque to the torque at maximum continuous rated power.

b. Within plus or minus two percent of the torque being measured, from the torqueat maximum continuous rated power to the transient (0.2 minute) torque limit.

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3.7.6.7 Engine Component Life Counter. The function of the engine life counter shall be accomplished by the engine monitoring system indentified in 3.1.6. The EMS shall document and record data and events that indicate usage of engine or component life as follows:

a. Mechanical and thermal LCF cyclesb. Stress rupture lifec. Engine operating and flight hoursd. Engine startse. Limit exceedancesf. Condition monitoring events

Interface requirements shall be provided in the RFP.

3.7.6.7.1 EMS Algorithms An EMS algorithm Validation, Verification, and Accreditation (VV&A) plan for the development of the engine algorithms shall be developed. This plan shall be described in the engine specification.

3.7.6.8 Speed Indication. The engine shall provide signals for rotor speed (rpm) indication as shown on the Electrical Installation Connection Diagram. For multi-rotor engines, a speed signal shall be provided for each rotor. The speed (rpm) for each rotor at 100 percent speed shall be specified in the engine specification. For the power turbine, 100% speed shall be defined as in 3.2.1.4.6

3.7.7 Engine Lubricating System. The system shall function without degradation throughout its operating envelope and provide windmilling capability per 3.2.1.5.7 without requiring a change in lubricant type or viscosity, except as allowed in 3.7.7.2.1. The complete lubricating system, including oil coolers and oil reservoir, shall be furnished as part of the engine. The engine oil system shall not provide a function for other aircraft accessories or components. The engine lubrication system shall be designed to prevent contamination of the oil by other fluids.

3.7.7.1 Lubricating System Interface.

3.7.7.1.1 Oil System Installation and Servicing. The location and pertinent features of oil system interfaces, such as the oil reservoir filler cap, pressure fill ports, drains, oil level indication, and vents, shall be shown on the Engine Configuration and Envelope Figure.

3.7.7.2 Lubricants.

3.7.7.2.1 Lubricating Oil. The engine shall use lubricating oils conforming to MIL-PRF-7808 and MIL-PRF-23699, and any of the variations in characteristics permitted by the oil specifications. No special provisions such as oil preheaters or oil dilution shall be required for starting and operation throughout the complete environmental conditions and operating envelope for the engine, except that with

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MIL-PRF-23699 oil, operation is not required at oil temperatures below that temperature corresponding to an oil kinematic viscosity of 13,000 centistokes.

3.7.7.3 Lubricating System Performance. The engine shall function satisfactorily throughout the complete environmental conditions and operating envelope of the engine, including any of the flight maneuver forces and attitudes of 3.2.1.5.1 when the oil reservoir contains more than the quantity of oil which is defined as “unusable.” The lubrication system shall provide its function within the oil pressure and temperature limits of 3.2.1.4.8 and shall not exceed the oil consumption rate specified in 3.2.1.4.9.

3.7.7.3.1 Oil Flow and Heat Rejection. The performance of the oil system and associated cooling provisions and cooling requirements, including sample calculations of the oil system heat balance, shall be as furnished in Appendix (contractor to specify).

a. The oil flow, heat rejection data and vent airflow, based upon the maximum limitingtemperatures of 3.1.2.8.1, shall be as furnished in Appendix (contractor to specify). If an oil-to-air heat exchanger is used, the required airflow and pressure drop for all applicable airspeed, altitude, and ambient temperature conditions shall be as specified in Appendix (contractor to specify).

b. When an oil-to-fuel heat exchanger is used, the data shall be based on 15°C(59°F), 57°C (135°F), and maximum fuel temperatures at the inlet connection on the engine together with the minimum fuel flow for each operating condition as would be encountered by an engine having the best possible projected fuel consumption. The values of and methodology for determining maximum and minimum values of fuel temperatures and flows used in the analysis shall be furnished in Appendix (contractor to specify).

c. A complete oil system heat balance shall be presented in Appendix (contractor tospecify) with the maximum limiting temperatures of 3.1.2.8.1 and with 15°C (59°F), 55°C (131°F), and maximum oil cooler cooling medium temperature. The heat balance shall take into consideration effects of maximum accessory gearbox loading, variations in engine internal cooling requirements, bleed air extraction, and the conditions that would impose the greatest cooling requirement on the system. A report verifying the heat balance analysis shall be shown in this engine specification and shall be submitted to the Using Service prior to the initiation of the qualification endurance test. The values of and methodology for determining worst-case conditions in the analysis shall be furnished in Appendix (contractor to specify). The instrumentation requirements and sensor locations for evaluation of oil system performance shall be as specified in Appendix (contractor to specify). Heat balance data shall be furnished in Appendix (contractor to specify) as specified above for idle power, 50 percent maximum continuous power, maximum continuous power, and maximum rated power for the following conditions:

1. Sea level, static, 55°C (131°F) ambient.2. Sea level, static, 15°C (59°F) ambient.3. Sea level, static, -54°C (-65°F) ambient (after 10 hours ground soak).4. 4000 ft altitude, zero flight speed, 45°C (113°F) ambient.5. 6000 ft altitude, zero flight speed, 45°C (113°F) ambient.

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6. Maximum altitude as defined in 3.2.1.4.3, zero flight speed, hotatmospheric conditions as defined in 3.2.1.4.1, and Figures 7 and 8.

7. The engine condition which produces the highest oil temperatures. Thevalues of and methodology for determining this engine condition shall be furnished in Appendix (contractor to specify).

3.7.7.3.2 Internal Oil Leakage. The lubricating system design shall be such that oil leakage within the engine shall not cause oil discharge from the engine upon subsequent starting after shutdown, adversely affect oil supply determination, cause contamination of bleed air, cause residual fires in the engine, cause deposits or produce visible smoke.

3.7.7.3.3 Oil Flow Interruption. The engine shall operate at intermediate power for a period of 30 seconds during which no oil is supplied to the engine oil pump inlet. As a result of this operation, there shall be no detrimental effects to the engine during the oil flow interruption period or during operation thereafter.

3.7.7.3.4 Loss of Oil. The engine oil system shall adequately lubricate the engine for not less than 6 minutes of operation at 75% maximum continuous power after loss of all oil supply to the engine bearing sumps followed by 30 minutes of operation at 75% maximum continuous power with lubrication restored.

3.7.7.4 Lubrication System Components and Features.

3.7.7.4.1 Oil Reservoir. Engine mounted oil reservoirs and mountings shall be constructed of corrosion resistant material capable of withstanding, without permanent deformation, the stresses imposed by reservoir pressurization, engine vibration, and cyclic stresses imposed by variations in ambient pressure and internal reservoir pressure. The oil reservoir shall withstand a differential pressure of 14.7 psid positive and negative or twice the maximum reservoir operating differential pressure within the continuous operating envelope of the engine, whichever is greater, without visible leakage or deformation. The reservoir shall also withstand satisfactorily 15,000 cyclic pressure reversals at the maximum reservoir operating pressure within the operating envelope of the engine without visible leakage or permanent deformation. The oil reservoir shall meet the requirements of 3.3.6.1.

3.7.7.4.1.1 Oil Reservoir Capacity. The total enclosed capacity of the oil reservoir, usable oil volume, gulping oil volume, unusable oil volume, and expansion space shall be specified in the engine specification. The oil reservoir shall contain an expansion space equal to or greater than 20 percent of the total oil quantity of the reservoir. The oil reservoir shall be of sufficient capacity to provide a "usable" oil quantity equal to a minimum of 12 times the maximum hourly oil consumption specified in the engine specification. The oil reservoir shall be of sufficient capacity to insure that with up to a maximum of one-half of the "usable" oil supply depleted, the low level warning will not be activated within the engine operating envelope under the forces and attitudes of figures 1 and 9.

Usable Oil ( quarts) Gulping Volume Oil ( quarts)

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91

Unusable Oil ( quarts) Expansion Space (equivalent oil) ( quarts) Total Enclosed Reservoir Volume (equivalent oil)

( quarts)

3.7.7.4.1.2 Oil Reservoir External Features. The oil reservoir shall contain the following external features necessary to determine the oil level and to service and drain the reservoir:

a. Gravity fill port including a scupper with an overboard drain and a port cap. Thecap shall not exceed a 2-inch diameter and shall be painted yellow. The oil reservoir shall be capable of being continuously filled over an oil tank pitch attitude range of -15° to +15°. The fill-to-spill feature of the tank shall function at all points within the specified fill range.

b. An indicating device on or in the oil reservoir which signifies to the maintenancepersonnel when the tank needs servicing with at least one quart of oil.

c. A low level oil warning device shall be provided and incorporated in the oilreservoir. The device shall furnish a warning signal for a cockpit warning light when a 2 hour minimum of “usable” oil remains in the reservoir over an oil tank pitch attitude range of -15° to +15°.

d. A conveniently accessible self closing drain valve shall be provided for drainingthe oil reservoir.

e. A flapper valve which prevents spillage when the oil tank cap is not properlyattached.

3.7.7.4.2 Oil Drains. Drain ports shall be provided at appropriate low points in the oil system for draining the engine oil while the engine is in attitudes ranging from horizontal to 15° nose up and 20° nose down. The quantity of oil remaining in the oil reservoir after three minutes draining shall be specified herein with the oil temperature at or above 10°C (50°F). A self-closing drain valve shall be provided at an optimum location in the oil system for obtaining representative oil samples for spectrometric analysis. The oil drain shall be constructed to prevent accidental opening during routine maintenance operations or during severe engine vibration cycles. Complete drainage must be possible during varying climatic conditions as specified in figure 24. The provisions for draining the engine oil and the drain locations are shown on the Engine Configuration and Envelope Figure.

3.7.7.4.3 Oil Filter. An oil filter shall be provided in the engine oil system. The type of element and capacity shall be specified herein. The filter shall have a 3 micron filtration rating. The filter assembly shall incorporate a pressure relief bypass and be of a design which will prevent the discharge of filter contaminant through the bypass. The bypass indicators shall provide both local and remote indications and incorporate provisions to preclude activation during low temperature starting. An automatic shutoff device shall be provided to prevent oil drainage when the filter bowl is removed . The filter assembly shall provide an electrical indication t o t h e e n g i n e m o n i t o r i n g s y s t e m . The oil filter element shall have a life of at least 500 hours prior to

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impending bypass indication and at least a duration of two hours at IRP between impending bypass and bypass.

3.7.7.4.4 Oil Debris and Condition Monitor. Oil debris and condition monitor(s) shall be installed in locations where wear particles in the oil would most likely collect and shall detect oil debris and determine fluid condition as specified herein. Oil debris monitors shall be designed for easy access by maintenance personnel and incorporate a checkout feature. The design shall be such as to prevent the loss of oil and introduction of contamination into the oil system when the detector is removed for inspection. Built in test capability shall be required. Electrical power shall be supplied by the engine electrical system. If the monitor indications are processed by the engine control system, they shall be transmitted to the cockpit warning and annunciation (CWA) system via the aircraft data bus. The data format and transmission rate shall be as described in 3.7.4.3.3.

3.7.7.4.5 Oil Coolers. The type and number of oil coolers used in the oil system shall be specified. Oil coolers shall be designed to facilitate internal cleaning and inspection. Coolers shall incorporate a pressure relief bypass valve.

3.7.7.4.6 Not used.

3.7.7.4.7 Oil Pressure Indication. A continuous remote indication of the oil pressure, at the location where the limit of 3.2.1.4.8 is determined, is required.

3.7.8 Not used.

3.7.9 Starting System.

3.7.9.1 Starting System Interface.

3.7.9.1.1 Starting Torque and Speed Requirements. The required starter torque and drive speeds shall be included and presented in the format of figure 17. The figures shall identify engine drag, breakaway torque, engine accessory drag, and accessory gearbox drag. The figures shall show the effects, singly and in combination, of (1) ambient temperatures, (2) altitudes from sea level to the maximum altitude of 3.2.1.4.4, and for air starting with starter assist. The above effects will be for the worst case fuels and oils of the engine specification. Figures shall be presented for each of the following:

a. No customer bleed air extraction, no customer power extraction.b. Maximum allowable customer bleed air extraction, no customer power extraction.c. No customer bleed air extraction, maximum allowable customer power.d. Maximum allowable customer bleed air extraction, maximum allowable customer

power extraction.

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3.7.9.1.2 Moment of Inertia of Rotating Parts. The maximum effective mass moment of inertia (slug-feet squared) of engine parts to be rotated by the starter, referred to the starter drive, and the speed ratio between the starter pad and the driven rotor system shall be specified herein.

3.7.9.1.3 Torsional Spring Constant. The torsional spring constant (lbf-ft per radian) for the engine starting drive system at the starter drive pad shall be specified.

3.7.9.1.4 Starter Train Backlash. The maximum backlash of the starting drive system in radians at the starter drive pad shall be specified.

3.7.9.2 Starting Requirements. Using the starting procedure of 3.7.9.3, the engine shall consistently make satisfactory starts, within the envelope, altitude, attitude, temperature, bleed air and power extraction limits of 3.2.1.4.4, 3.2.1.5.1, and 3.2.5.1. A satisfactory start shall be when the engine rotor is accelerated to idle speed from either rest or windmilling speed when using the procedure specified in 3.7.9.3 provided that:

a. The engine stays within operating limits.b. The total starting time for starts made with no ram, from sea level to the maximum

altitude defined in the RFP shall be equal to or less than those specified in Figure 18. c. The total starting time for air starts is equal to or less than 60 seconds (no power

extraction). Air starts are considered acceptable only if the engine lights-off within 30 seconds and accelerates to idle speed in a total elapsed time that is equal to or less than 60 seconds. Starting time shall be measured from the initiation of the starting sequence to the attainment of a stabilized uncorrected engine idle rotor speed. During air starts the engine rotor load shall be at least 5.0 percent of the total maximum accessory drive load specified for the drives shown in Table V.

d. The minimum assist torque is provided by a starter or ram air, as specified herein.

3.7.9.2.1 Restart Time.The minimum allowable time between shutdown and ground starts or between starting attemptsas determined by engine limitations, shall be specified herein. The times specified shall notexceed 30 seconds after the driven rotor system stops turning.The number of consecutive restart attempts shall be specified by the contractor and is subject to approval by the Using Service.

3.7.9.3 Starting Procedure. The starting procedure shall be simple and shall not require critical timing. For normal starting, after initiation of the starting sequence, the control shall provide for ground and air starting and satisfactory acceleration to stabilized idle operating conditions. The engine shall have the capability of being started with the condition lever in idle position and above and, after initiation of the starting sequence, being accelerated immediately to any selected steady-state operating condition. This shall be accomplished within specified engine operating limits and without adversely affecting the engine durability or structural integrity. During all starting, simultaneous manual operation or actuation of switches or levers or combinations thereof shall not be required.

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3.7.9.4 Starting Drive Train. The engine starting torque acceptance capability shall be specified and shall be at least 3.33 times the starting torque required to provide a 15 second acceleration from start initiation to minimum starter cutout speed, under the conditions specified under 3.7.9.1.1.d and based on sea level static, 15°C (59oF). The weakest part of the starting drive system (shear section, clutch, etc.) shall be external to the engine starting drive train and defined in the starter specification. Failure of the air vehicle provided engine starter shall not result in loss of the remaining gearbox functions. The starter drive pad characteristics shall be included in Table V. The direction of rotation when facing the starting pad on the engine shall be specified in the engine specification. No resonant frequency shall be transmitted to or from the starting drivetrain through the engine interface.

3.7.10 Exhaust Nozzle System.

3.7.10.1 Allowable Exhaust System Connection Loads. A brief description of the exhaust nozzle including the method of attachment shall be included in the engine specification. The engine structural integrity shall not be adversely affected by the engine exhaust nozzle due to resonance or fatigue. The nozzle shall not exhibit any resonance frequencies approaching known excitations outlined in MIL-STD-810G. The maximum allowable shear, axial, and overhung moment loads at the engine exhaust flange shall be specified in the engine specification for static (1g) conditions. The allowable loads at the engine exhaust flange shall also be specified for the maximum allowable maneuver loads as defined in Figure 1 (combined linear and angular effects).

Static (1g) Maximum allowable Maneuver loads

Shear TBD TBD Axial TBD TDB Overhung Moment TBD TBD

Crash Loads Factor 25 As a starting point the minimum Static (1g) capability shall be as follows:

Axial Load 50 pounds Shear Load 50 pounds Overhung Moment 40 pound-feet

3.7.10.2 Exhaust Nozzle. A complete dimensional description of the referee exhaust nozzle that the performance given in this AQP is based on shall be provided in Figure 26.

3.7.10.3 Jet Wake Diagrams. Diagrams showing the temperature and velocity profiles in the jet wake at sea level static standard conditions for idle, intermediate, and maximum power settings shall be IAW Figure 25. Areas hazardous to personnel and equipment shall be appropriately defined and marked on Figure 25.

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3.7.11 Not used.

3.7.12 Not used.

3.7.13 Wash System. An integral wash system shall be provided that shall adequately clean the engine with a minimum of cleaning fluid. The required pressure and flow of the washing medium shall be specified in the engine specification. The washing medium shall conform to MIL-PRF-85704C Type II or Type III. The wash system shall not allow standing water to accumulate in the engine. Wash systeminterface connection shall be MS 33656J.

3.8 Not used.

3.9 Quality Verification.

3.9.1 Engineering Evaluation Tests. Engineering evaluation tests shall be conducted for acquiring data (for safety, installation, maintainability, quality, etc.) demonstrating results to support analyses, and establishing an engine configuration capable of satisfactorily completing the EIT, PFR and QT. The configuration of each test article and its difference from the EIT, PFR or QT engine configuration shall be identified and justified in each test report. The engineering evaluation test reports shall be submitted as required by the applicable requirement or test paragraph for each test, or prior to the completion of EIT, PFR and QT as applicable, if no specific delivery dates are given. All required engineering evaluation tests shall be as specified in Section 4.4.

3.9.2 Engine Integrity Testing (EIT). EIT establishes the acceptability of the engine to continue on into Preliminary Flight Rating Testing. The engine is not required to meet full verification requirements for durability and reliability but is required to be the final production configuration. The acceptability of the engine shall be predicated on the satisfactory completion of all EIT analyses and tests in accordance with Appendix A and written acceptance by the Using Service of all test reports and analyses required for EIT.

3.9.3 Preliminary Flight Rating (PFR). The acceptability of the engine for use in experimental flight testing shall be predicated on the satisfactory completion of all PFR analyses and tests in accordance with Appendix B and written acceptance by the Using Service of all test reports and analyses required for PFR.

3.9.4 Qualification Test (QT) Rating. QT shall be predicated on the satisfactory completion of all tests in accordance with Appendix C and written acceptance by the Using Service of all required qualification test and inspection reports, and analyses. Failure or deficiencies in any of the tests will be considered as a failure to qualify the engine model.

3.9.5 Operational Capability Release (OCR) Test. The OCR Program shall include an Accelerated Endurance Test and two Accelerated Mission Tests. OCR tests shall be predicated on the satisfactory completion of all tests in accordance with Appendix D and acceptance by the Using Service of all tests, reports, analysis, and inspections.

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3.9.7 Engine Development Special Tests. Engine Development Special Tests shall be conducted during the PFR and QT phases of the engine development program. These tests shall be in accordance with 4.9.

3.9.8 Acceptance Test (AT). A test shall be conducted on each engine to be delivered to the Using Service and shall consist of those acceptance test requirements specified in 4.7. Engines submitted for engineering evaluation tests, EIT, PFR or QT need not be subjected to acceptance testing. A detailed Acceptance Test Procedure IAW 4.7 shall be prepared and submitted for approval by the Using Service.

Appendix D

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4 VERIFICATION.

4.1 General. Verification that the engine meets the requirements specified in Section 3 of this AQP shall be by inspection, analysis, similarity analysis to previously certified/qualified components, demonstration, or test and shall be as specified in this section.

4.1.1 Responsibility. The engine contractor is responsible for performing all verification requirements and quality assurance provisions.

4.2 Quality Conformance Inspections. Engines, components, and test apparatus shall be subject to inspection by authorized Government officials who will be given the necessary information and facilities to determine conformance with this AQP.

4.2.1 Quality Evidence. Quality evidence and records thereof shall be maintained as required by the contract.

4.3 Manner of Test and Reporting.

4.3.1 Test Surveillance. Each test, demonstration and teardown inspection described herein shall be subject to witnessing by authorized Government representatives. At convenient times prior to the tests and during teardown inspections, the engine and components shall be examined to determine if they conform to all requirements of the contract and specifications under which they were built. At no time during the EIT, PFR, QT or OCR testing shall any part of the engine or component be disassembled, adjusted, cleaned, replaced, or removed without prior approval of the Government representative.

4.3.2 Test Article Configuration. The configuration of each test article shall be identified by a specific parts list. Germane parts specifically under test as well as all engine parts used for a particular test shall be identified in test report documentation. The configuration of each test article and its differences from the EIT/PFR or QT (reference parts list of 3.3.9.2) endurance engine configuration shall be identified and justified in each detailed test procedure. A list of Material Review Board (MRB) approvals with details of approved nonconformances shall be provided in each test procedure. Design corrections and improvements as substantiated by development test are allowed in the identification of parts differences for each test engine; however, the mixing of parts of the same or different design, such as blades in a disk or the mixing of different vendor’s components and parts in a multiple assembly such as a segmented stator assembly, is not allowed. All parts shall be considered as having zero time at the start of a test. While all parts do not have to be new, any part which fails during the test shall be cause for rejection of that test article. The parts list for the engine which successfully completes the EIT, PFR or QT tests shall constitute the approved parts list for the respective EIT, PFR or QT engine model.

4.3.2.1 Test Engines. The particular engine intended for a specific test or demonstration shall be officially designated by the contractor prior to the start of testing.

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4.3.2.2 Method of Qualification. The requirements for Engine Integrity Testing (EIT) are in Appendix A, the requirement for Preliminary Flight Rating (PFR are in Appendix B, the requirements for a Qualification Rating (QT) are in Appendix C and the requirements for Operation Capability Release (OCR) are in Appendix D. The method of meeting the requirements of Appendix A, B, C or D are by test, similarity analysis to previously qualified components, similarity analysis to previously certified components in conjunction with additional analysis to substantiate the difference between the military test requirement and the commercial test requirement, field experience on same or similar components in conjunction with analysis to substantiate the design differences or by other analyses (such as a finite element analysis) that shows substantiation of the design relative to the test or design requirements. See 3.3.8.10.2.c for the requirements of a similarity analysis. Tests that cannot be met by similarity or analysis include the 150 hour endurance test (paragraph 4.5.1), the altitude test (paragraph 4.5.3), the low cycle fatigue engine test (paragraph 4.5.4.8), the electromagnetic environmental effects test (paragraph 4.5.5), the 300 hour endurance test (paragraph 4.6.1), the altitude test (paragraph 4.6.3), the low cycle fatigue engine test (paragraph 4.6.6.2.2), the electromagnetic environmental effects test (paragraph 4.6.4.14) , the Accelerated Endurance Test (paragraph 4.8.1) and the Accelerated Mission Test (paragraph 4.8.2).

4.3.3 Test Apparatus.

4.3.3.1 Automatic Recording Equipment. Automatic recording equipment shall be used to record data during the execution of those parts of the test requiring the evaluation of time versus engine variables.

4.3.3.2 Vibration Measuring Equipment and Response Characteristics. The engine vibration shall be measured with sensors. The vibration measurement and analysis equipment shall operate over a frequency band of at least 3 Hz to 5 kHz and produce acceleration spectrograms having a demonstrated accuracy with confidence level of 95 percent. The maximum allowable effective filter bandwidth of the spectrum analysis equipment shall be 3 Hz up to 1000 Hz and 6 Hz above 1000 Hz. The vibration measuring equipment shall be calibrated as a complete system. The frequency response of the system, when calibrated by applying a known sinusoidal motion to the sensor, shall not deviate by more than ±3.0 dB from the known input at frequencies from 3 Hz to 10 kHz. Filters are required when measuring overall velocity levels, they shall be not more than 3.0 dB down at the cutoff frequency which shall be 30, 70, or 100 Hz high pass, as appropriate, and they shall have a roll-off rate of at least 18 dB per octave. High pass filters shall not be used to produce velocity and acceleration spectrograms.

4.3.3.3 Test Stand Characteristics. Engines shall be subjected to qualification and acceptance testing in a test cell and with test equipment that is acceptable to the Using Service.

4.3.3.3.1 Dynamic Characteristics. Vibration shall be measured with the engine operating on a test stand which has natural frequencies with the engine installed no higher than 50 percent of the idle rotor speed in all modes of motion which can be excited by residual rotor unbalances. The test stand shall not induce damaging or detrimental resonance into the engine at any test or operating condition. 4.3.3.3.2 Power Absorption Characteristics.

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To provide realistic engine accelerations and decelerations, the engine power absorber shall have a polar mass moment of inertia (referred to the output shaft speed) within ±5% of the UH-60M referred rotor inertia for all engine qualification tests in paragraphs 4.5, 4.6 and 4.8 and subsequent subparagraphs.

4.3.3.4 Starter. During EIT/PFR and QT engine tests, starting shall be performed with a starter that has torque characteristics within 5.0 percent of the minimum required torque shown in Figure 17 during EIT/PFR Tests specified in 4.5.1, 4.5.3 and 4.5.4.8; and QT Tests specified in 4.6.1, 4.6.3, 4.6.4.1, 4.6.5.1, 4.6.5.5, 4.6.6.2.2 and 4.6.6.6.

4.3.4 Test Condition.

4.3.4.1 Servicing.

4.3.4.1.1 Oil Servicing. The oil system shall be drained and filled with new oil at the start of the specific engine test or demonstration. Oil used for engine testing shall be Government Furnished Material (GFM) and conform to paragraph 3.7.7.2.1 of this AQP. All additions of oil added after initial servicing shall be recorded. Oil shall be drained from the system and documented in the engine test log for Government surveillance. The use of external oil filters shall not be permitted for special development or qualification tests. Oil consumption shall be determined and reported after a specific number of cycles or hours as approved by the Using Service. Oil shall be added on an as required basis as indicated by the oil tank sight glass or low oil sensor.

4.3.4.2 Inlet and Exhaust Duct Connections. During the test of 4.6.1 and 4.6.6.6.6 the inlet and exhaust duct connections shall be loaded as specified in 3.1.2.10.2 and 3.7.10.1.

4.3.4.3 Bleed Air Connections. During the tests of 4.6.1 and 4.6.6.6.6, the bleed connection loads shall be as specified in 3.1.2.11.1.

4.3.4.4 Accessory Drive Gearboxes. During the tests of 4.5.1.3 and 4.6.1.3, the engine gearbox shall be loaded to the rated loads and overhung moments specified in 3.1.2.7.

4.3.4.5 Accreditable Test Time. Test time shall not be credited by increments shorter than fifteen minutes except when shorter periods are a test requirement.

4.3.4.6 Fuel Properties for Test. The fuels used for the EIT and QT shall be in accordance with grades JP-4 and JP-5 of MIL-DTL-5624U and JP-8 of MIL-DTL-83133H, except when otherwise noted in a particular test paragraph.

4.3.5 Data. Data shall be submitted or recorded during tests IAW the following subparagraphs.

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4.3.5.1 Test Plans and Procedures.

4.3.5.1.1 Engine and Engine Subassembles. The following shall be submitted to the Using Service for approval prior to initiation of each test or demonstration required in 4.4, 4.5, 4.6, 4.8 and 4.9 which is to be conducted on a specific test engine, or subassembly. One document for each test is required. Test plans shall include but not be limited to:

a. The detailed test plan to be used to define the conduct of the test. The test planshall contain the information specified in Table XXI as a minimum.

b. An appropriate configuration identification of the engine or assembly IAW 4.3.2.c. Service and wear limits. Service limits to be used by maintenance personnel in

the field and repair facilities (for OCR only). d. Pretest dimensional inspection results for selected hardware.

4.3.5.1.2 Components. For each component identified in Tables XX-A and XX-B, a single test plan shall be prepared in accordance with Parts 1 and 2 of Table XXV. Part 1 shall be submitted to the Using Service for approval prior to preparation and/or submittal of any Part 2 individual test procedures. Once the Part 1 summary is approved by the Using Service, individual test procedures shall be prepared in accordance with Part 2 and submitted for approval. Upon approval, the individual test procedures shall be inserted within the previously approved Part 1 summary as uniquely identifiable test procedure subsections.

a. One test plan, in accordance with the requirements above, shall be submitted forthe EIT group of tests specified in paragraph 4.5.2.5, excluding the software verification tests of 4.5.2.5.8.

b. One test plan, in accordance with the requirements above, shall be submitted forthe QT group of tests specified in paragraphs 4.6.2.2, 4.6.2.3, 4.6.2.4 and 4.6.2.5.

4.3.5.2 Preliminary Data. The dry engine weight, center of gravity location, photographs, and other pertinent data shall be obtained and recorded at the time the engine is being prepared for test. The weight shall be measured before the engine has been serviced with fuel or oil. These requirements apply to the 4.5.1, 4.6.1, 4.8.1, and 4.8.2 engine tests only. This data shall be presented in the respective engine test reports IAW 4.3.6.2.

4.3.5.3 Accuracy of Data. For all engine and component calibrations and tests, or demonstrations, reported data shall have a steady state accuracy within the tolerances shown below. The accuracy of transient data and the corresponding instrument calibration methods shall be subject to the approval of the Using Service and shall be described in the test plans. All instruments and equipment shall be calibrated as necessary to ensure that the required degree of accuracy is maintained. No allowance shall be permitted for measurement uncertainty to establish compliance with the requirements of this AQP (i.e., all measured values must be increased by the maximum specified measurement inaccuracy of the test instrumentation. When engine tests are performed, the measurement uncertainty errors shall be minimized. The instrument calibration methods shall be subject to Using Service approval prior to the initiation of the test. The “as tested” accuracy as defined by the test facility shall be included in the test report.

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DATA ITEM TOLERANCE Rotor speed(s) +0.2 percent of the value obtained at maximum ratingFuel Flow +0.5 percent of the value measured for maximum continuous

rating and above+0.5 percent of the value measured at maximum continuousrating for all values below intermediate rating

Torque +0.5 percent of the value measured for maximum continuousrating and above+0.5 percent of the value measured maximum continuous ratingfor all values below intermediate rating

Airflow +1.0 percent of the value measured for maximum continuousrating and above+1.0 percent of the value measured at maximum continuousrating for all values below intermediate rating

Temperatures ±1.0°C (1.8°F) up to 200°C (392°F) ±3.0°C (5.4°F) above 200°C (392°F) ±4.0°C (7.2°F) above 800°C (1472°F)

Engine Weight +1 lb or +0.1 percent of the weight being determined, whicheveris greater

Vibration Velocity +5.0 percent of specified engine limit during 4.4.5 vibrationsurvey, 4.6.2.4.2 vibration scan and resonant search, and4.6.6.6.6 engine vibration survey

Vibration Velocity +10.0 percent of specified engine limit for all other testsAll other data +2.0 percent of the value obtained at maximum rating

4.3.5.3.1 Measurement Uncertainty Analysis. A detailed measurement uncertainty analysis addressing all elemental precision and bias errors shall be performed in accordance with ASME PTC 19.1 or equivalent approved by the Using Service. The analysis will cover the three categories of errors: calibration errors, data acquisition errors, and data reduction errors. Analysis will be performed for the following measured parameters: airflow, fuel flow, torque, temperature, pressure, rotor speeds, SFC, output power, engine weight, vibration, barometric pressure, relative humidity, fuel lower heating value, and fuel specific gravity. The Using Service reserves the right to add additional measured parameters that require uncertainty analysis.

4.3.5.4 Steady-State Data. During operation at each specified steady state condition and after performance stabilization, applicable data per Table XII shall be recorded as specified in the test plan.

4.3.5.5 Transient Data. For each transient performed during the power transient operations, applicable data per Table XII shall be recorded as specified in the Test Plan.

4.3.5.6 Starting Data. During each start, applicable data per Table XII shall be recorded as specified by the Test Plan.

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4.3.5.7 Miscellaneous Data. The date, test title, engine model designation, and serial number shall be recorded on each log sheet.

4.3.5.8 Test Notes. Notes shall be placed on the log sheets of all incidents of the run, such as leaks, vibrations, and other irregular functioning of the engine or test equipment and any corrective measures taken or investigated. The log sheets shall also record start count, added oil, software faults or anomalies, and test plan deviations (approved by the Using Service). All test article software changes shall be recorded on the log sheets.

4.3.5.9 Barometer Reading. The barometer reading shall be corrected for temperature and shall be read and recorded with ambient air temperature at intervals not exceeding one hour, if required for engine performance calculations.

4.3.5.10 Relative Humidity Data. Water vapor pressure readings shall be taken at intervals not exceeding one hour, if required for engine performance calculations. The relative humidity shall be measured in the vicinity of the test article.

4.3.5.11 Fuel and Oil Data. Samples of the fuel and oil shall be taken at the start and completion of the test of 4.5.1 and 4.6.1 and for other tests as applicable. The fuel and oil samples shall be analyzed for physical and chemical properties including, but not limited to, fuel lower heating value and specific gravity to determine conformance with the applicable fuel and oil specifications. The results of these analyses shall be included as part of the applicable test reports. Additives to fuels and oils shall be specified in the test plan and data.

4.3.6 Reports.

4.3.6.1 Test Reports. Following the completion of each separate engine or component test, demonstration or consecutive group of tests or demonstrations conducted on any single assembly or component, a report shall be submitted. These contractor reports shall be submitted to the using service for review and approval. Each report shall contain but is not limited to the following items:

a. Cover (title number and source of report, date, name(s) of the author(s), andcontract number).

b. Title page (title number and source of the report, date, name(s) of the author(s),and contract number).

c. Abstract (a brief statement of the contents of the report, including the objective).d. Table of contents.e. List of illustrations providing figure numbers and captions of all illustrations.

Photographs, charts, and graphs shall be treated as illustrations and given figure numbers. List of tables with associated captions. When used in a separate series, tables shall be given Roman numerals. Examples: Figure 1, Figure 2, etc., Table I, Table II, and Table III).

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f. Summary (brief resume of the test conducted, including objective, procedure, results, conclusions, and recommendations referencing the applicable paragraph of the engine specification).

g. Body of the report.1. Brief general description of the engine or of the component and a detailed

description of all parts which differ from the approved parts list, if applicable. 2. List of all germane hardware and serial numbers, if applicable, specifically

under test and a list of all engine parts (not necessarily germane to the test) which were utilized. 3. List of all reference documents relied upon or referenced within the report

body. (Report/document number, title, author and publication date). 4. If approval is being requested, without test, based on similarity to a

component or assembly for which previous test approval was obtained, the requirements of paragraph 3.3.8.10.2.c shall be met including, but not limited to any physical or functional dissimilarities or differences in testing requirements with respect to the tested component and reference to the approved component test report shall be included. Rationale as to test similarity shall be provided.

5. Method of test (general description of test facility, equipment and methodsused in conducting the test).

6. Record of test (chronological history of events and incidents in connectionwith all of the testing, including details of all leaks, vibrations and other irregular functioning, and any adjustments, repairs, replacements of parts and the corresponding engine operating time).

7. Analysis of results (a complete discussion of all phases of the test, suchas probable reason for failure or unusual wear, comparison in performance with previous models, analysis of general operation, and any items that are significant from an engineering viewpoint).

8. Calibrations and recalibration data, for engine, assemblies andcomponents, including acceptance limits. (Data in uncorrected form and corrected form, if applicable, shall be provided in tabular form and shown by suitable curves. A plot of vibration characteristics in three planes shall be provided for both calibration and recalibration runs.)

9. For the altitude tests specified in 4.5.3 and 4.6.3, the performance of 3.2.1and 3.2.1.2 shall be verified.

10. Tabulated data for all pertinent instrument readings and requiredinstrument readings taken during the test.

11. Description of the condition of the engine or components at disassemblyinspection. A completed disassembly inspection and material discrepancy description shall be provided for each discrepancy.

12. Conclusions and recommendations, with respect to approval of the engineor components tests, supplemented by such discussion as is necessary for their justification.

13. Appendix (final approved test procedure).

4.3.6.2 Summary Reports. Following completion of the engine integrity test requirements and qualification requirements specified herein, a summary report for each shall be prepared. These reports shall contain the following items:

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a. Cover (title of report, number of the report, source of report, date, security markings).

b. Title page (title of the report, number of the report, source of report, name of authors, contract number).

c. Abstract (a brief statement of the contents of the report, including the objective).d. Table of contents.e. Summary (A brief resume and summary of each of the tests conducted, giving the

title of each test, test report number, the items tested, dates of testing, and a general statement of the results).

f. Conclusions and recommendations.

4.4 Engineering Evaluation Tests.

4.4.1 Customer Bleed Air. An engine, having substantially the same parts list and configuration as the endurance test engine, shall be subjected to a bleed air test to verify the total pressure, temperature, and bleed air flows available in accordance with 3.1.2.11 to be furnished for aircraft system use. Verifications at sea level from idle to maximum power for inlet air temperatures from -65oF to 131oF (a minimum of five test temperatures) shall be accomplished. The test is to be conducted in such a manner as to demonstrate the amount of bleed air available over and above that required for engine systems such as the acceleration and engine anti-icing systems. During the above verifications, customer bleed air shall be analyzed to determine whether the contaminant levels are within the limits specified in 3.1.2.11.3, for at least four power settings from idle to maximum. The contaminant verification is required at only one inlet air temperature.

4.4.2 Engine Heat Rejection and Oil Cooling. Engine heat rejection and cooling requirements data, including the oil system, will be obtained from engines having substantially the same parts list and configuration as the EIT/PFR endurance test engine. The parts list and configuration with differences identified shall be provided to the Using Service. This data shall include cooling requirements, heat rejection rates, and corresponding surface temperatures for various engine components and stations. The data necessary to define the installed cooling requirements shall be obtained for various engine operating conditions, customer power extraction, and bleed air conditions throughout the environmental conditions and engine operating envelope. The data shall verify the requirements of 3.1.2.8.2 and 3.7.7.3.1.

4.4.3 Oil Flow Interruption Test. An engine, having substantially the same parts list and configuration as the PFR endurance test engine, shall be subjected to an oil flow interruption test. The parts list and configuration with differences identified shall be provided to the Using Service. The engine shall be operated at the intermediate power setting for 30 seconds with only air supplied to the inlet of the oil pump. The engine shall operate without damage during the oil flow interruption period and for 30 minutes thereafter with the normal lubrication having been restored. The engine shall be disassembled and undergo a dirty and clean inspection to ensure that no damage has occurred.

4.4.4 Engine Electrical Power Failure Tests. Electrical power tests shall be conducted to substantiate compliance with 3.7.4.1 and 3.7.4.2.

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4.4.4.1 Engine System Electrical Power Tests. An engine, having substantially the same parts list and configuration as the PFR endurance test engine, shall be subjected to the following tests to demonstrate compliance with 3.7.2.2, 3.7.4.2, and 3.7.4.3.1. A report of the test results shall be submitted to the Using Service prior to completion of the EIT.

a. Operate the engine at intermediate power for five minutes, disconnect the engine supplied electrical power to the control system. The engine shall not exhibit any discernible change in output shaft power or speed. After the engine has run an additional five minutes, decelerate to maximum continuous power. The engine shall not exceed any of its limits nor exhibit any discernible performance or operability changes. There shall be no loss of control system functionality, of any kind, throughout the test.

b. Disconnect engine supplied electrical power to the control at the most critical point of an acceleration. During the time the electrical power is off and during the switching, the engine shall not exceed any of its limits nor exhibit any discernible performance or operability changes. The engine shall complete the transient satisfactorily.

c. Disconnect engine supplied electrical power to the control at the most critical point of a deceleration. During the time the electrical power is off and during the switching, the engine shall not exceed any of its limits nor exhibit any discernible performance or operability changes. The engine shall complete the transient satisfactorily. There shall be no loss of control system functionality, of any kind, throughout the test.

d. With the engine operating at maximum continuous power, switch off the normally supplied external. The engine shall remain within the stability requirements of 3.2.1.5.5 and shall have no loss of control system functionality, of any kind, throughout the test.

e. After the engine has been shut down, and with the engine supplied electrical power disconnected and the normally supplied external electrical power on, restart and accelerate to idle. Accelerate to intermediate power and operate there for five minutes. Decelerate to idle, operate at idle for two minutes and shut down. During this portion of the test the engine shall not exhibit any discernible performance or operability changes. There shall be no loss of control system functionality, of any kind, throughout the test.

f. With the engine supplied power operational, perform the following 50ms interrupts to the normally supplied external electrical power source; first on each control channel separately, then on all control channels simultaneously:

1. Engine steady state at maximum power.2. Engine steady-state at intermediate power.3. Engine steady-state at maximum continuous power.4. Engine steady-state at flight idle.5. Engine steady-state at ground idle.6. During engine acceleration.7. During engine deceleration.

There shall be no discernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test.

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g. With the engine supplied power disconnected, repeat tests f (1) thru f (6) with 50ms interrupts to the normally supplied external electrical power source, first on each control channel separately, then on all control channels simultaneously. During single channel interrupt testing, there shall be no discernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test, when the interrupted channel is the standby control. Interrupts of the controlling channel shall not result in unsatisfactory engine transients or control system behavior (as determined by the Using Service) during the interrupt and recovery period. However, channel switchover to the standby channel is allowable. Simultaneous channel interrupt testing shall not result in unsatisfactory engine transients or control system behavior (as determined by the Using Service) during the interrupt and recovery period. Engine fail fix operation or engine shutdown is allowable, dependent upon control system architecture.

4.4.4.2 Alternate Electrical Power Tests. The electronic control system, with engine supplied electrical power disconnected, shall be tested as follows. During all tests, the engine shall be at maximum continuous power.

a. Steady State Voltage. Simultaneously on all electrical power sources from theaircraft: 12.0 Vdc for 5 minutes minimum, and then 31.5 Vdc for 5 minutes minimum. There shall be no discernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test.

b. Transient Voltage. Simultaneously on all electrical power sources from the aircraft:1. Start at 29 Vdc maximum, pulse to 50 Vdc minimum for 12.5 ms minimum,

end at 29 Vdc maximum. 2. Start at 29 Vdc maximum, pulse to 38 Vdc minimum for 50 ms minimum,

end at 29 Vdc maximum. There shall be no discernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test.

c. Transfer Interrupts. Durations of 1ms, 2ms, 4ms, 6ms, 8ms, 10ms, 12ms, 14ms,16ms, 20ms, 30ms, 50ms:

1. Single Channel. During single channel interrupt testing, there shall be nodiscernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test, when the interrupted channel is the standby control. Interrupts of the controlling channel shall not result in unsatisfactory engine transients or control system behavior (as determined by the Using Service) during the interrupt and recovery period. However, channel switchover to the standby channel is allowable. The interrupted channel shall be fully functional within 500ms after power interrupt.

2. All Channels. Simultaneous channel interrupt testing shall not result inunsatisfactory engine transients or control system behavior (as determined by the Using Service) during the interrupt and recovery period. Engine fail fix operation or engine shutdown is allowable, dependent upon control system architecture. Interrupted channels shall be fully functional within 500ms after power interrupt.

d. Long-Duration Interrupts. Durations of (100ms, 200ms, 500ms, 1sec, 2sec, 5sec,10sec):

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1. Single Channel. During single channel interrupt testing, there shall be no discernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test, when the interrupted channel is the standby control. Interrupts of the controlling channel shall not result in unsatisfactory engine transients or control system behavior (as determined by the Using Service) during the interrupt and recovery period. However, channel switchover to the standby channel is allowable. The interrupted channel shall be fully functional within 500ms after power interrupt.

2. All Channels. Simultaneous channel interrupt testing shall not result in unsatisfactory engine transients or control system behavior (as determined by the Using Service) during the interrupt and recovery period. Engine fail fix operation or engine shutdown is allowable, dependent upon control system architecture. Interrupted channels shall be fully functional within 500ms after power interrupt.

e. Voltage Spikes. One channel at a time. There shall be no discernible changes in engine performance or operability nor loss of control system functionality, of any kind, throughout the test. Duration of 50μsec, amplitude (deviation from steady state) of +31.5V (positive spike), -63V (negative spike) or 10 A magnitude, whichever comes first (voltage or current); 50 spikes in each polarity. Allowable substitute test methods (any of the following):

1. RS06 of DO-160G, Section 17, Category A using above limits at 50μsec, or Spike #1 limited to 10 A maximum amplitude.

f. Ripple Voltage. Simultaneously on all electrical power sources from the aircraft. There shall be no discernible changes in engine performance or operability or loss of control system functionality, of any kind, throughout the test.

1. Frequency Scan. 10 Hz to 20 kHz, scan rate per MIL-STD-461F paragraph 4.3.10.4.1 and Table III (extended down to 10 Hz), amplitude using the guidance of MIL-HDBK-704-8, worst case.

2. Complex Ripple Voltage Waveform. 1.5Vrms AC-component, 2400 Hz dominant waveform frequency (zero-crossing about mean), and a fundamental down to 400 Hz, with a minimum of 9 separate harmonics with minimum amplitudes using the guidance of MIL-HDBK-704-8, worst case.

4.4.5 Engine Vibration Survey. A vibration survey which demonstrates compliance with 3.2.1.4.10 and 3.3.8.10 shall be conducted. The vibration survey shall include, but not necessarily be limited to, data showing true RMS velocity spectrograms and peak acceleration spectrograms for each sensor location at the highest vibration point in the operating envelope (which shall be identified) and at designated engine rating points. The spectrograms shall cover the frequency range of 3 Hz to 5 kHz. The method used for determining the overall true RMS velocity from the spectrogram and the maximum permissible overall true RMS velocity limit shall be described. The test stand shall comply with 4.3.3.3.

4.4.6 Starting Torque. Prior to start of PFR an engine shall be tested to demonstrate compliance with the starting torque and speed requirements of 3.7.9.1.1. The procedure for accomplishing this demonstration shall be submitted to the Using Service for approval prior to initiation of test.

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4.4.7 Not used.

4.4.8 Maintenance Test. See RFP for test requirements.

4.4.9 Not used.

4.4.10 Verification of Correction Factors. When correction factors or engine models are used to convert engine performance data to standard conditions, sea level and altitude tests shall be conducted over the necessary range of environmental conditions as defined in 3.2.1.4.1 to verify the correction factors or engine models used to correct performance data.

4.5 Engine Integrity Testing (EIT)/Preliminary Flight Rating (PFR).

4.5.1 Endurance Test

4.5.1.1 Pretest Verification.

4.5.1.1.1 Engine Dry Weight. Prior to initiation of calibrations, the dry weight of the engine, as specified in 3.2.2.1 shall be verified in accordance with 4.3.5.2.

4.5.1.2 Calibration.

4.5.1.2.1 Component Calibration. Functional bench calibration/checks shall be conducted on each of the components specified in Table XVIII. All fuel nozzles and fuel carrying components of the engine control system shall undergo bench calibrations using fluid in accordance with 3.7.3.3.1. Components shall conform to the design tolerance range required by the applicable design specifications. All control system sensors shall be calibrated. The temperature sensing system performance shall meet the tolerance and thermal response characteristics of 3.7.6.4.

4.5.1.2.2 Engine Calibration. The procedure during the engine calibration shall be such as to establish the performance characteristics of the complete engine. Prior to the beginning of the calibration, the engine shall be cleaned using the wash procedure specified in 3.7.13, and all engine controls shall be adjusted and shall not be readjusted throughout the calibration. During calibration, engine inlet air shall be controlled to the temperature specified for the Table IA engine ratings. Calibrations shall be made initially with no customer power extraction and no bleed air extraction other than that required for continuous engine operation. Data indicated for calibration in 4.3.5.4 and 4.3.5.5 shall be recorded. During calibration, conformance with the leakage requirement of 3.3.6.4 and shutdown drainage requirements of 3.3.6.5 shall be demonstrated. The fuel and oil used shall be the same as those used during the test of 4.5.1.3. The following data shall be obtained.

a. Data required establishing compliance with sea level performance ratings in TableIA and IB and 3.2.1 of the engine specification and to establish the accuracy of the torque sensor signal.

b. Data required establishing compliance with 3.2.1.5.5, 3.2.1.5.6, and 3.7.9 at sealevel, static conditions.

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c. Repeat items "a" and "b" with maximum permissible bleed air extraction, which includes customer and anti-icing bleed air flow.

d. Repeat item "c" with accessory power extraction as specified by the Using Service.

4.5.1.2.2.1 Customer Bleed Air Analysis. Prior to and at the completion of the endurance test, a customer bleed air analysis shall be performed. Customer bleed air shall be sampled from each bleed air outlet during a maximum continuous power run. A sample of air entering the engine inlet shall be taken at the same time the bleed air samples are obtained. The samples shall be properly identified and processed through laboratory analysis to determine whether the contaminant levels are within the limits specified in 3.1.2.11.3.

4.5.1.3 Endurance Test Procedure. An engine conforming to the EIT/PFR configuration shall be tested for 150 hours of the following durability test cycles. Following the calibration run, the engine shall be adjusted to permit operation at the maximum allowable steady state gas temperature at the first stage turbine rotor inlet as defined in 3.2.1.4.5 and Table VI. The engine inlet air shall be controlled where necessary during a test cycle to ensure engine operation at the specified test conditions. For the purpose of this endurance test, operation at maximum continuous through contingency test temperature settings shall be defined as operation at a gas temperature at or above the maximum allowable (min endurance temperature) gas generator first stage turbine rotor inlet temperature specified in Table VI for the maximum continuous through contingency power settings. The engine shall operate at the MGT that corresponds to the first stage turbine rotor inlet temperature (TRIT). The relationship between MGT and TRIT shall be established via a 4-point performance calibration at the following ratings, prior to the test and checked at the conclusion of every 10th cycle (±2 cycles) and prior to starting the contingency power run. Rating Contingency Maximum Intermediate Max Continuous A performance calibration shall be conducted at 15°C (59°F) to determine sea level, standard day performance if either the following conditions exist when segment “a” maximum power data points are compared with segment “i” maximum power data points.

SHP at MGT decreases greater than or equal to 1.0 percent. Ng at MGT changes greater than or equal to 0.5 percent

For the purpose of the endurance test, minimum output shaft speed shall be interpreted as that output shaft speed at which the output shaft torque is not less than the torque limit for the applicable power condition or the minimum output shaft governed speed, whichever occurs first. Maximum output shaft speed shall be interpreted as the output shaft maximum speed limit or the output shaft maximum governed speed, whichever occurs first.

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For this engine test, the test cell shall conform to the power absorption device requirements of 4.3.3.3.2. The engine shall be subjected to an endurance test consisting of twenty five cycles of six hours each IAW the schedule listed below using MIL-PRF-23699 oil and MIL-DTL-83133H, grade JP-8 fuel. The inlet fuel pressure and temperature conditions shall be within the operating limits specified in this AQP and shall be maintained throughout the endurance test. Each cycle shall be preceded by a 2 hour shutdown. The test runs in each cycle shall be conducted in the order given. The time for changing power shall be charged to the duration of the lower setting. The 2nd, 7th, 20th and 23rd cycles shall be accomplished with the anti-icing bleed air system in operation. The fuel manifold/combustible fluid drainage provisions specified in 3.3.6.5 shall be demonstrated after a false start and also after a shutdown from normal operation. The oil pressure, if adjustable, shall be adjusted at the beginning of the test to the minimum steady state value specified in 3.2.1.4.8. No further adjustment shall be permitted during the test except when authorized by the Using Service. Oil consumption shall be determined for the 25 endurance cycles. Oil shall be added on an as needed basis as indicated by the oil tank sight glass or the low oil sensor. The engine shall be run with the control system in a control failure mode during runs “a” and “e” of every 5th cycle, for the times designated, to verify engine control system performance and reliability as specified in 3.7.2.2 and 3.7.2.2.3. If the engine has a FADEC with multiple channels of equal capability, the engine shall be operated using the 1st channel (A) during cycles 1–13. During these cycles, the control failure mode will be the 2nd channel (B). The engine shall be operated using channel B or additional modes during cycles 14-25. During these cycles, the control failure mode will be channel A. Accessory pads shall be subjected to rated loads and overhung moments as specified in Table V. The actual torque loading and overhung moments imposed during the endurance test shall be stated in the test report. The angular misalignment of the power absorber drive shaft to the engine output shaft shall not be less than the maximum allowable angular misalignment specified in 3.1.2.14.1. The engine internal washing provisions shall be demonstrated at the end of each cycle. The procedure for demonstration of the internal washing provisions shall be specified in the engine specification.At the completion of the endurance runs, the engine oil drain provisions specified in 3.7.7.4.2 shall be demonstrated. This demonstration shall also verify the adequacy and the locations for the oil debris monitor(s). At the end of each cycle, special engine control system features (e.g., overspeed, overtemperature, torque matching) shall be cycled throughout their functions. Each cycle shall consist of the following runs:

a. Maximum-Idle Run. This run shall consist of six successive periods of 10 minutes each. Each period shall include 5 minutes at maximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA followed by 5-minutes at idle operating condition. If the engine provides for anti-icing, at the end of each period at maximum test temperature setting, anti-icing controls shall be operated for one minute with the maximum anti-icing bleed air, before the power setting is changed. During the 5th, 10th, 15th, 20th and 25th, the first 3 minutes of each 5 minute period at maximum test temperature setting shall be run with the control in the failure mode. Transient data recording systems are to be on when switching control modes.

b. Incremental Torque Run. This run shall consist of 96 minutes including:

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1. Eight minutes at the maximum output shaft speed and the output shaft torque available at maximum test temperature setting or at the maximum continuous torque limit, whichever is less.

2. Eight minutes at the minimum output shaft speed and the maximum testtemperature setting or at the maximum continuous output shaft torque limit, whichever is less.

3. Eight minutes at the minimum output shaft speed and 90 percent of the output shaft torque obtained in (2).

4. Eight minutes at the maximum output shaft speed and 90 percent of the output shaft torque obtained in (1).

5. Eight minutes at the maximum output shaft speed and 80 percent of the output shaft torque obtained in (1).

6. Eight minutes at the minimum output shaft speed and 80 percent of the output shaft torque obtained in (2).

7. Four minutes at the minimum output shaft speed and 60 percent of the output shaft torque obtained in (2).

8. Four minutes at the maximum output shaft speed and 60 percent of the output shaft torque obtained in (1).

9. Four minutes at the maximum output shaft speed and 40 percent of the output shaft torque obtained in (1).

10. Four minutes at the minimum output shaft speed and 40 percent of the output shaft torque obtained in (2).

11. Four minutes at the minimum output shaft speed and 20 percent of the output shaft torque obtained in (2).

12. Four minutes at the maximum output shaft speed and 20 percent of the output shaft torque obtained in (1).

13. Four minutes at the maximum output shaft speed and 10 percent of the output shaft torque obtained in (1).

14. Four minutes at the minimum output shaft speed and 10 percent of the output shaft torque obtained in (2).

15. Four minutes at the minimum output shaft speed and at zero output shafttorque.

16. Four minutes at the maximum output shaft speed and at zero output shafttorque.

17. Four minutes at idle.18. Four minutes at idle and at zero output shaft speed.

If test data indicates the existence of a critical compressor or turbine vibration condition within the operating speed range between idle and maximum test temperature setting, at the option of the Using Service, the following shall be substituted for 48 minutes of the incremental torque run of each cycle to be chosen by the Using Service.

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1. Sixteen minutes at gas generator rotor and output shaft speeds which produce the critical vibration conditions.

2. Eight minutes at gas generator rotor and output shaft speeds 2.0 percent more than those which produce the critical vibration conditions.

3. Eight minutes at gas generator rotor speed 2.0 percent more and an outputshaft speed 2.0 percent less than those which produce the critical vibration conditions.

4. Eight minutes at gas generator rotor speed and output shaft speeds 2.0 percent less than those which produce the critical vibration conditions.

5. Eight minutes at a gas generator rotor speed 2.0 percent less and anoutput shaft speed 2.0 percent more than those which produce the critical vibration conditions.

c. Power Transient Run. This run shall consist of 39 minutes of power transients:1. Four minutes at no load condition followed by one minute at maximum test

temperature setting at the output shaft speed specified for the maximum rating in Table IA. Repeat the above for a total of 20 minutes.

2. The remaining 19 minutes of the run shall consist of one minute atmaximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, followed by an immediate decrease to no load condition. As soon as the engine reaches the no load condition, the engine power shall be increased to the temperature and output shaft speed associated with the previous condition and maintained at this power level for a period of one minute before repeating the cycle.

d. Incremental Power Run. This run shall consist of 9 minutes of operation in thesequence of conditions and time duration as follows: One minutes idle, three minutes intermediate test temperature, one minute at maximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, three minutes intermediate test temperature, and one minute idle.

e. Intermediate Run. This run shall consist of 30 minutes:1. Fifteen minutes at intermediate test temperature at the maximum

governed shaft speed. Output torque shall be at least as high as that specified in Table IA for the intermediate rating.

2. Fifteen minutes at intermediate test temperature at the minimum outputshaft speed. Output torque shall be at least as high as that specified in Table IA for the intermediate rating. During the 5th, 10th, 15th, 20th and 25th cycles, the first 26 minutes of this run shall be with the control in the failure mode. Transient data recording systems are to be on when switching the control from the automatic to the failure mode and also when switching back to the automatic mode from the failure mode.

f. Maximum Continuous Run. This run shall consist of 20 minutes:1. Ten minutes maximum continuous test temperature setting at maximum

governed shaft speed. 2. Ten minutes maximum continuous test temperature at the minimum output

shaft speed. Output torque shall be at least as high as that specified in Table IA for the maximum continuous rating.

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g. Intermediate Run. This run shall consist of 15 minutes at intermediate test temperature setting at intermediate rated output shaft speed.

h. Maximum Continuous Run. This run shall consist of 15 minutes at maximum continuous test temperature setting at maximum continuous rated output shaft speed.

i. Intermediate-Maximum Run. This run shall consist of six periods of 5 minutes each alternating between:

1. Five minutes at intermediate test temperature setting at intermediate rated output shaft speed.

2. Five minutes at maximum test temperature setting at maximum rated output shaft speed.

j. Maximum Continuous Run. This run shall consist of 16 minutes at maximum continuous test temperature setting at maximum continuous rated output shaft speed.

k. Intermediate-Maximum Run. This run shall consist of 30 minutes:1. Fifteen minutes at intermediate test temperature setting at intermediate

rated governed shaft speed. 2. Ten minutes at maximum test temperature setting at maximum rated

output shaft speed followed by five minutes at intermediate test temperature setting at intermediate rated output shaft speed. At 5 minute intervals during the run, the anti-icing controls shall be operated for one minute with maximum anti-icing bleed air.

4.5.1.3.1 Starts. A minimum of 300 starts, commencing with the first test cycle of 4.5.1.3, shall be made on the endurance test engine. In addition to the 300 endurance test starts, there shall be ten false starts (a starting sequence without benefit of light-off followed immediately after the permissible engine draining time by a successful start), and ten restarts (a start within a maximum of 14 minutes time from shutdown). Starts shall be performed with a starter that performs within 5% of minimum starter torque curve. The engine shall be started and shutdown no less than six times each test cycle. Of the at least 300 endurance starts, 110 shall be accomplished following varied regulated shutdown periods. Those starts at the beginning of each endurance cycle shall follow a shutdown period of at least two hours. The shutdown period for 18 starts shall be regulated to provide intervals between starts of 5.0 minutes, 10 minutes, and increasing, thereafter, by 5.0 minute increments up to and including 90 minutes for the 18th start. Each of the 18 regulated shutdown periods shall be preceded by immediate engine shutdown without being held at idle, after engine operation for a duration of no less than two minutes at intermediate test temperature setting. For 22 starts, the shutdown period shall be regulated to provide an interval between starts of no less than 45 minutes. The shutdown period for the remaining at least 50 starts need not be controlled. During all endurance starts, immediately after the engine has reached stabilized idle speed, an acceleration to the next scheduled endurance test condition shall be accomplished by a control power signal change from idle to the appropriate test temperature setting.

4.5.1.3.2 Contingency Power Rating. If the engine has a contingency power capability, this capability shall be flight rated by four periods of operation at contingency power test temperature. Each operation at contingency power shall be preceded by and followed by a period of operation at maximum continuous power. The duration of each period of operation shall be 3t0.677 where "t" is the rating time for contingency power. All times are in minutes. This will be demonstrated on the endurance test engine after completion of the endurance test.

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4.5.1.4 Recalibrations.

4.5.1.4.1 Engine Recalibration. After completion of the tests specified in 4.5.1.3 through 4.5.1.3.2, a recalibration in accordance with 4.5.1.2.2 shall be made on the endurance test engine. The recalibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature conditions, the values of shaft power obtained during the initial calibration. The recalibration may be preceded by a specified run during which the cleaning procedure of 3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration.

4.5.1.4.2 Component Recalibration. Functional bench recalibration/checks shall be conducted in accordance with 4.5.1.2.1.

4.5.1.5 Engine Disassembly and Inspection. The engine completing the endurance test shall be completely disassembled for examination of all parts. Prior to cleaning, the engine parts shall be given a “dirty inspection” for evidence of leakage, oil coking, unusual heat patterns, and abnormal conditions. The “dirty inspection” shall be completed before any parts are cleaned. The engine parts shall then be cleaned and a “clean inspection” shall then be performed. Engine part measurements shall be taken as necessary to determine excessive wear and distortion. These measurements shall be compared with the engine manufacturer’s drawing dimensions and tolerances and with similar measurements made prior to the test. Inspection techniques may also include but not be limited to: magnetic particle, fluorescent penetrant, eddy current, X-ray, and ultrasonic. During the “clean inspection” a visual examination and condition assessment shall be conducted. Upon completion of the clean inspection, the Using Service shall be provided all results of nondestructive tests and recommendations for modification or redesign of deficient parts. The Using Service shall be notified of the inspection commencement date prior to each inspection. The following data shall be made available to the Using Service during both inspections:

a. Inspection forms filled out by the contractor listing all observed deficiencies andjudgment as to the part’s return to service capability.

b. Tabulation of all parts found deficient.c. Detailed configuration list of the component or system tested.d. Test logs and list of test events.e. Spectrometric oil analysis report.f. Other measurements and data as deemed necessary by the Using Service.

4.5.1.6 Endurance Test Completion. The endurance test will be considered to be satisfactorily completed when the engine has completed the endurance test of 4.5.1 and during the final engine recalibration, the steady state first stage turbine rotor inlet gas temperature does not exceed a value of the gas temperature obtained for the initial calibration plus 30 percent of the difference between the maximum allowable (minimum endurance) steady state first stage turbine rotor inlet gas temperature and the rated temperatures specified in Table IV for all ratings; the corrected specific fuel consumption does not exceed 105 percent of the initial calibration values; the engine meets all other specified performance requirements which can be checked by the calibration procedure; and, the test engine and components are meeting the requirements of this plan at the end of the tests, recalibrations do not reveal excessive performance deterioration, and teardown inspections do

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not disclose parts failure or indicate impending failures. Parts will not be judged to have passed the endurance test until they have successfully completed the entire 150 hours of the endurance test. First stage turbine rotor inlet gas temperature deterioration and SFC increase shall be determined at the pretest rated output shaft power values.

4.5.2 Engine Component Tests. The following tests shall be conducted on components conforming to the same parts list and configuration used on the EIT endurance test. Controls and external component specification compliance shall be verified by test and inspection.

4.5.2.1 Previous Component Approval. Engine components requiring testing as specified herein may have these tests waived at the option of the Using Service, if the component has been previously approved by the Using Service for use on another engine. All such components must conform to the same parts list and configuration as the components previously approved.

4.5.2.2 Component Acceptance Test. Engine components shall be acceptance tested IAW procedures developed by the contractor and approved by the Using Service. All component acceptance test procedures (ATP) results shall be recorded and provided to the Government in test reports to substantiate component compliance with performance specifications. The contractor shall provide a list of components that do not require an ATP. This list shall be subject to approval by the Using Service. Components not requiring ATP shall be operated under normal operating conditions to demonstrate satisfactory functioning and compatibility with other system components.

4.5.2.3 Component Re-Test, Disassembly, and Inspection. Upon completion of specific tests, component ATP shall be repeated utilizing the same configuration-controlled test procedure as was used in 4.5.2.2. Components not subjected to an ATP shall be operated under normal operating conditions to demonstrate satisfactory functioning and compatibility with other system components. All component ATP results shall be recorded and provided to the Government in test reports to substantiate component compliance with performance specifications. All components shall then be completely disassembled and inspected for indications of failure, impending failure or excessive wear.

4.5.2.4 Component Test Success Criteria. The component tests shall be considered to be completed when:

a. During the tests, component performance and function were within establishedlimits.

b. During the tests, there was no fluid leakage from any component other than thatof a nature and rate specified in the engine specification.

c. During the tests, there was no hang-up or hesitation of any component.d. Recalibrations indicate that no component has changed its calibration beyond

allowable service limits. The component teardown inspection shows no indication of failed, excessively worn, and distorted parts. Measurements shall be taken and compared with the contractor’s drawing dimensions and tolerances or with similar measurements made prior to the test.

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4.5.2.5 Component Test Procedures.

4.5.2.5.1 Explosive Atmosphere. All electrical components not hermetically sealed, or components with moving parts capable of creating a spark, shall be subjected to explosive atmosphere testing IAW MIL-STD-810G, Method 511.5, Procedure I. Components shall be subject to pre-test ATPs or functional checks as applicable, but need not pass any post-test checks, unless specified by the Using Service. The test shall be conducted as specified at test altitudes of sea level, 3000 m, 6000 m, and 9000 m. During the test, ambient air surrounding the component shall be maintained at the maximum allowable limiting temperature. During the test, components shall have maximum input voltage applied and shall be operated continuously at their maximum loads. During each test altitude sweep, all make and break contacts shall be operated at least ten times. If a component utilizes airframe power, power supply transients shall be applied to the component during each test altitude sweep. Ten of each of these power supply transients shall be applied during each test altitude sweep, at least four of which shall be applied during operation of make and break contacts, if any.

a. The application of spike voltages shall consist of:1. Five spikes of +600 volts for 10 microseconds each.2. Five spikes of -600 volts for 10 microseconds each.3. Five spikes of +200 volts for 50 microseconds each.4. Five spikes of -170 volts for 50 microseconds each.

b. Power supply transients for alternating current systems shall consist of:1. 180 volts RMS for at least 0.10 seconds.2. 160 volts RMS for at least 0.50 seconds.3. 140 volts RMS for at least 2.0 seconds.4. 125 volts RMS for at least 6.0 seconds.

c. Power supply transients for direct current systems shall consist of:1. 80 volts for at least 0.050 seconds.2. 60 volts for at least 0.50 seconds.3. 40 volts for at least 2.0 seconds.4. 35 volts for at least 3.0 seconds.

Ignition components or systems shall be operated continuously. Electrodes of spark ignitors shall be mounted in such a manner that the explosive vapor in the test chamber shall not be contacted. Additionally, all components shall be shown by analysis or test that no external or internal constituent part, which can plausibly be surrounded by a flammable mixture, can exceed the minimum auto-ignition temperature of JP-4 (+230°C/+446°F) under worst case conditions. If the minimum delta margin is less than +20°C (+36°F) to the above auto ignition temperature,, the contractor shall submit the analysis or test results to the AED for approval of the analysis/test assumptions and methodology.

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4.5.2.5.2 Vibration – Airframe and Engine. Components shall be subjected to a vibration test in accordance with MIL-STD-810C, Method 514.2, Procedure I, Categories b.1 or c, to:

a. Verify functional performance of the component when exposed to traditional sinusoidal resonance searches and dwells.

b. Determine that there are no unacceptable component resonance modes.Components may be tested in test assemblies or as individual units; however, components shall be tested to simulate installation orientation or other environmental or physical details. During the test, the component shall be subjected to its maximum limiting temperature (as specified in the engine model specification) and shall be operating to evaluate any functional effects of the vibration on specified performance. During post-EIT flight testing, the contractor shall make every effort to ascertain the actual installed vibration environment of each component in order to perform qualification testing as specified in 4.6.2.3.4.2a.

a. Unknown Application Environment. For engine-mounted components, ProcedureI, Part 1, Curve L, and Part 2, Curve AR (if normally equipped with vibration isolators) of Category b.1 shall apply. For airframe-mounted components, Procedure I, Part 1, Curve M, and Part 2,Curve B (if normally equipped with vibration isolators) of Category C shall apply.

b. Known Application Environment. If the vibration environment is known, the testmaybe performed in accordance with the functional performance qualification requirements of 4.6.2.3.4.2.a.1, with the test time per axis reduced from 4 hours to 2 hours. However, no qualification credit will be given unless the complete test time of 4 hours per axis is accomplished. The service life test, defined in subparagraph (2) of 4.6.2.3.4.2.a, is not required for EIT/PFR.

4.5.2.5.3 Fuel Pump Altitude Test. The portion of the fuel system from the engine fuel inlet to the engine fuel pump inlet, shall be simulated in the test assembly. This shall include lines, fittings, filters, and other items as applicable between the engine fuel inlet and fuel pump inlet as well as any elements of the fuel system downstream of the pump which might have an effect on the pump. All independent and separately replaceable fuel pumps shall be operated for:

a. One hour at the flow, pressure, and speed corresponding to those required bydesign maximum engine performance at sea level, 35°C (95°F) conditions with a fuel temperature of at least 57°C (135°F) and a fuel inlet pressure not more than the minimum specified in 3.7.3.3.4b, followed by one hour at the flow, pressure, and speed corresponding to those required by design idle engine performance and a fuel inlet pressure as established by the line restriction determined above.

b. One hour at the flow, pressure, and speed corresponding to those required bydesign maximum engine performance at 10000 ft, -5°C (23)°F conditions with a fuel inlet temperature of at least 57°C (135°F) and a fuel inlet pressure not more than the minimum specified in 3.7.3.3.4b, followed by one hour at the flow, pressure, and speed corresponding to those required by design idle engine performance and fuel inlet pressure as established by the line restriction determined above.

c. One hour at flow, pressure, and speed corresponding to those required by designmaximum engine performance at 20000 ft, -25°C (-13°F) conditions with a fuel inlet temperature of at least 43°C (109°F) and fuel inlet pressure not more than the minimum specified in 3.7.3.3.4b, followed by one hour at the flow, pressure, and speed corresponding to those required by design

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idle engine performance at and fuel inlet pressure as established by the line restriction determined above. The fuel shall be IAW MIL-DTL-83133H, grade JP-8. Except as stated above, the test procedure shall be IAW ARP 492C. During this test the fuel pump dry lift capability defined in 3.7.3.3.4.1 shall also be verified. The fuel pump altitude test shall be considered to be satisfactorily completed when, in the judgment of the Using Service, the pump performance during the course of testing and the performance deterioration, determined from ATP and/or calibration runs, conform to the requirements established by the applicable design specification and not adversely affect engine performance; the pump dry lift capability as defined in 3.7.3.3.4.1 has been verified; and the component teardown inspection does not disclose parts failure or impending failures.

4.5.2.5.4 Oil Reservoir Pressure Test. The tank with filler cap installed shall be subjected to positive and negative differential pressures in accordance with 3.7.7.4.1 for a period of 30 minutes. No leakage or detrimental deformation shall occur.

4.5.2.5.5 Fire Test. Lines, fittings, and components, including engine furnished oil tanks, which convey flammable fluids shall be tested to verify conformance with 3.3.6.1. Individual lines, fittings, components, or assemblies shall be tested as specified in AC20-135, AS 1055D and AS 4273A while conveying fluids at the lowest flow rate, highest pressure, and highest fluid temperature possible over the complete engine operating envelope. The requirements of 3.3.6.1 shall be considered verified, if at the completion of the test period, and for a period of 5 minutes after removal of the flame, there are no measurable leaks (this requirement applies only to components that are tested to a 5-minute fire resistant rating). For fire proof rated components, the post-flame leakage requirement is not applicable. All fire tests shall be conducted using JP-8 for fuel components, or a suitable substitute as approved by the Using Service. Fuel shutoff capability following the exposure of the fuel shutoff component to the flame shall be demonstrated. It is also acceptable for any shutoff to function automatically at any time during the period when flame is applied or at any time during the 5-minute post flame period.If the engine control system is located in a designated fire zone, the engine control system shall be tested to verify conformance to 3.3.6.7. During the five-minute flame application period, and for five minutes thereafter, the engine control system shall continue to control the engine in accordance with the requirements of this AQP, or shall cause the engine to fail safe.

4.5.2.5.6 Overheat Test. Overheat testing shall be conducted on all electronic engine control assemblies, unless the component is mounted in a location where the ambient environment surrounding the component cannot exceed the component limiting temperature as specified in the engine specification, even under failure conditions. If the engine control electronic assembly incorporates hardware or software to initiate hardware shutdown/shutoff, or other defined accommodation, in response to excessive internal temperatures, then this test shall be performed even if the external ambient temperature cannot exceed the defined component limiting temperature under failure conditions. The engine control system shall be tested to verify conformance to the fail-safe requirements of 3.3.6.7. The engine control system shall be subject to pre- and post-test ATPs or functional checks as applicable, and shall be operating in a ‘closed-loop’ mode wherein the engine/airframe is simulated during the test in order to assess the effects on the engine of the engine control system overtemperature condition. The engine control system need not pass a post-test ATP or functional check.

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Procedural conduct of this test is not defined and shall be agreed between the contractor and the Using Service in the pretest data. Note that this test is in addition to any fire test, since the failure mechanisms for the two tests are different; one being temperature elevation of the entire component while the other is a potential rapid burn-through of the housing and flame impingement on the circuit card assemblies.

4.5.2.5.7 Output Shaft Static Torque Test. The output shaft system shall be subjected to testing at the maximum allowable output torque. The shaft shall be heated to simulate its operating temperature within the engine at maximum test temperature. Torque of not less than the value specified in 3.2.1.4.11 shall be applied for 10 continuous cycles, each cycle consisting of a 15 minute period and a 10 minute period at the maximum allowable output torque (4 hours 10 minutes total). The torque shall be reduced to zero before each period. At completion of the test, the shaft shall be within allowable dimensional limits, torque accuracy shall still be within engine specification limits and there shall be no evidence of impending failure.

4.5.2.5.8 Impact (Shock) Test. Components located in a crew compartment shall be subjected to an crash hazard shock test in accordance with MIL-STD-810G, Method 516.6, Procedure V. The shock test spectrum shall be IAW Figure 516.6-8 (SRS), or Figure 516.6-10 (terminal peak sawtooth) as approved by the Using Service. Tests shall be conducted under room ambient conditions. The component need not be operating during the test nor pass a post-test ATP or functional check. With the approval of the Using Service, a dummy unit may be used if it duplicates the mounting arrangement, overall mass, and moments of inertia of the actual component.

4.5.2.5.9 Software/Programmable Hardware Verification In order to receive a Preliminary Flight Rating (PFR) for flight test, the control system shall demonstrate it meets the requirements of 3.7.2 and all subparagraphs by passing a formal qualification (FQT) in accordance with Appendix G of this AQP. Tests shall be conducted to verify and validate both engine control and EMS software. The software used for PFR shall be base-lined and under configuration control.

4.5.2.5.10 Electrical Power System Test Electrical power system tests shall be conducted on an engine to substantiate compliance with model specification requirements for engine safety and fail-fixed performance (if required).

4.5.2.5.10.1 Engine-Supplied Electrical Power During the time that electrical power is OFF, and during any switching events (both ON to OFF and OFF to ON), the engine shall not exceed any of its limits nor exhibit any discernible performance or operability changes. The engine shall complete all transients satisfactorily. There shall be no loss of control system functionality, of any kind, throughout the test.

a. Enable all engine-supplied and airframe-supplied power sources. Interrupt the airframesupplied electrical power source for 500 msec, first on each control channel separately(if more than one channel), then on all control channels simultaneously, at the followingengine operating conditions

1. Engine steady-state at ground idle2. Maximum engine acceleration

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3. Engine steady-state at maximum continuous power4. Maximum engine deceleration

b. Enable all engine-supplied and airframe-supplied power sources.

1. Start the engine and stabilize at ground idle for 1-2 minutes.2. Begin to accelerate the engine towards maximum continuous power and then

disconnect, fail or switch off engine-supplied electrical power to the control system atthe most critical point of the acceleration, as defined by the contractor and approvedby the Using Service.

3. Re-enable engine-supplied electrical power and operate the engine at maximumcontinuous power for one minute, then disconnect, fail or switch off the engine-supplied electrical power to the control system.

4. Re-enable engine-supplied electrical power and operate the engine at maximumcontinuous power for one minute. Disconnect, fail or switch off the airframe-suppliedexternal power to the control system.

5. Re-enable airframe-supplied electrical power and operate the engine at maximumcontinuous power for one minute. Begin to decelerate the engine towards ground idleand then disconnect, fail or switch off engine-supplied electrical power to the controlsystem at the most critical point of the deceleration, as defined by the contractor andapproved by the Using Service.

6. Perform a normal engine shutdown.

c. After the engine has been shut down, and with the engine-supplied electrical power stilldisconnected, restart the engine and accelerate to ground idle. After 1 minute,accelerate to maximum continuous power and operate there for 1 minute, thendecelerate to ground idle. Operate at ground idle for 1 minute and then perform a normalengine shutdown.

4.5.2.5.10.2 Aircraft-Supplied Electrical Power During all tests, the engine supplied electrical power shall be disconnected and, except as noted, the engine shall be at maximum continuous power. For the interrupt testing of subparagraphs a through c, the following performance requirements shall be met:

1. During the single-channel interrupt testing, there shall be no changes in engineperformance or operability, nor loss of control system functionality, of any kind,throughout the test, when the interrupted channel is the standby control. Interruptsof the controlling channel shall not result in unsatisfactory engine transients orcontrol system behavior (as determined by the Using Service) during the interruptand recovery period. However, switchover to the standby channel is allowable.

2. Simultaneous channel interrupt testing shall not result in unsatisfactory enginetransients or control system behavior (as determined by the Using Service) duringthe interrupt and recovery period. Engine fail-fix operation or engine shutdown isallowable, dependent upon control system architecture.

3. The interrupted channel shall be fully functional within 500 msec after powerinterrupt (after power is restored), unless otherwise specified.

a. Bus Transfer Interrupt. Interrupt the airframe-supplied electrical power source for 50msec, first on each control channel separately (if more than one), then on all controlchannels simultaneously, at the following engine operating conditions.

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1. After 1 minute at steady-state ground idle.2. During maximum engine acceleration.3. After 1 minute at steady-state maximum continuous power.4. During maximum engine deceleration.

b. Bus Transfer Interrupts. Repeat a(3) for each control channel (if more than one) and allchannels simultaneously for interrupt durations of 2 msec, 4 msec, 6 msec, 8 msec, 10msec, 12 msec, 14 msec, 16 msec, 20 msec, and 30 msec.

c. Long-Duration Interrupts. Repeat a(3) for each control channel (if more than one) andall channels simultaneously for interrupt durations of 200 msec, 500 msec, 1 sec, 2 sec, 5sec and 7 sec.

For the testing specified in subparagraphs d through g, there shall be no changes in engine performance or operability, nor loss of control system functionality, of any kind, throughout the test.

d. Steady-State Voltage. Simultaneously on all airframe-supplied electrical power sources:16.0 Vdc for 30 minutes minimum (or 12.0 Vdc if so specified in the model specification)and then 31.5 Vdc for 30 minutes minimum.

e. Transient Voltage. Simultaneously on all airframe-supplied electrical power sources:

1. Start at 28 Vdc, pulse to 80 Vdc minimum for 50 msec minimum, end at 28 Vdc.2. Start at 28 Vdc, pulse to 60 Vdc minimum for 500 msec minimum, end at 28 Vdc.3. IAW MIL-HDBK-704-8, LDC105 Normal Voltage Transients for MIL-STD-704A4. IAW MIL-HDBK-704-8, LDC302 Normal Voltage Transients for MIL-STD-704A

f. Ripple Voltage. Simultaneously on all airframe-supplied electrical power sources IAWMIL-HDBK-704-8, LDC 103 and LDC 104 for MIL-STD-704A.

4.5.2.5.11 Helicopter Drive System Torsional Stability In order to receive an AWR for flight, a torsional stability analysis of the helicopter drive system shall be submitted in accordance with the technical data requirements of paragraph 4.2 of ADS 9C when any newly developed component, or change to an existing component, is determined to affect the behavior of any of the control laws which govern the stability of the engine/rotor system. General guidelines for preparation of the analysis are described in ARP704.

4.5.3 Altitude Tests. An engine, conforming to the same parts list and configuration as the endurance test engine, shall be subjected to altitude tests which shall consist of operation and air starting checks at several selected conditions within the engine operating limits envelope specified for the engine and at least those given in this AQP per Figure 19. Pretest data shall be per 4.3.5.1. The test points shall include the effects of power extraction, bleed air extraction, inlet recovery and anti-ice operation on engine performance and stability. The effects of anti-ice operation shall be included in the demonstration of 4.5.3.2 sections a. b. and c. Data to be taken and recorded during the test shall be as specified in Table XII. Adjustments to the engine shall not be made without approval of the Government representative. The altitude tests shall be accomplished

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using MIL-PRF-23699 oil and MIL-DTL-83133H, grade JP-8 fuel. Fuel temperature shall be varied over a range sufficient to encompass all anticipated engine operating environments. For this engine test, the test cell shall conform to the power absorption device requirements of 4.3.3.3.2. Overall true RMS velocity measurements and acceleration spectrograms shall be obtained for each sensor mounted on the engine case and accessory gearbox case at the engine speeds and powers selected for the test. The points selected shall include at least the altitude rating points and the point in the operating envelope where the highest engine vibration levels are generated. Critical components of the engine shall be identified on each spectrogram.

4.5.3.1 Altitude Calibration. Prior to the initiation of the testing described in 4.5.3.2, the engine shall be calibrated IAW 4.5.1.2.2. No control readjustments shall be made after the initial adjustments at the beginning of the calibration. In addition, calibration shall be conducted on the control system components IAW 4.5.1.2.1.

4.5.3.2 Altitude Test Procedure. Operation at each test point shall be of sufficient duration to stabilize the engine and to establish the performance and operating characteristics of the engine. Engine operation with the control system in control failure modes shall be evaluated and the effects on engine performance and operating characteristics shall be determined during the test. The control failure modes to be evaluated shall be specified in the pretest data and shall be subject to approval by the Using Service. Operation shall be conducted to obtain the following data:

a. Altitude Rating Points. The test conditions shall be those specified for altitude ratings in Table IIA, IIB, IIC and IID of the engine specification. A sufficient number of additional engine power settings shall be selected for each specified altitude test point to establish operating and performance characteristics at rated conditions. The time elapsed versus engine speed, measured temperature, and fuel flow shall be obtained for stability verification with the power settings of idle, maximum continuous, intermediate, and maximum. The time period for stability verification shall be a minimum of five minutes at each power setting.

b. Transient Operation. The applicable transient performance specified in 3.2.1.5.6 shall be demonstrated at each rating condition.

c. Functional Test. The operating envelope of the engine shall be verified by operation at the extremities of the engine operating envelope. Engine steady state and transient characteristics shall be determined at each test point over the range of power settings. Determination of the engine operating characteristics shall be accomplished up to an altitude of 20000 ft.

d. Starts and Restarts. Flameouts, with the ignition system not operating, shall be accomplished by means specified in 3.2.1.5.3. Engine air starts and restarts, shall be accomplished at each of the specified air starting points shown on Figure 19.

4.5.3.3 Altitude Test Completion. Comparison of observed data obtained during the test to the specified performance and operating characteristics shall be made by a method acceptable to the Using Service to determine compliance with the engine specification. The test shall be considered to be satisfactorily completed when, in the judgment of the Using Service:

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a. The engine performance is at least that specified in the steady state performance computer program and pressure and temperature lapse rates for MGT, SHP, SFC, and Ng are ±2.0 percent of those obtained from the steady state performance computer program.

b. The altitude starting and transients conducted during the test are in accordance with the engine specification requirements.

c. The functional test points demonstrate satisfactory engine operation and do not show any discrepancies with the steady state performance computer program, altitude starting data, or transient data.

4.5.4 Structural Tests. Engines or components conforming to the parts list and configuration of the EIT endurance test engine shall be used for the following tests.

4.5.4.1 Component Vibration Characterization. A vibration test shall be conducted on an engine carcass to verify that external component resonant frequencies are not in the normal engine and airframe induced frequency ranges. The effect of any observed resonance will be evaluated on a case by case basis subject to approval by the Using Service. The test shall be performed on an engine carcass with all external components. The engine carcass shall not be assembled with rotating hardware in the gas generator, power turbine, and accessory gearbox including shafting. External components including tubes, fuel control, valves, main oil tank, IPS blower, etc. shall be instrumented with accelerometers or strain gages as appropriate. The engine carcass will be suspended in a free-free mode and driven by a portable vibration exciter (stinger) from 5 Hz to 2000 Hz.

4.5.4.2 Rotor Structural Integrity.

4.5.4.2.1 Overspeed. The most critical stage of each rotor system including the power turbine, gas generator turbine, compressor, and inlet particle separator, as defined by analysis and agreed to by the Using Service, shall be subjected to engine or spin pit operation for a stabilized period of at least five minutes duration at component temperatures predicted to occur during operation at stabilized maximum allowable first stage turbine rotor inlet temperature (3.2.1.4.5) at one of the following speed conditions. If an ambient temperature spin pit is used, additional demonstration speed shall be added to accommodate material property scatter (average to -3σ yield strength values), yield strength differences resulting from the temperature difference between engine operating conditions and spin pit test conditions, and actual versus average yield strength material properties in accordance with the guidance of the Using Service. If a heated spin pit or engine test is used, additional demonstration speed shall be added to accommodate material property scatter (average to -3σ yield strength values) and actual versus average yield strength material properties in accordance with the guidance of the Using Service. Prior to any testing, critical disk dimensions (such as rotor bores and other critical pilots) agreed to by the Using Service shall be recorded for comparison to post test measurements. Gas generator turbine and compressor rotors shall be subjected to 115 percent of the maximum allowable steady state speed limit specified in 3.2.1.4.6. Power turbine rotors shall be subjected to the greater of the following two methods:

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a. 115 percent of the transient speed limit specified in 3.2.1.4.6.b. 105 percent of the predicted peak speed attained in a loss of load analysis per the

requirements stated in 3.7.2.3.3.1, conducted throughout the engine operating envelope at worst case ambient temperature and altitude conditions. For calculation of this speed, the analysis shall assume loss of load at MRP redline conditions (min endurance test temperature per Table VI), 100% PT design speed per 3.2.1.4.6 and worst stackup of overspeed control system tolerances from the loss of load event until the point at which the overspeed device actuates. Following these overspeed tests the power turbine, compressor and gas generator turbine parts and assemblies shall be measured at the same critical locations as pre-test to determine residual growths. Analysis shall present why/how these residual growths are acceptable or unacceptable for continued disk operation after an overspeed event. The parts and assemblies shall show no evidence of impending failure, in the judgment of the Using Service.

4.5.4.2.2 Overtemperature. Using the same engine rotors that the completed the overspeed test, the engine shall be operated at a first-stage turbine rotor inlet temperature of at 45°C (81°F) in excess of the maximum allowable first stage turbine rotor inlet temperature as specified in 3.2.1.4.5a, at no less than maximum allowable steady state speed specified in 3.2.1.4.6 for all rotors, for five minutes. Following the test, parts and assemblies shall comply with 3.3.8.8 and there shall be no evidence of impending failure.

4.5.4.3 Engine Static Load Test. The engine cases and mounts of the endurance engine configuration shall be subjected to a static test to verify the requirements of 3.1.2.5 and 3.1.2.6. A static rig test utilizing the applicable engine static structure shall be conducted to demonstrate the capability of the engine and its supports to meet the requirements of this AQP during and after exposure to maximum externally applied forces specified in Figure 1 without permanent deformation of any component and 1.5 times those forces without failure of any component. In this test, maximum thrust loads, acceleration loads, gyroscopic moments, torque, and reaction loads shall be applied separately and then in combination. Stress and deflection data shall be obtained at critical locations as determined by analysis and preliminary stress coating tests. The limit loads shall be based on a mass factor consisting of the dry mass of the engine, increased by the specific mass allowed for all engine mounted accessories and operating fluids.

4.5.4.4 Attitude Test. The engine shall be subjected to an attitude test to demonstrate compliance with 3.2.1.5.1 and 3.7.7.3. Engine capability to operate for the time specified in 3.2.1.5.1 for both negative “g” and zero “g” conditions shall be verified by analysis or by a rig test of the lubrication oil system. The engine shall be started, and then operated at maximum rated rotor speeds specified in Table I for at least 30 minutes at each of the test points shown in the clear area of Figure 9. The engine shall also be operated at maximum rated rotor speeds for at least 30 seconds at each of the test points shown in the shaded area of Figure 9. This test shall be considered satisfactorily completed when, in the judgment of the Using Service, the engine starts satisfactorily, remains within all operating limits, and there is no evidence of mechanical damage.

4.5.4.5 Loss of Oil Test. The engine shall be operated at 75% maximum continuous power for 6 minutes after loss of all oil supply to the engine bearing sumps. Upon completion of the above: the engine will be shut down, the oil tank filled, the engine re-started within 15 minutes and operated at 75% maximum

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continuous power for 30 minutes. Post test condition of the bearings may indicate damage but in the judgment of the Using Service the bearings shall not be at an imminent failure condition.

4.5.4.6 Engine Crash Load Test. The engine cases and mounts of the endurance engine configuration shall be subjected to a static test to verify the design. A static rig test utilizing a full engine carcass with the applicable mass of an engine with all engine mounted accessories and operating fluids shall be conducted to demonstrate the capability of the engine and its supports to withstand to peak crash loads without catastrophic failure, per paragraph 3.1.2.5.

4.5.4.7 Not used.

4.5.4.8 Low Cycle Fatigue Engine Test. A low cycle fatigue test of 1875 cycles shall be performed. No manual adjustment of the engine control system components is required. The accessory pads shall be loaded to provide maximum continuous loads. The engine components shall be calibrated prior to the test and recalibrated after the test in accordance with paragraph 4.5.1.2.1.

For this engine test, the test cell shall conform to the power absorption device requirements of 4.3.3.3.2.

The engine shall be calibrated prior to the test and recalibrated after the test. The calibration shall be conducted in accordance with paragraphs 4.5.1.2.2a, 4.5.1.2.2b, and 4.5.1.2.2.1 except compliance with the starting torque requirement (paragraph 3.7.9) need not be shown. The recalibration shall be in accordance with paragraphs 4.5.1.2.2a, 4.5.1.2.2b, and 4.5.1.2.2.1 except that compliance with the starting torque requirement need not be shown. The recalibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature condition, the values of output shaft power obtained during the initial calibration. The recalibration may be preceded by a specified run during which the cleaning procedure of paragraph 3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration. The test will be considered to be satisfactorily completed when the engine has completed the test and during recalibration, the steady state first stage turbine rotor inlet gas temperature does not exceed a value of the gas temperature obtained for the initial calibration plus 30 percent of the difference between the maximum allowable (minimum endurance) steady state first stage turbine rotor inlet gas temperature and the rated temperatures specified in Table IV for all ratings; the corrected specific fuel consumption does not exceed 105 percent of the calibration values; the engine meets all other specified performance requirements which can be checked by the calibration procedure; and, in the judgment of the Using Service, the test engine and components are operating satisfactorily at the end of the tests and teardown inspections do not disclose parts failure or impending failures. The final corrected specific fuel consumption shall be reported. First-stage turbine rotor inlet gas temperature deterioration and SFC increase shall be determined at the same rated output shaft power.

The LCF test shall be run at constant power at a power turbine speed of 105%. The establishment of the maximum power settings at the start of the test shall be as follows:

For operation at maximum power this shall be at the power corresponding to the pre-test calibration at a gas temperature at or above the rated gas generator first stage turbine rotor inlet temperature specified in Table VI for the maximum rated power setting at 105% power turbine speed.

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Following completion of the test, the engine shall be disassembled and inspected for evidence of cracking and shall be within allowable limits. The low cycle fatigue test cycle shall be as follows:

Approximate Approximate total time schedule time (minute) (minute) Event 0.5 0.5 Start engine 2.5 2.0 Run at idle 2.6 0.1 Accelerate to maximum power 10.6 8.0 Run at maximum power 10.7 0.1 Decelerate to idle 12.7 2.0 Run at idle 15.0 2.3 Shutdown and cool down

4.5.4.9 Main Shaft, Seals and Bearing Mechanical Tests. A test program to evaluate bearing, seal, shaft and gear operating characteristics (heat rejection, stresses, loads, life, etc.) and to demonstrate that sump temperatures during soak back do not exceed oil coking temperatures shall be conducted. A total of at least 1000 hours of endurance, AMT, and/or LCF testing shall be accumulated on each main shaft bearing design with at least 500 hours accumulated on a single bearing of each design. Critical bearings shall be instrumented to include but not necessarily be limited to inner and outer races temperatures, oil input and output temperatures, cage speeds, bearing loads versus engine power condition, and vertical, lateral, and axial vibrations. Demonstration methods including engine, gas generator, or rig tests may be used to acquire inner race temperature data.

4.5.5 Environmental Tests

4.5.5.1 Electromagnetic Environmental Effects (E3). The requirements of 3.2.8.4 shall be verified by analysis and test.

4.5.5.2 Electromagnetic Interference (EMI). The requirements of 3.2.8.4.1 shall be verified by analysis and testing IAW ADS-37A-PRF.

4.6 Qualification Test (QT) Rating.

4.6.1 Endurance Test. The endurance test shall consist of 300 hours operation on a single engine IAW the following test schedules, calibrations, and procedures. The engine shall be tested using MIL-PRF-7808 oil and MIL-DTL-83133H, grade JP-8 fuel.

4.6.1.1 Pretest Verification.

4.6.1.1.1 Engine Dry Weight. Prior to initiation of the calibration, the dry weight of the engine as specified in 3.2.2.1 shall be verified in accordance with 4.3.5.2.

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4.6.1.2 Calibrations (QT).

4.6.1.2.1 Component Calibration. Functional bench calibration/checks shall be conducted on each of the components specified in Table XVIII. All fuel nozzles and fuel carrying components of the engine control system shall undergo bench calibrations using fluid in accordance with 3.7.3.3.1. Components shall conform to the design tolerance range required by the applicable design specifications. All control system sensors shall be calibrated. Temperature sensing system performance shall meet the tolerance and thermal response characteristics specified in 3.7.6.4 of this AQP.

4.6.1.2.2 Engine Calibration. The procedure during the engine calibration shall be such as to establish the performance characteristics of the complete engine. Prior to the beginning of the calibration, the engine shall be cleaned using the wash procedure specified in 3.7.13, and all engine controls shall be adjusted and shall not be readjusted throughout the calibration. During calibration, engine inlet air shall be controlled to the temperature specified for Tables IA, IB, and IC engine ratings. Calibrations shall be made initially with no customer power extraction and no bleed air extraction other than that required for continuous engine operation. Data indicated for calibration in 4.3.5.4 and 4.3.5.5 shall be recorded. During calibration, conformance with the leakage requirement of 3.3.6.4 and shutdown drainage requirements of 3.3.6.5 shall be demonstrated. The fuel and oil used shall be the same as those used during the test of 4.6.1.3. The following data shall be obtained:

a. Data required establishing compliance with sea level performance ratings inTables IA and IB and 3.2.1 of this specification and to establish the accuracy of the torque sensor signal.

b. Data required establishing compliance with 3.2.1.5.5, 3.2.1.5.6, and 3.7.9 at sealevel, static conditions.

c. Repeat items "a" and "b" with maximum permissible bleed air extraction, whichincludes customer and anti-icing bleed air flow.

d. Repeat item "c" with accessory power extraction as specified by the UsingService.

4.6.1.2.3 Customer Bleed Air Analysis. Prior to initiation of the endurance test, customer bleed air analysis shall be performed. The customer bleed air shall be sampled from each bleed air outlet during a maximum continuous power run. A sample of air entering the engine inlet shall be taken at the same time the bleed air samples are obtained. The samples shall be properly identified and processed through laboratory analysis to determine whether the contaminant levels are within the limits specified in 3.1.2.11.3. The results of the analysis, the methods and test apparatus used shall be detailed in the engine test report.

4.6.1.3 Endurance Test Procedure. An engine conforming to the QT configuration shall be tested for 300 hours of the following durability test cycles. Following the calibration run, the engine shall be adjusted to permit operation at the maximum allowable steady state gas temperature at the first stage turbine rotor inlet as defined in 3.2.1.4.5. The engine inlet air shall be controlled where necessary during a test cycle to ensure engine operation at the specified test conditions. For the purpose of the endurance test, operation at

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maximum continuous through contingency test temperature settings shall be defined as operation at a gas temperature at or above the maximum allowable (min endurance temperature) gas generator first stage turbine rotor inlet temperature specified in Table VI for the maximum continuous through contingency power settings. The engine shall operate at the MGT that corresponds to the first stage turbine rotor inlet temperature (TRIT). The relationship between MGT and TRIT shall be established via a 4-point performance calibration at the following ratings, prior to the test and checked at the conclusion of every 10th cycle (±2 cycles). Rating Contingency Maximum Intermediate Max Continuous A performance calibration shall be conducted at 15°C (59°F) to determine sea level, standard day performance if either the following conditions exist when segment “a” maximum power data points are compared with segment “i” maximum power data points. • SHP at MGT decreases greater than or equal to 1.0 percent• Ng at MGT changes greater than or equal to 0.5 percentFor the purpose of the endurance test, minimum output shaft speed shall be interpreted as that output shaft speed at which the output shaft torque is not less than the torque limit for the applicable power condition or the minimum output shaft governed speed, whichever occurs first. Maximum output shaft speed shall be interpreted as the output shaft maximum speed limit or the output shaft maximum governed speed, whichever occurs first. For this engine test, the test cell shall conform to the power absorption device requirements of 4.3.3.3.2. The engine shall be subjected to a 300 hour endurance test consisting of 50 cycles of six hours each in accordance with the schedule listed below using the oil and fuel specified in 4.6.1. Each cycle shall be preceded by a 2 hour shutdown. The test runs in each cycle shall be conducted in the order given. The time for changing power shall be charged to the duration of the lower setting. If the engine does not have a maximum rating, a suitable test temperature setting shall be specified by the Using Service and shall be substituted for maximum test temperature setting throughout the test schedule. Every 5th cycle shall be accomplished with the anti-icing bleed air system in operation. The fuel manifold/combustible fluid drainage provisions specified in 3.3.6.5 shall be demonstrated after a false start and also after a shutdown from normal operation. Inlet fuel shall be maintained at minimum specified fuel pressure throughout the test. During five successive cycles prior to the tenth cycle and five successive cycles after the 40th cycle of the endurance test, the temperature of the fuel shall be maintained at the maximum temperature specified in 3.7.3.1.3.3. The inlet fuel temperature and pressure conditions shall be within the operating limits of 3.7.3.1.3.3. The 4th, 9th, 14th, 19th, 24th, 29th, 34th, 39th, 44th, and 49th cycles shall be accomplished with the anti-icing bleed air system in operation. For all operations during cycles 1–10, 21–30, 41–50 the oil temperature shall be maintained at no less than the maximum oil temperature specified in 3.2.1.4.8. The oil pressure, if adjustable, shall be adjusted at the beginning of the test to the minimum steady state value specified in 3.2.1.4.8. No further adjustments shall be permitted during the test except when authorized by the Using Service. Oil consumption shall be determined and reported after each block of 10

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cycles. Oil shall be added on an as needed basis as indicated by the oil tank sight glass or the low oil sensor. Oil drained for analysis shall not be charged to engine oil consumption and shall be replaced by an equivalent amount of new oil. The engine shall be run with the control system in the control failure mode during runs “a” and “e” of every 5th cycle, for the times designated, to verify engine control system performance and reliability as specified in 3.7.2.2 and 3.7.2.2.3. If the engine has a FADEC with dual channels of equal capability, the engine shall be operated using the 1st channel (A) during cycles 1–25. During these cycles, the control failure mode will be the 2nd channel (B). The engine shall be operated using channel B during cycles 26–50. During these cycles, the control failure mode will be channel A. Accessory pads shall be subjected to rated loads and overhung moments as specified in Table V. The actual torque loading and overhung moments imposed during the endurance test shall be stated in the test report. During the test, the exhaust duct, intake duct, and bleed air duct connections shall be loaded as specified in 3.7.10.1, 3.1.2.10.2, and 3.1.2.11.1 respectively. If the engine is supplied with an infrared suppression system, it shall operate continuously during runs “e” and “f”.If the engine provides special features such as fuel heaters, indicator lights, and switching functions, these items shall be actuated during selected test runs as specified in the detailed test procedures as approved by the Using Service. The angular misalignment of the power absorber drive shaft to the engine output shaft shall not be less than the maximum allowable angular misalignment specified in 3.1.2.14.1. The engine internal washing provisions shall be demonstrated once every 10 cycles. The procedure for demonstration of the internal washing provisions shall be as specified in the pretest data. At the completion of the endurance runs, the engine oil drain provisions specified in 3.7.7.4.2 shall be demonstrated. This demonstration shall also verify the adequacy and the locations for the oil debris monitor(s). At the end of each cycle, special engine control system features (e.g., overspeed, overtemperature, torque matching) shall be cycled throughout their functions. Each cycle shall consist of the following runs:

a. Maximum-Idle Run. This run shall consist of six successive periods of 10 minutes each. Each period shall include 5 minutes at maximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, followed by 5.0 minutes at idle operation condition. If the engine provides for anti-icing, at the end of each period at maximum test temperature setting, anti-icing controls shall be operated for one minute with the maximum anti-icing bleed air, before the power setting is changed. During the 5th, 10th, 15th, 20th, 25th, 30th, 35th, 40th, 45th, and 50th cycles, the first 3.0 minutes of each 5.0 minute period at maximum test temperature shall be run with the control in the failure mode. Transient data recording systems are to be on when switching the control from the automatic to the failure mode, and also when switching back to the automatic mode from the failure mode.

b. Incremental Torque Run. This run shall consist of 96 minutes including:1. Eight minutes at the maximum output shaft speed and the output shaft

torque available at maximum test temperature setting or at the maximum continuous torque limit, whichever is less.

2. Eight minutes at the minimum output speed and the maximum testtemperature setting or at the maximum continuous output shaft torque limit, whichever is less.

3. Eight minutes at the minimum output shaft speed and 90 percent of theoutput shaft torque obtained in (2).

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4. Eight minutes at the maximum output shaft speed and 90 percent of the output shaft torque obtained in (1).

5. Eight minutes at the maximum output shaft speed and 80 percent of the output shaft torque obtained in (1).

6. Eight minutes at the minimum output shaft speed and 80 percent of the output shaft torque obtained in (2).

7. Four minutes at the minimum output shaft speed and 60 percent of the output shaft torque obtained in (2).

8. Four minutes at the maximum output shaft speed and 60 percent of the output shaft torque obtained in (1).

9. Four minutes at the maximum output shaft speed and 40 percent of the output shaft torque obtained in (1).

10. Four minutes at the minimum output shaft speed and 40 percent of the output shaft torque obtained in (2).

11. Four minutes at the minimum output shaft speed and 20 percent of the output shaft torque obtained in (2).

12. Four minutes at the maximum output shaft speed and 20 percent of the output shaft torque obtained in (1).

13. Four minutes at the maximum output shaft speed and 10 percent of the output shaft torque obtained in (1).

14. Four minutes at the minimum output shaft speed and 10 percent of the output shaft torque obtained in (2).

15. Four minutes at the minimum output shaft speed and at zero output shafttorque.

16. Four minutes at the maximum output shaft speed and at zero output shafttorque.

17. Four minutes at idle.18. Four minutes at idle and at zero output shaft speed.

If the test data indicates the existence of critical compressor or turbine vibration conditions within the operating speed range of the engine between idle and maximum test temperature setting, at the option of the Using Service, the following shall be substituted for 48 minutes of the incremental torque run of each cycle to be chosen by the Using Service.

1. Sixteen minutes at gas generator rotor and output shaft speeds whichproduce the critical vibration conditions.

2. Eight minutes at gas generator rotor and output shaft speeds 2.0 percentmore than those which produce the critical vibration conditions.

3. Eight minutes at a gas generator rotor speed 2.0 percent more and anoutput shaft speed 2.0 percent less than those which produce the critical vibration conditions.

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4. Eight minutes at gas generator rotor and output shaft speeds 2.0 percent less than those which produce the critical vibration conditions.

5. Eight minutes at gas generator rotor speed 2.0 percent less and outputshaft speed 2.0 percent more than those which produce the critical vibration conditions.

c. Power Transient Run. This run shall consist of 39 minutes of power transients.1. Four minutes at no load condition followed by one minute at maximum test

temperature setting at the output shaft speed specified for the maximum rating in Table IA. Repeat the above for a total of 20 minutes.

2. The remaining 19 minutes of the run shall consist of one minute at maximum testtemperature setting at the output shaft speed specified for the maximum rating in Table IA, followed by an immediate decrease to no load condition. As soon as the engine reaches the no load condition the engine power shall be increased to the temperature and output shaft speed associated with the previous condition and maintained at this power level for a period of one minute before repeating the cycle.

d. Incremental Power Run. This run shall consist of 9 minutes of operation in thesequence of condition and time duration as follows: One minute idle, three minutes intermediate test temperature, one minute at maximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, three minutes intermediate test temperature, and one minute idle.

e. Intermediate Run. This run shall consist of 30 minutes:1. Fifteen minutes at the intermediate test temperature setting at the

maximum governed shaft speed. 2. Fifteen minutes at intermediate test temperature at the minimum output

shaft speed. Output torque shall be at least as high as that specified in Table IA for the intermediate rating. During the 5th, 10th, 15th, 20th, 25th, 30th, 35th, 40th, 45th, and 50th cycles, the first 26 minutes of this run shall be with the control in the failure mode. Transient data recording systems are to be on when switching the control from the automatic to the failure mode and also when switching back to the automatic mode from the failure mode.

f. Maximum Continuous Run. This run shall consist of 20 minutes:1. Ten minutes maximum continuous test temperature setting at maximum

governed shaft speed. 2. Ten minutes maximum continuous test temperature at the minimum output

shaft speed. Output torque shall be at least as high as that specified in Table IA for the maximum continuous rating.

g. Intermediate Run. This run shall consist of 15 minutes at intermediate testtemperature setting at intermediate rated output shaft speed.

h. Maximum Continuous Run. This run shall consist of 15 minutes at maximumcontinuous test temperature setting at maximum continuous rated output shaft speed.

i. Intermediate-Maximum Run. This run shall consist of six periods of 5 minuteseach alternating between:

1. Five minutes at intermediate test temperature setting at intermediate ratedoutput shaft speed.

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2. Five minutes at maximum test temperature setting at maximum rated output shaft speed.

j. Maximum Continuous Run. This run shall consist of 16 minutes at maximum continuous test temperature setting at maximum continuous rated output shaft speed.

k. Intermediate-Maximum Run. This run shall consist of 30 minutes:1. Fifteen minutes at intermediate test temperature setting at intermediate

rated output shaft speed. 2. Ten minutes at maximum test temperature setting at maximum rated

output shaft speed followed by five minutes at intermediate test temperature setting at intermediate rated output shaft speed. At 5 minute intervals during the run, the anti-icing controls shall be operated for one minute with maximum anti-icing bleed air.

4.6.1.3.1 Starts. A minimum of 600 starts, commencing with the first test cycle of 4.6.1.3, shall be made on each endurance test engine. In addition to the 600 endurance test starts, there shall be 20 false starts (a starting sequence without benefit of light-off followed immediately after the permissible engine draining time by a successful start) and 20 restarts (a start within a maximum of 14 minutes time from shutdown). Starts shall be performed with a starter that is acceptable to the Using Service. The engine shall be started and shutdown not less than six times each endurance test cycle. Of the at least 600 endurance starts, 220 shall be accomplished following varied regulated shutdown periods. Those starts at the beginning of each endurance cycle shall follow a shutdown period of at least two hours. The shutdown period for 36 starts shall be regulated to provide intervals between starts of 2.5 minutes, 5.0 minutes, 7.5 minutes, and increasing, thereafter, by 2.5 minute increments up to and including 90 minutes for the 36th start. Each of the 36 regulated shutdown periods shall be preceded by immediate engine shutdown, without being held at idle, after engine operation for a duration of not less than two minutes at intermediate test temperature setting. For 134 starts, the shutdown period shall be regulated to provide an interval between starts of no less than 45 minutes. The shutdown period for the remaining at least 380 starts need not be controlled. During all endurance starts, immediately after the engine has reached stabilized idle speed, an acceleration to the next scheduled endurance test condition shall be accomplished by a control power signal change from idle to the appropriate test temperature setting.

4.6.1.3.2 Contingency Power Qualification Rating. The engine contingency power capability shall be qualified by four periods of operation at contingency power test temperature. Each operation at contingency power shall be preceded by and followed by a period of operation at maximum continuous power. The duration of each period of operation shall be 3t0.677 where "t" is the rating time for contingency power. All times are in minutes. This will be demonstrated on the endurance test engine after completion of the endurance test.

4.6.1.4 Recalibrations

4.6.1.4.1 Engine Recalibration. After completion of the tests specified in 4.6.1.3 through 4.6.1.3.3, a re-calibration in accordance with 4.6.1.2.2 and 4.6.1.2.3 shall be conducted on each endurance test engine. The re-calibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature conditions, the values of output shaft power obtained during the initial calibration. The re-calibration may be preceded by a specified run during which the cleaning procedure of

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3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration.

4.6.1.4.2 Component Re-calibration. Functional bench calibration/checks shall be in accordance with 4.6.1.2.1.

4.6.1.5 Engine Disassembly and Inspection. Each engine completing the endurance test shall be completely disassembled for examination of all parts. Prior to cleaning, the engine parts shall be given a “dirty inspection” for evidence of leakage, oil coking, unusual heat patterns, and abnormal conditions. The “dirty inspection” shall be completed before any parts are cleaned. The engine parts shall then be cleaned and a “clean inspection” shall then be performed. Engine part measurements shall be taken to determine excessive wear and distortion. These measurements shall be compared with the engine manufacturer’s drawing dimensions and tolerances and with similar measurements made prior to the test. Inspection techniques may also include but not be limited to: magnetic particle, fluorescent penetrate, X-ray, and ultrasonic. During the “clean inspection” a visual examination and condition assessment shall be conducted. Upon completion of “clean inspection,” the Using Service shall be provided all results of nondestructive tests and recommendations for modification or redesign of deficient parts. The Using Service shall be notified of the inspection commencement date prior to each inspection. The following data shall be made available to the Using Service during both inspections:

a. Inspection forms filled out by the contractor listing all observed deficiencies and judgment as to the part’s return-to-service capability.

b. Tabulation of all parts found deficient.c. Detailed configuration list of the component or system tested.d. Test logs and list of test events.e. Spectrometric oil analysis report.f. Other measurements and data as deemed necessary by the Using Service.

4.6.1.6 Not used.

4.6.1.7 Endurance Test Completion. The endurance test will be considered to be satisfactorily completed when the engine has completed the endurance test of 4.6.1 and during the final engine recalibration, the steady state first stage turbine rotor inlet gas temperature does not exceed a value of the gas temperature obtained for the initial calibration plus 30 percent of the difference between the maximum allowable (minimum endurance) steady state first stage turbine rotor inlet gas temperature and the rated temperatures specified in Table IV for all ratings; the corrected specific fuel consumption does not exceed 105 percent of the initial calibration values; the engine meets all other specified performance requirements which can be checked by the calibration procedure; and, , the test engine and components are meeting the requirements of this plan at the end of the tests, and teardown inspections do not disclose part failures or impending failures. Parts will not be judged to have passed the endurance test until they have successfully completed the endurance test. First stage turbine rotor inlet gas temperature deterioration and SFC increase shall be determined at the same rated output shaft powers.

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4.6.2 Engine Component Tests. The following tests shall be conducted on all components listed under 3.1.3 of this AQP. All components shall conform to the same parts list and configurations as those used on the endurance test engines.

4.6.2.1 Previous Component Approval. Engine control and accessory components requiring testing as specified herein may have these tests waived at the option of the Using Service, if the component has been previously approved by the Using Service for use on another engine. All such components must conform to the same parts list and configuration as the components previously approved. In order to receive consideration for test waiver, a similarity argument report in accordance with the requirements of 4.4.4.2 shall be prepared and submitted to the Using Service for approval.

4.6.2.2 Simulated Operational Component Tests. The following tests pertain to the fuel system, ignition system, engine anti-icing system, and engine control system, including temperature sensing and actuation components. For each component listed in section 3.1.3, a single serialized unit shall be used for the testing specified in the following subparagraphs. Tests shall be conducted in the following order:

a. Low Lubricity Fuel Test (4.6.2.2.3.1)b. High Temperature (4.6.2.2.3.2)c. Room Temperature & Contaminationd. Low Temperaturee. Temperature & Vibration Cyclingf. Engine Fuel Pump System Cavitation Endurance

Tests shall be conducted on the same test assemblies, consisting of groups of related components so arranged and interconnected as to simulate their normal relationship and function on the engine. However, subassemblies or components of a system may be tested separately if such separation does not prevent simulation of the complete function of the components or subassemblies. Insofar as practicable, components shall be mounted in their normal position as mounted on the engine.

4.6.2.2.1 Component Acceptance Test. Components shall be acceptance tested IAW procedures developed by the contractor and approved by the Using Service. All component ATP results shall be recorded and provided to the Government in test reports, to substantiate component compliance with performance specifications. The contractor shall provide a list of components that do not require an ATP. This list shall be subject to approval by the Using Service. Components not requiring ATP shall be operated under normal conditions to demonstrate satisfactory functioning and compatibility with other system components. Once component testing has begun, it shall not be acceptance tested, recalibrated or disassembled without specific prior approval of the Using Service.

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4.6.2.2.2 Component Test Procedures. All components shall be cleaned of oil, grease, or other corrosion preventive compounds used for preparation for storage, prior to the start of testing. Test assemblies or components shall be subjected to operating loads simulating those encountered on the engine. Sufficient instrumentation shall be provided to indicate the performance of each component and to indicate that the functional relationships of components are maintained as required by the applicable test schedule or performance specification. Functional checks shall be performed at the end of each test or group of tests and at other times, at the option of the contractor, to verify that the calibrated component has not changed its calibration beyond allowable service limits and that the function of is unimpaired. All components shall be supplied with such fluids as they normally handle or contact, except components normally in contact with fuel will be supplied with test fluids as specified for the individual tests. All shaft driven accessories shall be operated under maximum allowable axial and angular misalignment conditions at the drive pad.

4.6.2.2.2.1 Component Test Cycles (excluding Ignition System). All engine components shall have a test cycle defined by the contractor and submitted to the Using Service in the pretest data. Except for the ignition system (4.6.2.2.2.2), test cycles shall be consistent with the following requirements:

a. Each component shall pass through its maximum range of functions at least once during each cycle.

b. Components in test assemblies shall function in their normal sequence of operation on the engine.

c. Cycling shall be controlled by varying simulated inputs to the test assembly or component. Pilot controlled inputs, such as the control power signal shall be varied in single step changes over their extreme range and shall not be changed again until all variables have reached the demanded values. Engine supplied inputs shall be varied in their usual relations to component outputs.

d. Input variables substantially independent of other control inputs, such as altitude pressure, shall be cycled at a rate faster or slower than the basic functional cycle so that every component shall eventually have accomplished each part of its function at each value of the independent variables.

e. When manual or automatic transfer from one mode of operation to another is provided, the manual or automatic means shall be used to obtain transfer during the cycle.

f. Components designed to prevent the engine from exceeding its operating limits, but which are not actuated by simulated operation, shall be actuated at least once during each of the first ten cycles of the High, Room Temperature and Low Test (4.6.2.2.3.2, 4.6.2.2.3.3, 4.6.2.2.3.4). The pretest data shall include a list of inputs to be cycled, corresponding ranges for each input, and procedures to be used in testing. Disturbing functions such as variations in fuel pressure and bleed airflow shall be included in the list of inputs. Appropriate recording of input and output parameters versus time, suitable for detection of deviation from performance specifications (e.g., ATP) shall be taken throughout the test.

4.6.2.2.2.2 Ignition System Test Cycle. The self contained ignition system or component test assemblies shall be tested IAW the following applicable cycle. For the purpose of these tests, the minimum and maximum voltages and frequencies shall correspond to those extreme conditions permitted on the engine for

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satisfactory functioning of the ignition system. The engine contractor shall submit, in the test report, the ignition system’s output characteristics. The ignition system shall be tested with the same number of igniters connected as used on the engine. The igniters shall be installed in a suitable chamber and the chamber shall be purged with air or nitrogen at a rate specified in the pretest data. The chamber pressure shall be regulated to simulate the internal engine pressure from minimum windmilling pressure to the maximum pressures the igniters will see in the engine operating envelope. For the high temperature test, the complete ignition system shall be placed in a suitable chamber at 95°C (203°F) ambient temperature or the maximum component limiting temperature of 3.1.2.8.1, whichever is hotter. The ignition system shall be tested to the applicable cycle listed below, depending upon the design type:

Light OFF Duty Cycle Continuous Duty Cycle On Off On Off

a. NOM voltage 1 minute 1 minute 60 minutes 20minutes b. MAX voltage 1 minute 1 minute 40 minutes 20minutes c. MIN voltage 1 minute 1 minute 30 minutes 20minutes d. NOM voltage 1 minute 60 minutes 40 minutes 10 minutes

4.6.2.2.3 Order of Testing. The sequence of tests shall be as specified in paragraph 4.6.2.2. Any proposed deviation from the specified sequence shall require approval by the Using Service

4.6.2.2.3.1 Low Lubricity Fuel Test. Fuel system assemblies or components normally requiring fuel for self-lubrication shall undergo a simulated mission operational test for at least 100 hours. The test shall be in accordance with ARP 1797A except that the components shall be subjected to twenty cycles of five hours of operating time followed by one hour of non-operating time. The fuel system components shall be subjected to an acceptance test prior to and after the low lubricity test. Performance shall conform to new part standards. The fuel shall be in accordance with MIL-PRF-5624, grade JP-4 degraded to provide the lubricity value equivalent to a 0.030 - 0.035 inch Wear Scar Diameter (WSD) as defined by ASTM D5001-10.

4.6.2.2.3.2 High Temperature. The test assemblies or component shall be operated as specified in the following subparagraphs.

a. Engine Components (excluding Electronic Assemblies, components with nomoving parts, or Ignition System). Each test assembly or component shall be functionally cycled in accordance with the appropriate test cycle of 4.6.2.2.2 for 100 hours or 600 cycles, whichever represents the longer period. During this cycling, ambient and fluid temperatures shall be maintained as follows:

1. The ambient temperature shall be maintained at 85°C (185°F) for 60minutes. The ambient temperature shall then be increased, within five minutes, to the maximum allowable temperature for the component specified in 3.1.2.8.1 and maintained at this temperature for 120 minutes. The ambient temperature shall then be returned to 85°C (185°F) within five minutes. This cycle shall be repeated until completion of the test.

2. For components normally in contact with fuel, it shall be as specified in theengine specification, but with an aromatic content of at least 25 percent. Toluene may be added

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to the fuel to attain the required aromatic content. Fuel shall be supplied controlled to the maximum temperature specified in 3.7.3.1.3.3.

3. Other fluids used for cooling or control purposes shall be maintained at their maximum allowable temperatures.

b. Electronic Assemblies. The high temperature test need not be performed, since it is included as part of the temperature-vibration cycling test defined in 4.6.2.2.3.5.

c. Components with No Moving Parts. The high temperature test need not be performed, since it is included as part of the temperature-vibration cycling test defined in 4.6.2.2.3.5.

d. Ignition System. The ignition system shall be operated for either for 100 light off duty cycles or 25 continuous duty cycles IAW 4.6.2.2.2.2 and at the maximum component limiting temperature given in 3.1.2.8.1. A 30 minute shut-down following each light off duty cycle, or a two hour shutdown following each continuous duty cycle, shall be performed at the temperature conditions corresponding to the requirements of 3.1.2.8.1c. At the conclusion of testing, checks shall be made of insulation resistance, over voltage capability, igniter output energy, and spark rate.

4.6.2.2.3.3 RoomTemperature & Contamination.

Each test assembly or component shall undergo functional cycling in accordance with the following subparagraphs. During the testing, fuel filters, if furnished with the engine fuel system, shall be serviced as recommended by the engine manufacturer, but at intervals representing a cumulative fuel flow equivalent to not less than that obtained in 12 hours of continuous operation at intermediate rated power.

a. Components utilizing bleed air or requiring pneumatic input signals. : Shall be subjected to air at pressure and temperature values corresponding to those occurring throughout the range of engine operation. During the first hour and each succeeding tenth hour of testing, this air shall be contaminated as follows:

1. Three parts per million engine lubricating oil by mass.2. A salt concentration of 0.2 parts salt (NaCl) per million parts of air by mass

(salt shall be introduced using a 4.0 percent water solution). 3. Distilled water to saturate the air at 52C (126F) at an ambient absolute

pressure of 14.7 psi. 4. ISO 12103-A1 Fine Test Dust, 1.46 x 10-4 pound of quartz per pound of

air. b. Engine Components (excluding Electronic Assemblies, components with no

moving parts. or Ignition System). Each test assembly or component shall undergo functional cycling in accordance with the appropriate test cycle of 4.6.2.2.2 for at least 300 hours or 1800 cycles, whichever represents the longer period. Control of ambient or fluid temperatures shall not be required during this test. Test assemblies normally in contact with fuel shall be supplied with fluid conforming to MIL-PRF-7024F or any of the primary fuels listed in 3.7.3.2.1. The functional cyclic test shall be run in six 50-hour segments. During the first forty-five hours of each segment, the fluid shall be contaminated down-stream from the fluid tank with at least the amount of contaminant specified in Table XI. During the remaining five hours of each segment, the fluid shall be contaminated with at least the amount of contaminant specified in Table X. The solid contaminant shall not be recirculated. The test assembly shall remain idle for at least 18 hours

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following the second and fifth segments and for at least 6 hours following the first, third, and fourth segments. During these idle times, the assembly shall contain the contaminated residual test fluid from the previous test segment. The idle time shall not count as part of the cyclic test time.

c. Electronic Assemblies. The room temperature test need not be performed, since it is included as part of the temperature cycling test defined in 4.6.2.2.3.5.

d. Components with No Moving Parts. The room temperature test need not be performed, since it is included as part of the temperature cycling test defined in 4.6.2.2.3.5.

e. Ignition System. The ignition system shall be operated at an ambient temperature between 16°C (61°F) and 38°C (100°F). Throughout the test, broadband random background vibration shall be applied. Unless otherwise specified, the applied vibration spectrum shall be 0.02 g2/Hz over a bandwidth of 5 to 2000 Hz for airframe-mounted components, and 0.03 g2/Hz over a bandwidth of 15 to 2000 Hz for engine-mounted components. The system shall be operated in accordance with either the light off duty cycle or the continuous duty cycles outlined in 4.6.2.2.2.2. For the light off duty cycle, the system shall be operated for 300 cycles IAW the schedule outlined in 4.6.2.2.2.2. For the continuous duty cycle, the system shall be operated for 75 cycles. IAW the schedule outlined in 4.6.2.2.2.2.

4.6.2.2.3.4 Low Temperature. Upon completion of the room temperature test, each test assembly or component shall be soaked in an ambient temperature of lower than -54°C (-65°F) for a minimum period of ten hours. Upon completion of soaking, the soaking temperature of -54°C (-65°F) shall be maintained while each test assembly or component is operated as detailed below. During the entire low temperature test, fluid conforming to MIL-PRF-5624U, grade JP-4 shall be present in each test assembly or component part normally coming in contact with fuel. Prior to each cycling period the test fluid temperature shall be reduced to below -54°C (-65°F). Other fluid temperatures may rise as anticipated in service operation under similar ambient conditions. If -29°F is reached before completion of a cycling period, the cycling shall be stopped and restarted when the fluid temperature has been reduced to below -54°C (-65°F).

a. Engine Component (excluding Electronic Assemblies, components with no moving parts, or Ignition System,). Each assembly or component shall be functionally cycled in accordance with the appropriate test cycle of 4.6.2.2.2 for at least a total of 20 hours or 120 cycles, whichever represents the longer period. The test shall consist of at least ten separate runs with the cycle sequence of operation in each run as defined in 4.6.2.2.2. Prior to each test period, the test fluid temperature shall be reduced to below -54°C (-65°F) for a minimum of one hour, after which functional cycling may commence. While cycling, fluid temperatures may rise as anticipated in service operation under similar ambient conditions. If -34°C (-29°F) is reached before completion of a test period, the test shall be stopped and restarted when the fluid temperature has been reduced to below -54°C (-65°F) for a 15-minute period.

b. Electronic Assemblies. The low temperature test need not be performed, since it is included as part of the temperature-vibration cycling test defined in 4.6.2.2.3.5.

c. Components with No Moving Parts. The low temperature test need not be performed, since it is included as part of the temperature-vibration cycling test defined in 4.6.2.2.3.5.

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d. Ignition System. The ignition system shall be tested at an ambient temperature of -54°C (-65°F). The system shall be operated in accordance with either the light off duty cycle or the continuous duty cycle outlined in 4.6.2.2.2.2.For the light off duty cycle, the system shall be operated for 12 cycles IAW the schedule outlined in 4.6.2.2.2.2 followed by a 10 hour minimum inoperative soaking period, and a final 12 cycles. For the continuous duty cycle, the system shall be operated for 3 cycles IAW the schedule outlined in 4.6.2.2.2.2 followed by a 10 hour minimum soaking period, and a final 3 cycles.

4.6.2.2.3.5 Temperature & Vibration Cycling. Each electronic assembly or other component without moving parts shall undergo functional cycling in accordance with the appropriate test cycle of 4.6.2.2.2 for a period of 100 cycles or 300 hours, whichever is longer. The contractor shall propose an ambient and cooling media (if applicable) temperature profile that generally replicates component usage in the intended engine/airframe application. The profile shall be approved by the Using Service. The item under test shall be monitored at all times when electrical power is applied. However, performance in compliance with the item under test specification is only required when the unit is operating between the minimum and maximum allowable temperatures. Operation outside these limits shall not result in damage to the item under test, nor shall it result in the item under test indicating any failure condition to other subsystems. Once the temperature has returned to the operating limits, the item under test shall meet its performance requirements without requiring power cycling or other means of reset.

a. The ambient temperature shall be cycled between –54°C (-65°F) and the maximum allowable temperature for the component specified in 3.1.2.8.1. The ambient temperature profile described in the following subparagraphs provides general guidance for the contractor in defining the proposed profile.

1. The ambient temperature cycle shall begin and end at 25°C (77°F)..2. The Steady-State dwells shall occur at the following temperatures, in the

order specified: (a) 25°C (77°F)(b) Maximum operating(c) Maximum non-operating(d) Maximum operating(e) 25°C (77°F)(f) Minimum operating(g) Minimum non-operating(h) Minimum operating

3. All steady-state dwells shall be 30 minutes minimum.4. Transitions between operating temperatures shall be 5°C (9°F) per minute,

minimum. If the equipment is engine mounted, transitions in the increasing direction shall occur at 10°C (18°F) per minute, minimum.

5. Transitions between operating and non-operating temperatures shall be2°C (3.6°F) per minute, maximum.

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b. If the component is electrically powered, power shall be applied and removed in accordance with typical component usage in relation to the imposed ambient temperature profile.

1. When the component is at its maximum allowable temperature, applied voltage shall be at the maximum steady state voltage of 31.5 Vdc. If the component under test is powered from another source, the applied power shall be no less than the maximum allowable specified in the component specification. If multiple sources power the component, the applied power shall be from the source which maximizes the internal power dissipation of the component.

2. When the component is at its minimum allowable temperature, applied voltage shall be at the minimum steady state voltage of 12 Vdc, whichever is the lower. If the component under test is powered from another source, the applied power shall be no greater than the minimum allowable specified in the component specification. If multiple sources power the component, the applied power shall be from the source which minimizes the internal power dissipation of the component.

3. When the maximum non-operating temperature is reached (simulating a soakback condition), electrical power shall be removed.

4. At the beginning of the transition from the minimum operating temperatureto the minimum non-operating temperature, electrical power shall be removed to assist in cooling.

5. At the beginning of the transition from the non-operating temperature tothe operating temperature, electrical power shall be applied in accordance with b (1) or b (2).

c. If fluid or air is used for cooling the component, or is used internal to the component for other purposes (for example, a combustor pressure sensor within an engine control system), the following apply:

1. It shall be maintained at its maximum allowable temperature, concurrent with the maximum allowable ambient.

2.) It shall be maintained at its minimum allowable temperature, concurrent with the minimum allowable ambient.

d. During all portions of the test, except for those times specified below, broadbandrandom background vibration shall be applied simultaneously with the requirements specified in a, b and c. Unless otherwise specified, the applied vibration spectrum shall be 0.02 g2/Hz over a bandwidth of 5 to 2000 Hz for airframe--mounted components and 0.03 g2/Hz over a bandwidth of 15 to 2000 Hz for engine-mounted components.

1. During the steady-state dwell at the specified maximum non-operatingtemperature of the component, or during transitions to/from that condition.

2. During the steady-state dwell at the specified minimum non-operatingtemperature of the component, or during transitions to/from that condition.

4.6.2.2.3.6 Engine Fuel System Cavitation Endurance. For the engine fuel system cavitation test, the portion of the fuel system from the engine fuel inlet to the engine main fuel pump inlet (if they are not the same) shall be included in the test assembly. This shall include engine mounted boost pumps, lines, fittings, filters, etc., as well as any elements of the engine fuel system downstream of the main fuel pump which might have an effect on endurance. Except as stated below, the test procedure shall be in accordance with ARP 492C. Prior to the start of this test, the system shall have had fuel passed through it at maximum

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continuous engine fuel flow for five hours contaminated with at least the amount specified in Table X. Un-weathered clean fuel may be used to conduct the test. Each cycle shall consist of 45 minutes at the maximum speed and flow required by the engine at the altitude corresponding to the fuel tank pressure, followed by 15 minutes at the minimum idle speed and flow required by the engine at the altitude corresponding to the fuel tank pressure. Backpressure to the engine main fuel pump shall be set in accordance with the operating conditions. The test shall consist of two parts. For part one, subparagraphs a through f, the fuel shall conform to MIL-DTL-83133H, grade JP-8. For subparagraphs g through l, it shall conform to MIL-DTL-5624U, grade JP-4. A restriction shall be introduced ahead of the engine fuel inlet to provide a fuel pressure at the engine fuel inlet, during the maximum speed and flow portion of each cycle, of not more than 1 psi above the true vapor pressure of the fuel or not more than the fuel pressure which corresponds to a fuel vapor to liquid ratio of 1.0, whichever is the higher pressure. This restriction shall be maintained during the idle speed and flow portion of each cycle. Part one shall consist of:

a. Twenty cycles with pressure on the fuel tank maintained at 14.7 psia and with an engine fuel inlet temperature of 71°C (160°F).

b. Twenty cycles with pressure on the fuel tank maintained at 13.7 psia and with an engine fuel inlet temperature of 71°C (160°F).

c. Twenty cycles with pressure on the fuel tank maintained at 9.3 psia and with an engine fuel inlet temperature of 71°C (160°F).

d. Twenty cycles with pressure on the fuel tank maintained at 13.6 psia and with an engine fuel inlet temperature of 57°C (135°F).

e. Twenty cycles with pressure on the fuel tank maintained at 11.4 psia and with an engine fuel inlet temperature of 57°C (135°F).

f. Twenty cycles with pressure on the fuel tank maintained at 7.0 psia and with an engine fuel inlet temperature of 57°C (135°F).

g. Twenty cycles with pressure on the fuel tank maintained at 11.9 psia and with an engine fuel inlet temperature of 43°C (109°F).

h. Twenty cycles with pressure on the fuel tank maintained at 9.7 psia and with an engine fuel inlet temperature of 43°C (109°F).

i. Twenty cycles with pressure on the fuel tank maintained at 5.4 psia and with an engine fuel inlet temperature of 43°C (109°F).

j. Twenty cycles with pressure on the fuel tank maintained at 11.0 psia and with an engine fuel inlet temperature of 29°C (84°F).

k. Twenty cycles with pressure on the fuel tank maintained at 8.9 psia and with an engine fuel inlet temperature of 29°C (84°F).

l. Twenty cycles with pressure on the fuel tank maintained at 5.4 psia and with an engine fuel inlet temperature of 29°C (84°F). For part two, the fuel shall conform to MIL-DTL-83133H, grade JP-8. A restriction shall be introduced ahead of the engine fuel inlet to provide a fuel pressure at the engine fuel inlet, during the maximum speed and flow portion of each cycle, of not more than 35 percent of the fuel tank pressure or 2.2 psi absolute, whichever is the higher pressure. This restriction shall be maintained during the idle speed and flow portion of each cycle. The fuel temperature at the

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engine fuel inlet shall be maintained to provide a fuel viscosity of not less than 12 centistokes. Part two shall consist of:

a. Twenty cycles with pressure on the fuel tank maintained at 14.7 psia absolute.b. Twenty cycles with pressure on the fuel tank maintained at 8.7 psia absolute.c. Twenty cycles with pressure on the fuel tank maintained at 4.4 psia absolute.

4.6.2.2.3.7 Fuel System Bubble Ingestion. The engine fuel system must demonstrate capability of continuously ingesting slugging flow at its inlet without metered fuel flow interruption throughout the engine operating envelope and power range. The flow structure shall generally be alternating liquid fuel and fuel vapor (dissolved air and fuel vapor released from the fuel) in 12 inch lengths respectively, in a vertical pipe of diameter equal to the same cross sectional area as the engine fuel inlet. Except as stated below, the test procedure shall be in accordance with ARP4028.

The test shall be performed using actual engine fuel pumping elements (e.g. boost pump, main pump, etc…) but other elements of the engine fuel system may be simulated provided it can be shown that inclusion of these components does not negatively impact the result of the test. Approval to use less than a complete engine fuel system for this test shall be obtained prior to test commencement.

Each cycle shall consist of a 12 inch slug of fuel vapor, followed by a 12 inch slug of liquid fuel. Maximum Continuous Power speed is defined as the fuel pump speed that is associated with the maximum continuous power of the engine. Minimum Ground Idle speed is defined as the fuel pump speed that is associated with the minimum ground idle speed of the engine. Subparagraphs a through d shall use JP8. Subparagraph e through h shall use JP4.

a. Fifteen (15) cycles with fuel pump at maximum continuous power speed, fuel tankpressure maintained at 14.7 PSIA and engine fuel inlet temperature at 71°C (160 °F).

b. Fifteen (15) cycles with fuel pump at minimum ground idle speed, fuel tankpressure maintained at 14.7 PSIA and engine fuel inlet temperature at 71°C (160 °F).

c. Fifteen (15) cycles with fuel pump at maximum continuous power speed, fuel tankpressure maintained at 10.11 PSIA and engine fuel inlet temperature at 71°C (160 °F).

d. Fifteen (15) cycles with fuel pump at minimum ground idle speed, fuel tankpressure maintained at 10.11 PSIA and engine fuel inlet temperature at 71°C (160 °F).

e. Fifteen (15) cycles with fuel pump at maximum continuous power speed, fuel tankpressure maintained at 14.7 PSIA and engine fuel inlet temperature at 29°C (84 °F).

f. Fifteen (15) cycles with fuel pump at minimum ground idle speed, fuel tankpressure maintained at 14.7 PSIA and engine fuel inlet temperature at 29°C (84 °F).

g. Fifteen (15) cycles with fuel pump at maximum continuous power speed, fuel tankpressure maintained at 10.11 PSIA and engine fuel inlet temperature at 29°C (84 °F).

h. Fifteen (15) cycles with fuel pump at minimum ground idle speed, fuel tankpressure maintained at 10.11 PSIA and engine fuel inlet temperature at 29°C (84 °F).

4.6.2.2.4 Component Acceptance Test or Recalibration. Upon completion of specific tests, component ATP shall be repeated utilizing the same configuration-controlled test procedure as was used in 4.6.2.2.1. Components not subjected to ATP shall be operated under normal operating conditions to demonstrate satisfactory functioning and compatibility with other system components. All component ATP results shall be recorded and provided to the Government in test reports, as necessary to substantiate component

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compliance with performance specifications. All components shall then be completely disassembled and inspected for indications of failure, impending failure or excessive wear.

4.6.2.2.5 Component Test Completion. The simulated operational tests shall be considered to be satisfactorily completed when, in the judgment of the Using Service:

a. During the tests, component performance and function were within establishedlimits.

b. During the tests, there was no fluid leakage from any component other than at thelocation and rate specified in 3.3.6.4.

c. Acceptance tests indicate that the component is still within its allowable servicelimits (if any).

d. The component teardown inspection shows no indication of failed, excessivelyworn, or distorted parts, or impending part failure. Measurements shall be taken and compared with the drawing dimensions and tolerances and with similar measurements made prior to the test.

4.6.2.3 Environmental Component Tests. The following tests apply to all components or subcomponents as identified in 3.1.3 in accordance with Tables XX-A and XX-B herein. These tests may be conducted on test assemblies or individual components. For each component specified in section 3.1.3, a single serialized unit shall be used for the testing per section 4.6.2.3 and subparagraphs. At the option of the contractor, two units may be used to complete the test sequence. In this case, tests 4.6.2.3.4.1 through 4.6.2.3.4.8 shall be conducted on one unit, while tests 4.6.2.3.4.9 through 4.6.2.3.4.14 shall be conducted on the other unit.

4.6.2.3.1 Component Acceptance Test or Calibration. Components shall be acceptance tested IAW procedures approved by the Using Service. All component ATP results shall be recorded and provided to the Government in test reports to substantiate component compliance with performance specifications. The contractor shall provide a list of components that do not require ATP. This list shall be subject to approval by the Using Service. Components not requiring ATP shall be operated under normal conditions to demonstrate satisfactory functioning and compatibility with other system components. Once component testing has begun, it shall not be acceptance tested, recalibrated or disassembled without specific prior approval of the Using Service.

4.6.2.3.2 Order of Testing. The sequence of tests shall be as specified in Tables XX-A and XX-B. Any proposed deviation from the specified sequence shall require approval by the Using Service. However, if the testing is conducted on a single serialized unit, then the explosive atmosphere test shall be conducted last

4.6.2.3.3 Component Test Success Criteria The component tests shall be considered to be satisfactorily completed when, in the judgment of the Using Service:

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a. During the tests, component performance and function were within establishedlimits.

b. During the tests, there was no fluid leakage from any component other than thatof a nature and rate specified in the engine specification.

c. During the tests, there was no hang-up or hesitation of any component.d. Acceptance tests indicate that the component is still within its allowable service

limits (if any). e. The component teardown inspection shows no indication of failed, excessively

worn, or distorted parts, or impending part failure. Measurements shall be taken and compared with the contractor’s drawing dimensions and tolerances or with similar measurements made prior to the test.

4.6.2.3.4 Component Test Procedures. All components shall be cleaned of oil, grease, or other corrosion preventive compounds used for preparation for storage, prior to the start of testing. Test assemblies or components shall be subjected to operating loads simulating those encountered on the engine. Sufficient instrumentation shall be provided to indicate the performance of each component and to indicate that the functional relationships of components are maintained as required by the applicable test schedule or performance specification. Functional checks shall be performed at the end of each test or group of tests and at other times, as an, to verify that the component still meets allowable service limits and that the function of is unimpaired. All components shall be supplied with such fluids as they normally handle or contact, except components normally in contact with fuel will be supplied with test fluids as specified for the individual tests. All shaft driven accessories shall be operated under maximum allowable axial and angular misalignment conditions at the drive pad. When a hermetically sealed component is used in a series of the tests below, the component need not be disassembled for inspection until the last test of such series has been completed. At this time, the components will be inspected for defects or damage which may have been incurred during any of the tests performed. In addition, hermetically sealed components need not be subjected to the explosion-proof, sand and dust, and fungus tests. Prior to disassembly, a test to determine hermetic seal integrity shall be performed. Failure of the hermetic seals during any test shall disqualify that component.

4.6.2.3.4.1 Temperature – High, Low, Transient. Components shall be subjected to temperature tests in accordance with subparagraphs a, b and c, below. Components shall be subjected to pre- and post-test ATPs or functional checks as applicable, and shall be operating and verified for specified performance during the test. Criteria for passing the test shall be as defined in MIL-STD-810G. The three tests specified in the subparagraphs below shall generally be combined so that a single temperature cycle can be defined which transitions rapidly from the hot steady state condition to the cold steady state condition, and back again. The high and low temperature extremes (as well as the transition between these two steady state conditions) may be greater than those encountered in service since the intent is to stress the unit beyond its normal temperature range. If an SOT is being conducted on a separate unit, consideration shall be given to utilizing the defined SOT temperature cycle with modification to increase the steady state limits as well as the rate of transition between the two limits. If no cycle is defined, then the steady state conditions shall be maintained for approximately 2 hours after reaching that temperature, meaning that a ‘cycle’ is approximately 4 hours in duration (neglecting

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transitions). The recommended number of cycles is 5 to 10, for a total test time of approximately 20 to 40 hours. The transitions shall be in the range of 10C to 20C per minute. Because of the unrealistically rapid temperature transitions, performance may not meet specified requirements, in which case allowance must be made by increasing the performance criteria during the transitions. However, once steady state temperatures are reached, measured performance must meet established criteria; typically in the form of the approved ATP. In general, the component shall be operated continuously while it is exposed to the test environment, unless the extremes include soakback conditions which drive the temperature beyond the normal range and during which the component is not expected to be operating in any case. For Electronic Assemblies deriving electrical power from the airframe bus, power shall be left ON during a hot soakback condition to replicate the fact that the airframe power is not turned off when the engine is shut down. For cold soak conditions below –40°C (-40°F), it is necessary to remove electrical power to simulate overnight soaks at –54° (-65°F). After reapplying electrical power to simulate the aircraft being powered up, performance testing can be resumed when the unit temperature recovers to –40°C (-40°F), if so specified in the component specification. Note that if the component to be tested also conveys fluid (fuel, oil or air), then the steady state maximum temperature to which it is exposed may require control of not only the ambient air temperature, but the fluid temperature as well. Furthermore, the ‘worst-case’ temperature may not be maximum ambient nor maximum fluid temperature, but some other combination which yields the maximum internal component temperatures.

a. High Temperature: Component testing shall be in accordance with MIL-STD-810G, Method 501.5, Procedure II (storage/non-operating) and II (operating); constant temperature method.

b. Low Temperature: Component testing shall be in accordance with MIL-STD-810G, Method 502.5, Procedure II (storage/non-operating) and II (operating).

c. Temperature Shock/Transient: Component testing shall be in accordance with MIL-STD-810G, Method 503.5, Procedure I-C, with the number of cycles increased to five (minimum). The test shall consist of an initial steady-state temperature cycle followed by five (5) transient cycles. If the total time period for the five transient cycles is insufficient to completely determine UUT compliance with performance requirements, the number of cycles shall be increased as necessary to establish complete compliance.

For the initial steady-state temperature cycle, each temperature dwell shall be maintained for two (2) hours minimum after the UUT has stabilized at that test temperature. For the subsequent transient temperature cycles, the dwells shall be maintained until the UUT has stabilized at the test temperature, or for ten (10) minutes, whichever is the longer period.

The minimum temperature ramp rate for the steady-state cycle shall be 5°C (9°F) per minute for airframe-mounted components, and 10°C (18°F) per minute for engine-mounted components. For the transient cycles, the respective ramp rates shall be 15°C (27°F) and 20°C (36°F) per minute. During the high-rate transient cycles, unit signal conditioning accuracy may not meet specified requirements, in which case widening the pass/fail limits may be allowable with AED concurrence. However, when the measured UUT temperature rate of change slows to 5°C per minute (airframe-mounted components) or 10°C per minute (engine-mounted components), signal conditioning accuracy must meet specified requirements.

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For electronic assemblies deriving electrical power from the airframe bus, power shall be left ON during a hot soakback condition to replicate that airframe power may be left on after the engine is shut down. For cold soak conditions below -40°C (-40°F), remove electrical power to simulate overnight soaks at -54°C (-65°F). After re-applying electrical power to simulate the aircraft being powered up, performance testing shall be resumed when the unit recovers to the minimum operating temperature as specified in the component specification.

If the component also conveys fluid (fuel, oil or air), then the maximum temperature to which it is exposed to may not be maximum ambient nor maximum fluid temperature, but some other combination which yields the worst-case internal component temperatures. This may require control of not only the ambient air temperature, but the fluid temperature as well.

4.6.2.3.4.2 Vibration – Airframe and Engine. Components shall be subjected to a vibration test in order to:

a. Verify functional performance of the component when exposed to the expected worst-case vibration environment encountered during service.

b. Demonstrate acceptable component service life when exposed to the ‘typical’ vibration environment expected over the life cycle of the equipment. Testing shall be in accordance with MIL-STD-810G, Method 514.6, Procedure I. All components shall be subjected to the requirements specified in paragraph a. If the component is engine mounted, the requirements of paragraph b shall also apply.

a. Category 14, Figure 514.6D-3, Table 514.6D-III “On/Near Drive System Elements”

b. Category 22, Figure 514.6D-10, paragraph 2.11c .”Multiple rotors”

For Category 22 testing, the frequency ranges and amplitudes for the four narrowband random vibration spikes shall be as listed below, unless specified otherwise. Note that as an alternative to these narrowband spikes, the Using Service will consider use of linearly swept sinusoids to represent the gas generator and power turbine rotors, and the accessory gearbox shafting. This alternative approach is described in Annex D, Section 2.11, of MIL-STD-810G.

Any such proposal by the contractor shall be submitted to the Using Service for concurrence, including supporting data or analysis for the recommended amplitude and sweep rate of the sinusoids. In general, the sinusoidal levels described by curve “L”, Figure 514.2-2 of MIL-STD-810C are considered acceptable default levels for these sinusoidal sweeps. Whichever method is selected, the contractor shall be able to use the results of this qualification testing to generate maximum allowable installed vibration levels for the component. These levels are typically measured on-wing with accelerometers in terms of sinusoidal accelerations, velocities, or displacements.

Narrowband Spike Frequency Range ASD, g2/Hz

f0 Gas Generator: Ground Idle to GI + 10% 1.0 (default)

f1 Gas Generator: 90% to 105% 1.0 (default)

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f2 Power Turbine: 95% to 105% 1.0 (default)

f3

Power Turbine: Single Engine GI to Dual Engine GI OR

AGB Shafting: Dominant Frequency + 5%

1.0 (default)

Components may be tested in test assemblies or as individual units; however, components shall be tested to simulate installation orientation or other environmental or physical details. During the test, the component shall be subjected to its maximum limiting temperature (as specified in the engine model specification) and shall be operating to evaluate any functional effects of the vibration on specified performance. Two alternate procedures for qualification are described below; known application environment and unknown application environment. All qualification shall be in accordance with the requirements of paragraph c.

c. Known Application Environment. If the vibration environment is known, the boththe expected ‘worst-case’ vibration environment to be encountered during service and the ‘typical’ (e.g., mission-weighted) vibration environment expected over the life cycle of the equipment shall be defined. Substantiating data for the defined vibration levels shall be presented and obtain concurrence from the Using Service. The test shall be run in two parts, as specified below.

1. Functional Performance. Test for 1 hour per axis at the worst-case levelsdetermined in a, above. Performance shall be monitored, recorded, and verified throughout the entire test period. At the conclusion of testing in all three axes, a calibration or ATP shall be conducted to fully verify component functionality.

2. Service life test, demonstrating component life in accordance with therequirements of 3.3.8.1e, at the typical levels defined in a, above. The component shall be operational and monitoring shall be employed in order to determine if and when a failure occurs for the purpose of analysis. At the conclusion of testing in all three axes, a calibration or ATP shall be conducted to fully verify component functionality. This testing may be accomplished after each axis of test. Test time per axis may be reduced by increasing the test spectrum levels above the defined typical value(s). However, in no circumstances shall the test level be more than:

(a) For sinusoids. 1.5 times the worst-case level or the default testlevels (see paragraph b), whichever is greater.

6

×=

LevelTest

LevelTypicalhoursaxisper peakg

peakgLifeServiceTimeTest

(b) For random profiles. 2.0 times the worst-case level or the defaulttest levels (see paragraph b), whichever is greater.

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4

2

2

×=

LevelTest

LevelTypicalhoursaxisper

Hzg

Hzg

LifeServiceTimeTest

If component functionality can be reliably verified at the service life test spectrum level, and the test level is greater than the expected worst-case environment, then the then the functional test may be omitted. The functional test defined under 1, with the proviso that component performance shall be monitored, recorded, and verified for 30 uninterrupted minutes (minimum) during the first and last hours, and at the midpoint, of the service life test, in each axis.

d. Unknown Application Environment. The default test spectrums described insubparagraphs a and b are to be used if the actual environment(s) are unknown. These default test spectrums are considered to be worst-case and therefore performance requirements must be met under these conditions.

1. Functional Performance. Test for 1 hour per axis. Performance shall bemonitored, recorded, and verified throughout the entire test period. At the conclusion of testing in all three axes, a calibration or ATP shall be conducted to fully verify component functionality.

2. Service Life. Demonstrate component life in accordance with therequirements of the engine model specification. If component life is unspecified, 6000 hours shall be demonstrated, requiring 8 hours of vibration testing per axis IAW the default spectrums. Monitoring shall be employed in order to determine if and when a failure occurs. At the conclusion of testing in each axis, a calibration or ATP shall be conducted to fully verify component functionality, or with the Using Service concurrence, this testing may be accomplished after completion of all three axes. Test time per axis may be adjusted by increasing/decreasing the test spectrum levels, or to accommodate a different service life requirement. However, under no circumstances shall the test level be greater than:

(a) For sinusoids. 1.5 times the default level. The test time per axis shall becalculated as follows:

6

×=

LevelTest

LevelTypicalhoursaxisper peakg

peakgLifeServiceTimeTest

(b) For random profiles. 2.0 times the default level. The test time per axis shallbe calculated as follows:

4

2

2

×=

LevelTest

LevelTypicalhoursaxisper

Hzg

Hzg

LifeServiceTimeTest

3. Segment I. Functional performance test for 2 hours per axis at the levelsdetermined in (2), above. Performance shall be monitored, recorded, and verified throughout the

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entire test period. At the conclusion of testing in all three axes, a calibration or ATP shall be conducted to fully verify component functionality. This testing may be accomplished after each axis of test.

4. Segment II. Service life test for 8 hours per axis at the default test spectrum levels. The component need not be operating during the test, nevertheless it is recommended that a minimum level of monitoring be employed in order to determine if and when a failure occurs for the purpose of analysis.

5. Segment III. Functional performance test for 2 hours per axis at the levels determined in (b), above. Performance shall be monitored, recorded, and verified throughout the entire test period. At the conclusion of testing in all three axes, a calibration or ATP shall be conducted to fully verify component functionality. This testing may be accomplished after each axis of test. If the service life test spectrum level is adjusted to be greater than or equal to the default level, and component functionality can be reliably verified at this higher level, then the contractor may omit the functional portion of the vibration test. In this case, performance of the fully operational component shall be monitored, recorded, and verified for 30 uninterrupted minutes (minimum) during the first and last hours, and at the midpoint, of the test, in each axis.

4.6.2.3.4.3 Impact (Shock). Components shall be subjected to an operational shock test in accordance with MIL-STD-810G, Method 516.6, Procedure I. The shock test spectrum shall be in accordance with Figure 516.6-8 (SRS) and Table 516.6-I, or Figure 516.6-10 (terminal peak sawtooth) and Table 516.6-II. Tests shall be conducted under room ambient conditions in conjunction with the functional vibration test of 4.6.2.3.4.2.c(1) or d(2) utilizing the same test fixture and monitoring/recording equipment. Components shall be subjected to pre- and post-test ATPs or functional checks as applicable, and shall be operating and verified for specified performance during the test.

Components located in a crew compartment shall be subjected to a crash hazard shock test in accordance with MIL-STD-810G, Method 516.6, Procedure V, if it is determined that the component could become a hazard if it breaks free of its mounts during a crash. The shock test spectrum shall be in accordance with Figure 516.6-8 (SRS) and Table 516.6-I, or Figure 516.6-10 (terminal peak sawtooth) and Table 516.6-VII. Tests shall be conducted under room ambient conditions. The component need not be operating during the test nor pass a post-test ATP or functional check. With the approval of the Using Service, a dummy unit may be used if it duplicates the mounting arrangement, overall mass, and moments of inertia of the actual component.

4.6.2.3.4.4 Gunfire Shock. Components shall be subject to gunfire testing in accordance with MIL-STD-810G, Method 519.56. Components shall be subjected to pre- and post-test ATPs or functional checks as applicable, and shall be operating and verified for specified performance during the test. This procedure can be tailored for the aircraft if the location of engine components, type of gun(s) and gun location(s) are known. Tailoring of the procedure shall be approved by the Using Service.

4.6.2.3.4.5 Sustained Acceleration. Components shall be subjected to an acceleration test in accordance with MIL-STD-810G, Method 513.6, Procedure II. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, and shall be operating and verified for specified performance during the test. Test time may be increased beyond the minimum specified if necessary to

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determine proper operation. Recorded test data shall include the applied acceleration g-level (or equivalent) and the performance signals versus time. The data shall show acceleration at 0 g, reaching the specified value, remaining there for at least the specified test time, and then returning to 0 g. If a component has no moving parts and it has passed a vibration test, this test may be waived if an analysis report is provided, subject to approval by the Using Service. The analysis must show that 1), the internal component deflections exhibited during the vibration test are at least as great as those expected to occur during application of the sustained acceleration loads, and 2), the cumulative time spent at these deflections, or above, exceeds the sustained acceleration test time.

4.6.2.3.4.6 Low Pressure (Altitude). Electronic assemblies and electro-mechanical components shall be subjected to a low pressure test in accordance with MIL-STD-810G, Method 500.5, Procedures I (storage) and II (operation), if environmentally sealed. Components shall be subjected to pre- and post-test ATPs or functional checks as applicable, and shall be operating and verified for specified performance during testing in accordance with Procedure II. Unless otherwise specified, the non-operational and operational altitudes and temperatures utilized for Procedures I and II, respectively, shall be as defined in 4.6.2.3.4.1. If the component is located in an area where the local external temperature environment is higher than the default value for Procedure II, the test temperature shall be the worst-case maximum. Unless otherwise specified, the non-operational and operational altitudes and temperatures utilized for Procedures I and II, respectively, shall be as defined below. If the component is located in an area where the local external temperature environment is higher than the default value for Procedure II, the test temperature shall be the worst-case maximum.

Environment Operational Non-Operational Low Temperature -54ºC (-65 ºF) -54ºC (-65 ºF)

High Temperature Airframe Mounted 85ºC (185 ºF) 85ºC (185 ºF) Engine Mounted 149ºC (300 ºF) 174ºC (345 ºF)

Low Pressure Altitude 6,096 m (20,000 ft) 15,240 m (50,000 ft) Temperature -25ºC (-13 ºF) -54ºC (-65 ºF)

4.6.2.3.4.7 Rain. Components shall be subjected to a rain test in accordance with MIL-STD-810G, Method 506.5-, Procedure I (blowing rain). Electronic Assemblies shall be tested to Procedure I, if mounted on the engine, or Procedure III (drip), if located in a portion of the airframe that may be exposed to condensation or rain ingress. Components shall be subject to pre- and post-test ATPs or functional checks as applicable. Components shall be operating and verified for specified performance when tested IAW Procedure III, but need not be operating during testing conducted IAW Procedure I.

4.6.2.3.4.8 Explosive Atmosphere. All electrical components not hermetically sealed, or components with moving parts capable of creating a spark, shall be subjected to explosive atmosphere testing in accordance with MIL-STD-810G, Method 511.5, Procedure I. Components shall be subject to pre-test ATPs or functional checks as applicable, but need not pass any post-test checks, unless specified by the

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Using Service. The test shall be conducted at sea level and at the operational altitudes specified in 4.6.2.3.4.6 above. During the test, components shall have maximum input voltage applied to them and shall be operated continuously at their maximum loads, with ambient air surrounding the component maintained at its maximum allowable limiting temperature. During each altitude condition, all make and break contacts shall be operated at least ten times. If a component utilizes airframe power, power supply transients shall be applied to the component during each altitude condition. At least four of these power supply transients shall be applied during operation of make and break contacts. Power supply transients shall be as stated in 4.6.2.3.4. Ignition components or systems shall be operated continuously. Electrodes of spark igniters shall be mounted in such a manner that the explosive vapor in the test chamber shall not be contacted. Failure criteria shall be as defined in MIL-STD-810G. Additionally, all components shall be shown by analysis or test that no external or internal constituent part, which can plausibly be surrounded by a flammable mixture, can exceed the minimum auto-ignition temperature of JP-4 (+230°C/+446°F) under worst case conditions. If the minimum margin is less than +20°C (+36°F), the contractor shall submit the analysis or test results to the Using Service for approval of the analysis/test assumptions and methodology.

4.6.2.3.4.9 Fungus. Components shall be subjected to a fungus test in accordance with MIL-STD-810G, Method 508.6, or evidence must be provided in an analysis report which indicates that all exposed materials of the component are fungus inert or have been previously tested. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test. Criteria for passing the test shall be as defined in MIL-STD-810G. Analyses reports, shall list all parts and chemical-based substances (including glues, sealants, etc.) employed in the design of the component, and supplier certifications as to the non-nutrient character of the constituent materials.

4.6.2.3.4.10 Humidity. Components shall be subjected to a humidity test in accordance with MIL-STD-810G, Method 507.5, Procedure II – Aggravated Cycle (Figure 507.5-7) Components shall be subjected to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test other than as specifically required by the procedure. Criteria for passing the test shall be as defined in MIL-STD-810G.

4.6.2.3.4.11 Salt Fog. Components shall be subjected to a salt fog test in accordance with MIL-STD-810G, Method 509.5. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test. Criteria for passing the test shall be as defined in MIL-STD-810G.

4.6.2.3.4.12 Sand and Dust. Components shall be subjected to a sand and dust test in accordance with MIL-STD-810G, Method 510.5, Procedures I (blowing dust) and II (blowing sand). Procedure I testing shall be conducted at the component’s specified limiting temperature for a total duration of 6 hours with the component operating and verified for specified performance throughout the test period.

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Procedure II testing shall be conducted at +55°C (+131°F) IAW the duration specified in the procedure. The component need not be operating during this test. Components shall be subject to pre- and post-test ATPs or functional checks as applicable.

4.6.2.3.4.13 Contamination by Fluids. Engine-mounted components, or those located in the engine compartment, shall be subjected to a fluids contamination test in accordance with MIL-STD-810G, Method 504.1, Procedure I, 4.5.5b Intermittent, or evidence must be provided in an analysis report which indicates that all exposed materials of the component are unaffected by the specified fluids or have been previously tested. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test. Criteria for passing the test shall be as defined in MIL-STD-810G. Unless otherwise specified, the fluids identified below shall be utilized for the test; applied individually. Fluids that are determined to be miscible may be applied simultaneously at the highest of the specified fluid temperatures.

Group Type Test Fluid Fluid Temp ±2°C(±4°F)

Fuels Kerosene MIL-DTL-83133H

Aviation turbine fuel JP-8 (NATO F-34)

+70°C(+158°F)

Hydraulic Oils

Hydrocarbon base (synthetic)

MIL-PRF-83282D (NATO H-537) or MIL-PRF-46170D (NATO H-544)

+70°C(+158°F)

Petroleum base MIL-PRF-5606H (NATO H-515) +70°C(+158°F)

Phosphate ester base (synthetic)

SAE AS1421C contaminated with 1% water by weight (Skydrol® LD-4)

+70°C(+158°F)

Lubricating Oils Polyol ester base (synthetic) MIL-PRF-23699F (NATO O-156) +150°C

(+302°F)Transmission

Oils Synthetic base DOD-PRF-85734A +150°C(+302°F)

Solvents & cleaning fluids trans-1,2-dichloroethylene (replaces 1.1.1-Trichloroethane)

+23°C(+73°F)

Deicing & antifreeze fluids Ethylene or propylene glycol mixtures (e.g., A-A-52624A)

+23°C(+73°F)

Runway deicers Potassium-acetate based solution (e.g., Cryotech E-36)

+23°C(+73°F)

4.6.2.3.4.14 Proof Pressure. For any engine component pressurized by fuel, oil or air, the contractor shall demonstrate that the component can continue to operate in accordance with performance specifications after exposure to 1.5 times (fuel/oil) or 2.0 (time (air) the design maximum operating pressure for a period of 10 minutes. At the end of the test period, pressure shall be returned to the design maximum operating pressure and held for a period of 5 minutes. This test shall be performed on the same unit that has first undergone the pressure cycling test of 4.6.2.4.11. The test pressure shall be applied at the component’s limiting temperature. The test may be conducted at a lower temperature if the applied pressure is increased to compensate. Any such

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request by the contractor must be supported by analysis justifying the pressure/temperature combination, and shall be submitted to the Using Service for approval. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test unless required in order to perform the test. During the test, and for a period of 5 minutes immediately following the the component shall exhibit no seepage or leakage from gasket or other sealing surfaces, and there shall be no permanent deformation. The component shall be considered to be fully serviceable and may be used for other qualification testing as concurred by the Using Service

4.6.2.4 Special Components Tests.

4.6.2.4.1 Oil Reservoir. The filler cap and other fittings shall be installed, the tank mounted in a manner similar to that as found on the engine, and the entire test assembly shall be subjected to the following tests:

a. Cyclic Fatigue Test. The oil tank shall be cycled between the minimum andmaximum differential pressure limits of 3.7.7.4.1 at a rate no more than four times per minute for a minimum of 15,000 cycles. The differential pressure to be used in the cyclic fatigue test shall be stated in the pretest data. For the purpose of this test, the differential pressure shall mean the absolute value of the difference between the external and internal pressure of the oil tank. During the first 7,500 cycles, the oil reservoir shall be kept at the nominal operating temperature, and during the last 7,500 cycles the reservoir temperature shall be kept at the maximum oil operating temperature. Throughout this cycling, no leakage or detrimental deformation of the oil reservoir, filler cap, or fittings shall occur.

b. Proof Pressure Test. Upon successful completion of the cyclic fatigue test, thesame oil reservoir assembly shall be subjected to a proof pressure test to demonstrate compliance with 3.7.7.4.1. The proof pressure shall be held for a minimum of 10 minutes with the wall temperature of the oil reservoir at the maximum oil operating temperature. The pretest data shall specify the pressure to be used for the proof pressure test. No leakage or permanent deformation of the oil reservoir, filler cap, or fittings shall occur.

c. Valve Tests. If the oil reservoir assembly incorporates a pressurizing valve orpressure relief valve, the assembly shall be tested to demonstrate proper functioning. pretest data shall specify the procedure to be used.

4.6.2.4.2 Accessory Drive. The engine drive train and the external engine gearboxes which drive engine components and customer accessories shall be subjected to a 300 hour endurance test as specified in paragraph 4.6.1.3. After running the endurance test, the gearbox shall undergo a static torque test wherein the input shaft shall be held stationary and all drive pads, except the starter pad, simultaneously loaded to 150 percent of the maximum static torque values specified in the engine specification for a period of five seconds. During this five second period the starter drive shall be loaded to 250 percent of the maximum starting torque specified in the engine specification for this pad and in a direction that will not unload any other component of the accessory drive train. Following the static torque test, the gearbox, with all accessories and components installed, shall be subjected to a vibratory scan and resonant search while operating throughout the speed range from idle to 125 percent maximum speed under normal pad loads. The endurance test of paragraph 4.6.1.3 shall be run on the gearbox with all accessories and components installed. During the test, all drive pads shall be loaded to at least the maximum overhung moment rating of the drive pads and subjected to at least the maximum spline misalignment allowed.

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Following the test, scavenge oil filter and magnetic plug residue shall be collected and separated into magnetic and nonmagnetic material. This material shall then be weighed, analyzed, and the results reported. At test completion the gearbox shall be completely disassembled, cleaned, and parts inspected. There shall be no evidence of material defects, undue wear, or impending failure. Conformance with the fluid leakage requirements of 3.3.6.4 shall be demonstrated during this test.

4.6.2.4.3 Alternator Test. The following tests, in addition to the environmental tests, shall be conducted on each engine alternator. The alternator shall be subject to pre- and post-test ATPs or functional checks as applicable for tests a and b only:

a. Load Test. The alternator shall be operated at a speed corresponding to 115percent of maximum allowable engine speed of the driving rotor under full rated electrical load for one hour. During the test, the alternator shall be subjected to its maximum component limiting temperature. At completion of the test there shall be no evidence of mechanical or electrical damage or impending failure.

b. Containment. The generator/alternator shall be operated at a speedcorresponding to 122% of the maximum allowable engine steady state speed for a minimum of five minutes without failure to demonstrate design integrity. If the rotor burst capability is less than required above, a containment demonstration will be performed as follows: The alternator shall be operated at a speed corresponding to the maximum allowable engine transient speed of the driving rotor in a manner to cause a mechanical failure of the rotor system. All damage shall be contained within the alternator housing.

4.6.2.4.4 Heat Exchangers. Heat exchangers for cooling or heating of engine fluids or components shall be subjected to the following tests. If the heat exchanger assembly incorporates a bypass valve, regulator, or indicating feature, appropriate tests shall be conducted to demonstrate proper functioning and shall be specified in the pretest data.

a. Flow, Pressure, and Temperature Cycling Test. The exchanger shall be subjectedto a flow, pressure, and temperature cycling test for 15,000 cycles or one design life whichever is greater. The number of cycles shall be specified in the pretest data with a cycle defined as follows:

1. Oil shall be induced into the unit at a flow rate determined from designdata, pressure and oil temperature will also be determined from design data, based on an engine IRP at SLSD condition.

2. Air shall be induced at the same time into the oil cooler at a rate dependenton inlet ambient pressure and at an air temperature commensurate with conditions defined in a(1).

3. The unit shall be allowed to stabilize and the outlet conditions of oil circuitwill be recorded. This outlet condition will be used to determine when the following process is completed.

(a) Oil and air shall be induced into the unit, simultaneously, at thegiven inlet conditions as stated above in this paragraph.

(b) These conditions shall be maintained until the oil outlettemperature is within ±15°C (±27°F) of the stabilized temperature.

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(c) When this point is established the oil circuit will be returned to a zero flow and pressure state. The air circuit shall remain on to speed for the cooling process of the oil cooler.

(d) When the air outlet temperature is within ±15°C (±27°F) of the air inlet temperature, the oil circuit shall be turned on to begin a new cycle. At test completion there shall be no evidence of leakage or permanent deformation.

b. Heat Exchanger, Proof Pressure. Upon successful completion of the flow, pressure, and temperature cycling the same heat exchanger shall be subjected to a proof pressure test. Each fluid side of the heat exchanger shall be individually subjected to twice its maximum working pressure for at least two successive times and held two minutes for each pressure application. During the application of pressure to one side the other element shall be empty and at atmospheric pressure. There shall be no evidence of external leakage or internal leakage into the dry side. Following this test, both sides of the heat exchanger shall be simultaneously subjected to their maximum working pressures for at least two successive times and held two minutes for each pressure application. At test completion there shall be no evidence of leakage or permanent deformation.

4.6.2.4.5 Fire Test. Lines, fittings, and components, including engine furnished oil tanks, which convey flammable fluids shall be tested to verify conformance with 3.3.6.1. Individual lines, fittings, components, or assemblies shall be tested as specified in AC20-135, AS 1055DB and AS 4273A while conveying fluids at the lowest flow rate, highest pressure, and highest fluid temperature possible over the complete operating envelope of the engine. The requirements of 3.3.6.1 shall be considered verified, if during the flame application period (five minutes), and for a period of 5 minutes after removal of the flame, there are no measurable leaks (fire resistant rating). Oil tanks, lines, fittings, and hoses shall have the flame applied for 15 minutes in order to be rated fire proof. For fire proof rated components, the post-flame leakage requirement is not applicable. The above test may be waived for identical components which have successfully completed the test of 4.5.2.5.5.

All fire tests shall be conducted using JP-8 for fuel components. Fuel shutoff capability following the exposure of the fuel shutoff component to the flame shall be demonstrated. It is also acceptable for any shutoff to function automatically at any time during the period when flame is applied, or at any time during the 5-minute post flame period. If the engine control system is located in a designated fire zone, the engine control system shall be tested to verify conformance to 3.3.6.7. During the five-minute flame application period, the engine control system shall continue to control the engine in accordance with the requirements of the engine model specification, or shall cause the engine to fail safe (engine shutdown, or fail fixed if so designed). Additional general guidance may be found in AC 33.17-1A.

4.6.2.4.6 Output Shaft Static Torque Test The output shaft system shall be subjected to testing at the maximum allowable output torque. The shaft shall be heated to simulate its operating temperature within the engine at maximum power test temperature. Torque of not less than the value specified in 3.2.1.4.11 shall be applied for 50 continuous cycles, each cycle consisting of a 15 minute period and a 10 minute period at the maximum allowable output torque (20 hours and 50 minutes total). The torque shall be reduced to zero before each period. At completion of the test, the shaft shall be within

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allowable dimensional limits, torque accuracy shall still be within specification limits and there shall be no evidence of impending failure.

4.6.2.4.7 Overheat Test. Overheat testing shall be conducted on all electronic engine control assemblies, unless the component is mounted in a location where the ambient environment surrounding the component cannot exceed the component limiting temperature as specified in the engine specification, even under failure conditions. If the engine control electronic assembly incorporates hardware or software to initiate hardware shutdown/shutoff, or other defined accommodation, in response to excessive internal temperatures, then this test shall be performed even if the external ambient temperature cannot exceed the defined component limiting temperature under failure conditions.

The engine control system shall be tested to verify conformance to the fail-safe requirements of 3.3.6.7. The engine control system shall be subject to pre- and post-test ATPs or functional checks as applicable, and shall be operating in a ‘closed-loop’ mode wherein the engine/airframe is simulated during the test in order to assess the effects on the engine of the engine control system overtemperature condition. The engine control system need not pass a post-test ATP of functional check. Note that this test is in addition to any fire test, since the failure mechanisms for the two tests are different; one being temperature elevation of the entire component while the other is a potential rapid burn-through of the housing and flame impingement on the circuit card assemblies.

4.6.2.4.8 PMA Neutral Short-to-Ground Test. If applicable, a PMA neutral short-to-ground test shall be performed to demonstrate that the engine control system is neither damaged, nor causes unacceptable engine control system behavior if the neutral of the alternator (not normally brought out to a connector) shorts internally to the alternator case and thereby to engine ground. Such a failure causes excessive current to be drawn through the engine control system’s power supply with the potential for cascading damage.

4.6.2.4.9 Overspeed and Containment. For any engine component containing rotating elements (fuel pumps, hydromechanical metering units, air turbine starter, etc.), the engine contractor shall prepare analyses and conduct component testing to verify that the rotating elements shall not fail at the following overspeed conditions, or that the parts shall be contained if they do fail. The component shall be operated at a speed which corresponds to:

a. 115 percent of the maximum allowable steady state engine speed of the drivingrotor, for 5 minutes, if the component is driven by the gas generator accessory gearbox.

b. 105 percent of the highest speed attained during the engine overspeed controlsystem test of paragraph 4.6.6.9, for 1 minute, if the component is driven by the power turbine accessory gearbox.

c. The highest possible speed attainable by the component under failure conditionsof the supply air, for 1 minute, if the component is air driven. At the completion of the test the component will be disassembled and there shall be no evidence of mechanical damage or impending failure. If impending failure is indicated, a containment demonstration will be performed as follows:

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The component shall be operated at a speed corresponding to the maximum allowable engine transient speed of the driving rotor in a manner to cause a mechanical failure of the rotor system. All damage shall be contained within the component housing.

4.6.2.4.10 Burst Pressure. For any engine component pressurized by fuel, oil or air, it shall be demonstrated that the component can safely meet the requirements during and after exposure to 2.0 times (fuel/oil) or 2.5 times (air) the design maximum operating pressure for a 1-minute period. At the end of the test period, pressure shall be returned to the design maximum operating pressure and held for a period of 5 minutes. This test shall be performed on the same unit that has first undergone the proof pressure test of 4.6.2.4.12. The test pressure shall be applied at the component’s limiting temperature. The test may be conducted at a lower temperature if the applied pressure is increased to compensate. Any such request by the contractor must be supported by analysis justifying the pressure/temperature combination, and shall be submitted to the Using Service for approval. During the test, and for a 5-minute period immediately following, the component shall exhibit no leakage from, or fracture of, the pressurized cavities, during and after a one-minute application of the burst pressure. Seepage or leakage from gasket or other sealing surfaces is allowable.

No post-test ATP or functional check is required since the component shall be considered to be non-serviceable after the test, unless indicated otherwise. The test asset may be usable for other qualification testing depending on its specific condition and the additional testing to be performed. Any further use of the component shall be concurred by the Using Service.

4.6.2.4.11 Pressure Cycling. For any engine component pressurized by fuel, oil or air, it shall be demonstrated that the component can meet the requirements of this plan during and after exposure to a minimum of 15,000 cycles between the minimum and the maximum design operating pressures. The pressure cycling shall be applied at the component’s maximum operating temperature unless the worst case is shown by analysis to be some other combination of pressure and temperature. The test may be conducted at a lower temperature if the applied pressure is increased to compensate. Any such request by the contractor must be supported by analysis justifying the pressure/temperature combination, and shall be submitted to the Using Service for approval. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test unless required in order to perform the test. During and at the completion of the test, the component shall exhibit no leakage from gasket or other sealing surfaces, however minor seepage may be considered allowable if concurred by the Using Service.

Any component requiring testing in accordance with this paragraph shall also require testing IAW 4.6.2.4.12 (proof pressure) and 4.6.2.4.10 (burst pressure). These three tests shall be performed on the same serial number unit, in the following order: 1) pressure cycling, 2) proof pressure, 3) burst pressure. Performance of proof pressure testing in the normal course of product acceptance shall not result in waiver of the proof pressure test requirement.

4.6.2.4.12 Proof Pressure. For any engine component pressurized by fuel, oil or air, it shall be demonstrated that the component can continue to operate in accordance with performance specifications during and after exposure to 1.5 times (fuel/oil) or 2.0 (time (air) the design maximum operating pressure for

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a period of 10 minutes. At the end of the test period, pressure shall be returned to the design maximum operating pressure and held for a period of 5 minutes. This test shall be performed on the same unit that has first undergone the pressure cycling test of 4.6.2.4.11. The test pressure shall be applied at the component’s limiting temperature. The test may be conducted at a lower temperature if the applied pressure is increased to compensate. Any such request by the contractor must be supported by analysis justifying the pressure/temperature combination, and shall be submitted to the Using Service for approval. Components shall be subject to pre- and post-test ATPs or functional checks as applicable, but need not be operating during the test unless required in order to perform the test. During the test, and for a period of 5 minutes immediately following the test, the component shall exhibit no seepage or leakage from gasket or other sealing surfaces, and there shall be no permanent deformation. The component shall be considered to be fully serviceable and may be used for other qualification testing as concurred by the Using Service.

4.6.2.4.13 Pressure Wash Test. The engine (or an engine carcass) with all engine mounted components installed shall be subjected to a pressure wash test in accordance with the general requirements described below. Components shall be subject to pre- and post-test ATPs or functional checks as applicable. Components shall not be operating during the test.

a. The engine inlet and exhaust may be covered.b. Each accessible surface shall be sprayed from a distance of 18 inches to 24

inches using a fan pattern nozzle, with the fan making an acute angle of between 15° and 45° to the surface being sprayed.

c. Each surface shall be swept with the fan spray a total of 8 times; twice eachcoming from the left, right, top and bottom directions as viewed when face on to the surface.

d. Pressure shall be maintained between 2500 psi and 3000 psi throughout the test.No special requirements are levied for the water spray solution.

e. At the completion of the test, without removing the component from the engine, itshall be dried using lint free towels so that there is no dripping or standing water visible on any surface.

f. Any covers or other component features which lead to air cavities shall beremoved and the interior cavity examined for evidence of water ingress. If any water is found inside a cavity, the quantity shall be estimated.

g. The engine and components shall be disassembled to the extent necessary todetermine water ingress. Oil samples shall be taken from the sumps to determine water contamination and oil quality. There shall be no water ingress that could result in an immediate impact on engine or component operation or a long term effect on operation and/or reliability that could result in an early engine or component removal from service. This test may be conducted on individual components, or groups of components, provided that the test set up creates a reasonable facsimile of the actual installation of the component on the engine. However, testing of components individually or in groups does not waive the requirement to conduct this test on the engine as a whole.

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4.6.2.4.14 Ignition System Fouling. The following tests, in addition to the other environmental tests, are applicable to all ignition systems in order to demonstrate that the ignition system will be capable of consistently starting the engine under fouling conditions. The power input for the test shall be the minimum defined by 4.6.2.2.2.1.

4.6.2.4.14.1 Carbon Fouling. The spark igniter of the ignition system test assembly shall demonstrate sparking performance per 3.7.5.3.1 with spark gaps covered, filled, or bridged with a generous application of an amorphous carbon/oil mixture. Sparking rate shall be measured during the test, while energy delivery shall be measured both before and after the test.

4.6.2.4.14.2 Water Fouling. With the spark igniters positioned in a manner simulating the mounted position on the engine, the minimum power input specified in this AQP shall be supplied. The spark igniters shall meet the requirements of 3.7.5.3.2.Sparking rate shall be measured during the test, while energy delivery shall be measured both before and after the test.

4.6.2.4.15 Electrical Loads Analysis The contractor shall prepare an analysis, supported by test data that establishes the worst-case power consumption for the component, if it utilizes airframe power as either a primary or backup power source. The analysis shall verify that the component maximum current draw does not exceed the circuit breaker capacity under the worst case combination of loads, ambient temperature, and failure of one power source, if redundant. Supporting data shall include a Load Measurements test performed under the guidelines of MIL-HDBK-704-8, LDC101.

4.6.2.4.16 Short Circuit Protection For any short circuit condition specified in subparagraphs a through c, the condition shall only result in temporary failure of the shorted interface. Subsequent to the removal of the short circuit condition, the component shall be undamaged and fully functional.

a. Short circuit between any two adjacent pins within the component connector.b. Short circuit between any component connector pin and ground.c. Short circuit between any two pins (including non adjacent) of any input sensor,

output effector, or other external input or output device or subsystem which interfaces with the component.

Additionally, a short of any pin to 28 Vdc shall not result in the propagation of a permanent failure in the component beyond the interface or circuit affected by the short circuit condition.

4.6.2.4.17 Data Bus Specification Compliance Electronic units which incorporate data bus communication interfaces shall be verified for compliance with protocol specifications.

a. MIL-STD-1553B. Data bus remote terminal functionality shall be tested inaccordance with SAE AS4111.

b. ARINC 429. Data bus hardware characteristics shall be tested in accordancewith ARINC specification 429, Part 1, Appendix A.

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4.6.2.5 Software Verification. In order to receive a qualification test (QT) rating, the control system shall demonstrate it meets the requirements of 3.7.2 and all subparagraphs by passing a formal qualification (FQT) in accordance with Appendix G of this AQP.

4.6.2.5.1 Engine Testing (Software FQT) The contractor may elect to use data gathered from engine testing as part of software verification and/or system validation. However, for process repeatability all such testing shall be fully documented with respect to test case description, expected outcome, pass/fail criteria, and traceability to specific requirements. Additionally, implicit with the use of such means is the need to repeat the engine testing when a change affects those system or software requirements tracing to these particular test cases. This may become overly burdensome, hence relying on engine testing for software verification or system validation shall be minimized.

4.6.2.5.2 Flight Testing (Software FQT) Flight testing is the only feasible method to validate certain features, functions or performance of the control system software, or finalize airframe/engine integration parameters within the control. When this is the case, for process repeatability all such testing shall be fully documented with respect to test case description, expected outcome, pass/fail criteria, and traceability to specific system-level requirements. Since flight testing will have to be repeated when a change affects those system or software requirements tracing to these particular flight test cases, it shall be only utilized when other means are not possible.

4.6.2.6 Complex Power Interrupt Test (Closed Loop). The contractor shall conduct a comprehensive test of engine control system response to DC power interrupts when operating in a closed loop bench test environment. The Using Service will provide guidance in preparing the necessary test cases based on the particular system architecture and defined power sources. It shall be demonstrated that the engine control system can tolerate and safely recover from single, multiple and asymmetric interrupts of varying duration when:

a. Applied to each channel and both channels, with channels in both a healthy anda failed condition.

b. Applied during the engine start regime, prior to alternator power being available.Note that the above requirement is in addition to power interrupt testing conducted as part of MIL-STD-704 power quality verification. Effects of power interrupts must be verified in a closed-loop environment, since it is not possible to ascertain if the engine control system will behave properly, particularly since engine control system power up logic often depends on whether or not the engine is running at the time of the interrupt. Furthermore, channel swapping algorithms and power-up BIT strategies are very complex.

4.6.2.7 EMS Lifing Algorithms. Tests and analyses shall be conducted in accordance with the VV&A plan outlined in 3.7.6.7.1 to show that the lifing algorithms calculate accurate usage and remaining life for life limited parts during actual engine operation. EMS Lifing Algorithms shall be actively used during the 150 hour EIT endurance test (4.5.1), 300 hour QT endurance test (4.6.1), the LCF engine test (4.6.6.2.2), the 600 hour AET (4.8.1) and the AMT (4.8.2), to demonstrate algorithm performance. At the conclusion of each test above, the EMS calculated usage and remaining life must be within 10%

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of the pretest predictions calculated from a finite element analyses for all of the components being tracked.

4.6.2.8 Common Mode Multiple Signal Failure Test. The contractor shall demonstrate that the engine control system can safely accommodate simultaneously occurring, multiple signal failures due to a common cause (a connector backing off or ballistic damage to a harness branch).

The contractor shall demonstrate by test that:

a. For any single connector disconnecting, acceptable engine control system performance and/or behavior is maintained.

b. For any electrical harness branch sever, acceptable engine control system performance and/or behavior is maintained.

c. Each test case defined in both a and b, above, shall be executed at max continuous power.

d. Each test case defined in a, above, shall be conducted at an engine speed between light-off and ground idle (approximately 25% to 35%).

The test shall be conducted in a closed-loop bench test environment or on an actual engine, or a combination of both. If the signal group for a specific test case is expected to have a more pronounced effect at an engine operating condition other than those specified above, these may be tested in place of the specified conditions.

For test purposes, the actual physical connector can be disconnected, or the group of signals associated with that single connector or branch opened simultaneously (defined as all pins open within 50 msec maximum). If a connector or branch contains only one signal or parameter, then it need not be tested. This test requirement applies to all electrical connectors and/or harness branches within the engine control system, not just those for the electronic control unit.

4.6.3 Altitude Tests. An engine, conforming to the same parts list and configuration as the endurance test engine, shall be subjected to altitude tests which shall consist of operation and air starting checks at selected conditions throughout the operating envelope specified for the engine and at least those given in the engine specification per Figure 19. The test points shall include the effects of power extraction, bleed air extraction, windmilling, anti-ice operation, inlet recovery and inlet distortion on engine performance and stability. Pretest data shall be per 4.3.5.1. The effects of anti-ice operation shall be included in the demonstration of 4.6.3.2 sections a., b., and c. Adjustments to the engine shall not be made without approval of the Government representative. The altitude tests shall be accomplished using the oil and fuel combinations of: (1) MIL-PRF-23699 oil and MIL-DTL-83133H, grade JP-8 fuel and (2) MIL-PRF-7808 oil and MIL-DTL-5624U, grade JP-4 fuel. Fuel temperature shall be varied over a range sufficient to encompass all anticipated engine operating environments. All power take-off pads shall be loaded as specified by the Using Service for a particular test phase. Data to be taken and recorded during the test shall be as specified in Table XII. For this engine test, the test cell shall conform to the power absorption device requirements of 4.3.3.3.2.

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Overall true RMS velocity measurements and acceleration spectrograms shall be obtained for each sensor mounted on the engine case and accessory gearbox case at engine speeds and powers selected for the test. The points selected shall include at least the altitude rating points and the point in the operating envelope where the highest engine vibration levels are generated. Critical components of the engine shall be identified on each spectrogram.

4.6.3.1 Altitude Calibration. Prior to initiation of the testing described in 4.6.3.2, the engine shall be calibrated IAW 4.6.1.2.2 and 4.6.1.2.3. No control readjustments shall be made after the initial adjustments at the beginning of the calibration. In addition, calibration shall be conducted on control system components IAW 4.6.1.2.1.

4.6.3.2 Altitude Test Procedure. Operation at each test point shall be of sufficient duration to stabilize the engine and to establish the performance and operating characteristics of the engine. If the control system has defined ‘failure’ modes of operation (e.g., fail fixed, degraded capability, manual backup, etc.), each mode shall be declared and demonstrated. Engine operation in these failure modes shall be evaluated and the effects on engine performance and operating characteristics shall be determined during the test. The control failure modes to be evaluated shall be specified in the pretest data and shall be subject to approval by the Using Service. Operation shall be conducted to obtain the following data:

a. Altitude Rating Points. The test points shall be those specified for altitude ratings in Table IIA, IIB, IIC and IID of the engine specification. A sufficient number of additional engine power settings shall be selected for each specified altitude test condition to establish operating and performance characteristics at each rated point. The time elapsed versus engine speed, measured temperature, and fuel flow shall be obtained for stability verification with the power setting at idle, maximum continuous, intermediate, and maximum. The time period for stability verification shall be a minimum of five minutes at each power setting.

b. Transient Operation. The applicable transient performance specified in 3.2.1.5.6 shall be demonstrated at each rating condition.

c. Functional Test. The operating envelope of the engine shall be verified by running the engine at the extremities of the operating envelope. Engine steady state and transient characteristics shall be determined at each test point over the range of power settings.

d. Inlet Distortion. Stable engine operation and performance effects of inlet airflow distortion at the limits defined in Appendix (contractor to specify) shall meet the requirements of 3.1.2.10.3 and be demonstrated at three flight conditions: 1) sea level, static, standard day, 2) 10000 ft, 0.4 Mach number (Mn), hot day, and 3) 20000 ft, 0.2 Mn, standard day and shall meet the requirements of 3.1.2.10.4 and be demonstrated at sea level, static, standard day conditions. These steady-state demonstrations shall be performed at the engine ratings defined in Table IA. Stable compressor operation during engine transients, shall be demonstrated for the Qualification Test Distortion Patterns defined in Appendix (contractor to specify), Figure (contractor to specify).

e. Starts and Restarts. Flameouts with the ignition system not operating shall be accomplished by means specified in paragraph 3.2.1.5.3 of this AQP. Engine air starts and restarts shall be accomplished at each of the specified air starting points shown on Figure 19 without power extraction.

f. Altitude Windmilling Test. Altitude windmilling tests shall be conducted within thewindmilling envelope to verify the requirements of 3.2.1.5.7. In addition, testing shall be accomplished to verify that the lubricating system will provide proper lubrication, as defined in

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the pretest data, and operate without excessive loss of oil during windmilling operation. Engine shafts may be driven by any suitable external means.

4.6.3.3 Altitude Test Completion. Comparison of observed data obtained during the test to the specified performance and operating characteristics shall be made by a method acceptable to the Using Service. The test shall be considered to be satisfactorily completed when, in the judgment of the Using Service:

a. The engine performance is at least that specified in the steady state performance computer program and the pressure and temperature lapse rates for MGT, SHP, SFC, and Ng are within ±2.0 percent of the those obtained from the steady state performance computer program.

b. The altitude starting and transients conducted during the test are in accordance with the requirements specified in 3.7.9.2 and 3.2.1.5.6.1 respectively.

c. The functional test points demonstrate satisfactory engine operation and do not show any discrepancies with the steady state performance computer program, altitude starting data, or transient data.

4.6.4 Engine Environmental and Ingestion Tests. The tests in the following subparagraphs shall be conducted on engines having the same parts list and configuration as the endurance test engine of 4.6.1. Unless otherwise specified in the individual test, the engine shall be calibrated before and recalibrated after each test, but only to the extent necessary to determine any deterioration in steady state or transient performance capability which occurred during the course of testing. All starts shall be performed with a starter in accordance with 4.3.3.4. Unless otherwise specified for a particular test, the test shall be conducted at the ambient conditions which prevail at the test site.

4.6.4.1 Low and High Temperature Starting and Acceleration Test. The test engine shall be subjected to low and high temperature tests to demonstrate compliance with 3.2.1.4.8, 3.2.1.5.6, 3.2.5.1, 3.7.2.2.2, 3.7.9.2, and 3.7.9.3. All data required in 4.3.5.6 shall be recorded during each start. Starting and operating capabilities shall be accomplished with the torque loading as specified in Figure 17-1, applied to the accessory drive pads. Recalibration shall not be required. The point within the engine where temperature shall be measured for determining the engine soak temperature shall be specified herein.

a. Low Temperature Test. The engine, serviced with the oil specified in this AQP shall be subjected to a soaking period of at least ten hours duration at an ambient temperature of -54°C (-65°F). The ten hour soaking period shall be started after the point specified in this AQP has reached -54°C (-65°F). At the end of the low temperature soaking period, the electrical connectors shall be disconnected and reconnected to verify the requirements of 3.7.4.3.2. In addition, the oil reservoir filler cap and all other servicing features shall be checked to demonstrate their proper functioning under cold soak conditions. After the soak period and when supplied with fuel and inlet air at a temperature of -54°C (-65°F), a start shall be made with a control power signal request to idle. Immediately (within 30 seconds) after the engine has reached idle speed, an acceleration to maximum shall be accomplished by a control power signal request. The engine shall then be returned to idle and shutdown. The above procedure, including the soak period, shall be repeated twice. The complete test above shall be repeated using each of the primary and restricted fuels and oils. The soak, oil, air, and fuel temperatures shall be subject to fuel or oil temperature limitations of 3.2.5.1, 3.7.3.2 and 3.7.7.2.1 respectively.

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The test will be considered to be satisfactorily completed when, in the judgment of the Using Service, the above three successive starts on each fuel/oil combination have been satisfactorily accomplished within the time limits specified in Figure 18; the engine has demonstrated its ability to accelerate to maximum without exceeding any engine starting or operating limits; there were no fuel or oil leaks; and functional checks of electrical connectors and servicing features have revealed no damage or difficulties during operation.

b. High Temperature Test. The engine, serviced with the oil specified in the engine specification shall be subjected to a soaking period of at least ten hours duration at an ambient temperature of 57°C (135°F). The ten hour soaking period shall be started after the point specified in the engine specification has reached 57°C (135°F). After the soak period and when supplied with fuel at a temperature of 57°C (135°F) and air at 57°C (135°F), a start shall be made with a control power signal request to idle. Immediately (within 30 seconds) after the engine has reached idle speed, an acceleration to maximum shall be accomplished by a control power signal request. The engine shall then be returned to idle and shutdown. The above procedure, including the soak period shall be repeated. The complete test above shall be repeated using each of the primary and restricted fuels and oils. The soak, oil, air, and fuel temperatures shall be subject to fuel or oil temperature limitations of 3.2.5.1, 3.7.3.2 and 3.7.7.2.1 respectively. The test will be considered to be satisfactorily completed when, in the judgment of the Using Service, the above two successive starts on each fuel have been satisfactorily accomplished within the time limits specified in Figure 18; the engine has demonstrated its ability to accelerate to maximum without exceeding any engine starting or operating limits, and there were no fuel or oil leaks.

4.6.4.2 Environmental Icing Test. The engine shall be subjected to an environmental icing test to demonstrate compliance with 3.2.5.2. For this test, the engine shall be operated under the free air conditions listed in Table XIII. For each test run, the liquid water content and droplet size shall be measured at a distance within 5.0 ft of the engine inlet face and still within the engine inlet duct. The liquid water content measured at this station shall correct to the free air conditions as specified in Table XIII. This meteorological data shall be recorded at suitable intervals during each test run. The method and procedure for collecting and determining the water droplet size and liquid content shall be specified in the pretest data. During the test, torque, speed, MGT, and vibration shall be continuously recorded and video coverage of the engine inlet shall be provided. It shall be demonstrated that ice from the portion of the test facility not under test does not interfere with the test results. The base line for determining engine performance loss shall be established by operating the engine with no power extraction or customer bleed and under the inlet temperature conditions of Table XIII with air between 80 and 100 percent relative humidity and zero liquid water content. The output shaft power and specific fuel consumption losses shall be determined by comparison of engine performance when operating at the icing conditions defined in Table XIII with the aforementioned base line values.For engines that use oil as the anti-icing fluid, the oil temperature shall be maintained at the minimum operating oil temperature specified in the engine specification, or less, during all runs except at idle condition where the oil shall be maintained at the same temperature as the engine inlet air. If the engine anti-icing system uses compressor bleed air as an energy source and allows customer bleed air extraction, the icing test shall be conducted with maximum allowable customer bleed air extraction.

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The icing tests (part a. and part b.) shall be conducted with a test inlet (bellmouth) that is representative of the installed engine. The test inlet (bellmouth) must be anti-iced, but in a manner that allows no more than a 14°C (25.2°F) surface temperature rise in the engine inlet at a distance 1 inch downstream from the engine aerodynamic inlet plane. The test shall consist of two parts as follows:

a. This part shall consist of two runs at each of several engine power settings under each of the conditions in Part 1 of Table XIII. The engine power settings shall include: idle, no load, 25 percent maximum continuous, 50 percent maximum continuous, 75 percent maximum continuous, maximum continuous, intermediate, and maximum. At each icing condition and at each power setting, the engine shall be operated for a period of not less than ten minutes. During each period, at intervals after ice buildup, the engine shall be rapidly accelerated to maximum to demonstrate acceleration response. Prior to the start to idle at -5°C (23 oF), the engine shall be cold soaked until oil temperature throughout the oil tank equals inlet air temperature.

b. This part shall consist of a one hour run at idle with no throttle movement, followed by an acceleration to maximum at the end of the period. During this run the engine shall be operated under the conditions listed in Part 2 of Table XIII. Prior to the start to idle at -5°C (23 oF), the engine shall be cold soaked until oil temperature throughout the oil tank equals inlet air temperature. At each of the test power demand settings above, power shall be measured to verify performance loss, without any power lever or control power signal request movement. If the engine incorporates an anti-icing system, the above tests shall be performed using the anti-icing system to demonstrate the requirements of 3.2.5.2 and 3.7.1. The testing will be considered satisfactorily completed when, in the judgment of the Using Service, there is no damage to the engine and performance is within the requirements of 3.2.5.2 and 3.7.1.

4.6.4.3 Corrosion Susceptibility Test. A new or newly overhauled test engine shall be subjected to a corrosion susceptibility test to demonstrate compliance with 3.2.5.5 and operated in accordance with Table XIV and Figure 23. The engine will be subjected to two hours of sand ingestion testing in accordance with 4.6.4.7.2 prior to the start of this corrosion susceptibility test. Prior to starting the corrosion susceptibility test, the engine shall be disassembled sufficiently to inspect the surface condition of all parts normally exposed to atmospheric conditions. Detailed photographic coverage of these parts shall be taken. Upon reassembly and after an initial calibration, including data required to establish compliance with transients of 3.2.1.5.6, the engine shall be subjected to the cycles of 3.2.5.5 for 24 hours each in accordance with Table XIV. Anti-icing and customer bleed air shall be extracted for ten minutes of the Table XIV three-hour phase 1 operating cycle. Should engine performance during the test deteriorate more than the amount specified in 3.2.5.5 from that determined during the initial calibration, water washing shall be accomplished in accordance with 3.7.13. Water wash shall be used only to correct for performance loss. External water wash is not permitted. If performance cannot be recovered after water washing, the engine shall be disassembled and inspected to determine the effect of the corrosion testing on performance loss. If the unrecoverable performance loss is determined to be caused by a problem not related to corrosion, the engine shall be repaired and reassembled and the test continued. During the test, the engine shall be subjected to internal inspections after every tenth cycle to detect any evidence of corrosion or progression of corrosion of internal parts. Additional inspection may be conducted with approval of the Using Service. After completion of the corrosion susceptibility test, the cleaning procedure of 3.7.13 shall be accomplished prior to recalibration of the engine. During recalibration the applicable transient

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performance specified in 3.2.1.5.6 shall be demonstrated. Following recalibration, the engine shall be disassembled and inspected for evidence of corrosion. Detailed photographs shall be taken of all parts which show evidence of corrosion. Metallurgical analyses that completely characterize the types of corrosion found and test specimen evidence shall be available for review. The corrosion susceptibility test will be considered to be satisfactorily completed when upon recalibration:

1. The engine non-recoverable performance deterioration at steady state measured temperatures at maximum continuous and above does not exceed the deterioration specified in 3.2.5.5, for the same value of measured temperature obtained during the initial calibration,

2. The engine exhibits not more than the increase in specific fuel consumption specified in 3.2.5.5 for the same values of measured temperature obtained during the initial calibration at idle and above,

3. The transients conducted are in accordance with the engine specification requirements, and

4. The extent of corrosion evident at test completion satisfies the following criteria for all engine parts:

(a) All internal parts exposed to gas path air, upon cleaning, shall show no impairment of their function due to corrosion. Minor corrosive attack is acceptable only when the part design criteria (e.g., fatigue resistance) are not affected.

(b) All corrosion-protected parts shall indicate no effects of corrosion upon cleaning and stripping of the protective schemes. Recoating of parts designed for recoating shall restore such parts to a fully functional condition.

(c) All other parts shall show no corrosion that affects component integrity or prescribed maintenance procedures.

4.6.4.4 Bird Ingestion Test. The engine shall be subjected to a bird ingestion test conducted to verify the requirements of 3.2.5.6.1. Selection of test birds shall be conducted in a manner such that the first weighed birds that fall between 3.0 and 3.5 ounces shall be selected for test. Further, the test birds shall be randomly selected for the part a., part b., and part c. bird shots. The birds shall be ingested in a random sequence and dispersed over the inlet area, to simulate an encounter with a flock. No synthetic birds may be used for testing. The pretest data shall specify the critical target areas for each bird size and the procedure to be used for ingestion. High speed photographic coverage of the inlet is required during the ingestion test. The test will be considered to be satisfactorily completed when the performance criteria of 3.2.5.6.1 have been met, there is no evidence of major structural damage that could cause the engine to fail, and any questions or issues have been resolved that have arisen during the conduct of the test or the posttest hardware inspections so that the Using Service can verify specification compliance, approve the subject test report, and/or otherwise qualify the engine, assembly, or component.

4.6.4.5 Foreign Object Damage Test. Components shall be subjected to a foreign object damage test to demonstrate compliance with 3.2.5.6.2. Subject to the approval of the Using Service, the foreign object damage test shall be conducted by bench testing on individual blades or stators or rig testing on full-scale compressor. The component test plan shall be presented to and approved by the Using Service prior to the component test. Simulated foreign object damage shall be applied to three of the most critical

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blades or stators at the most critical leading edge location. The damage applied shall produce at least a fatigue notch factor (Kf) of 3.0. Following the foreign object damage application, the damaged blades shall be tested to the requirements in 3.2.5.6.2 using a stair-step cycle to accumulate 107 cycles at primary resonant frequencies for this blade or stator. Sufficient instrumentation for monitoring the structure of the components shall be included in the test hardware. At the completion of the test there shall be no evidence of blade or stator failure or cracking as the result of foreign object damage.

4.6.4.6 Ice Ingestion Test. The test engine shall be subjected to an ice ingestion test to demonstrate compliance with 3.2.5.6.3. Hailstones shall be ingested at typical takeoff (maximum power), cruise, and descent conditions. Sheet ice shall be ingested at typical takeoff and cruise conditions. The temperature of the ice shall be between -2°C - 0°C (28°F - 32°F). The test procedure for sheet ice ingestion shall require the most severe ice velocities representing ice shedding off the aircraft inlet lip. The pretest data shall specify the procedures to be used for introduction of ice at the engine inlet and the engine power settings and speed at which the ice or hailstones are to be ingested. The test procedure shall require the engine to run for at least five minutes following ice ingestion, before it is shut down for inspection. The time for engine power recovery shall be recorded. During the tests, high speed photographic coverage of the inlet is required. Sufficient instrumentation for monitoring the structure of the engine shall be included in the test engine. The test will be considered to be completed when, in the judgment of the Using Service, the criteria of 3.2.5.6.3 have been met.

4.6.4.7 Sand Ingestion Tests.

4.6.4.7.1 Fine Sand Ingestion Test. An engine shall be tested with 0 to 80 micron sand and dust ingested at the concentration levels specified in 3.2.5.6.4.1. The engine shall be tested for 9 hours at maximum, 27 hours at intermediate, and 18 hours at maximum continuous, for a total of 54 hours with not less than 27 starts. The test cycle shall be 10 minutes at maximum, 20 minutes at maximum continuous and 30 minutes at intermediate. All ratings shall be initially set to measured gas temperature associated with rated rotor inlet temperature. After test initiation, the engine shall be run at a constant gas generator speed for each power condition unless the measured gas temperature associated with the first stage turbine rotor inlet temperature deterioration limit is reached in which case the measured gas temperature will be held constant. This test shall be terminated and the engine shall be removed and disassembled if power deterioration based on measured gas temperature exceeds 15 percent or if engine failure or impending failure is evident. The engine shall be tested with the IPS if it is an integral part of the engine design. During each hour of operation, at least two decelerations to idle and acceleration to maximum continuous speed shall be made, with power lever movements or control power signal request of 0.5 seconds or less. A calibration test shall be performed at the beginning, at 25 hours and end of the test run. The engine shall be inspected at 25 hours by borescope and other visual techniques that do not require disassembly of the engine. Upon completion of the test run the engine shall be disassembled in such a way that the contamination displacement is minimized. The engine shall be disassembled to determine the extent of sand erosion, and the degree to which sand may have entered critical areas in the engine. Each major rotating and stationary component subject to the effects of sand ingestion shall be weighed at engine build, disassembly and after cleaning. The test shall be considered satisfactorily completed when 54 hours of testing have been

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completed, the engine is able to conduct the accelerations of 3.2.1.5.6.1 without surge or stall, the performance criteria of 3.2.5.6.4.1 have been met and the teardown inspection reveals no failure or evidence of impending failure in the judgment of the Using Service.

4.6.4.7.2 Coarse Sand Ingestion Test. The test engine shall be subjected to a run of 50 hours duration at maximum continuous rated measured gas temperature with 0-1000 micron sand contaminant in accordance with 3.2.5.6.4.2 introduced into the engine inlet. The engine shall be tested with the IPS if it is an integral part of the engine design. During each hour of operation, at least one deceleration to idle and acceleration to maximum continuous rated measured gas temperature shall be made with control power signal requests. Ten periods of one minute operations of the anti-icing system shall be performed during the first hour of each five hour cycle. The engine shall be shut down and cooled at least 12 hours following each five hours of sand ingestion. The engine internal washing provisions shall be demonstrated once during each shut down. During the entire test, maximum customer bleed air shall be extracted from the engine. The bleed air shall be continually filtered, the total deposits measured and results reported. Following the post test performance check, the engine shall be disassembled as necessary to inspect for the extent of sand erosion and the degree to which sand may have entered critical areas in the engine’s internal air cooling system. The air-oil cooler inlet shall be considered an engine inlet for the purpose of this test. However, a separate rig test may be run to satisfy the sand ingestion requirements of the air-oil cooler for this paragraph. The test will be considered to be satisfactorily completed when, in the judgment of the Using Service, the performance criteria of 3.2.5.6.4.2 have been met and teardown inspection reveals no failure or evidence of impending failure.

4.6.4.8 Atmospheric Water Ingestion Test. The test engine shall be subjected to a water ingestion test to demonstrate compliance with 3.2.5.6.5. With the engine operating at maximum power, 2.0, 3.5, and 5.0 percent of the total airflow mass in the form of liquid water shall be introduced into the inlet of the engine with 50 percent of the liquid water entering the engine inlet through a segment equivalent to one third the inlet area. Water ingested at rates of 2 to 3 percent of the total airflow mass shall be accomplished with no airbleed, while water ingested at rates of 3.5 to 5 percent of the total airflow mass shall be accomplished with air bleed. The engine shall be operated at each water ingestion condition for five minutes. Idle-max-idle throttle transients shall be conducted at each test condition. The above procedure shall be repeated with the engine operating at idle. During the test, the effects of the water ingestion on engine performance shall be noted and recorded. At test completion, the engine shall be shutdown and allowed to cool to ambient temperature before making the posttest performance check. Following the performance check, the engine shall be disassembled sufficiently for inspection. This test shall be considered to be satisfactorily completed when, in the judgment of the Using Service, adequate clearances were maintained, no damaging rub, or detrimental rubbing occurred during the test, the post-test performance has not deteriorated from the pre-test performance calibration and the gas path parts show no damage.

4.6.4.9 Engine Component Limiting Temperature Test. The engine shall be subjected to an engine component limiting temperature test to demonstrate compliance with 3.1.2.8.1. The test shall be conducted as follows:

a. The engine shall be run in an uninstalled configuration. Uninstalled is defined asat least five feet of airspace on all sides of the engine.

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b. During the test, ambient air shall be at the maximum stagnation air temperature of 55°C (131°F). Although the test is run in still air, forced convection shall be allowed as required to exhaust the heat rejected by the engine in order to maintain a constant 55°C (131°F).

c. The engine shall be run at the intermediate (Table IC) engine rating. This condition shall be stabilized for 30 minutes or until engine oil temperature is increasing at a rate of less than 1°C (1.8°F) in 15 minutes, whichever is longer. Engine oil temperature shall be measured by the engine supplied oil temperature sensor.

d. Once the operating condition is stabilized, the power shall be increased to the maximum (Table IC) engine rating for ten minutes and all surface temperature data shall be recorded for all parts under test.

e. Immediately upon recording the data specified in part d the engine shall be shutdown. Immediately upon commanding a shutdown, any exhaust flow used to maintain the maximum stagnation temperature shall be shutoff. All electrical power shall be removed from the engine once gas generator speed is less than 8% with the exception that any components powered directly from the airframe electrical busses shall remain powered until completion of part f.

f. Once gas generator speed is less than 8%, surface temperature data shall be recorded once every 30 seconds for a period of 15 minutes.

4.6.4.10 Noise Survey. The survey shall be conducted to substantiate the requirements of 3.2.5.7. The engine shall be mounted on an outdoor test stand with a minimum clearance of seven feet between the lowest part of the engine and the ground. Microphones shall be located in relatively flat terrain, free of excessive ground absorption characteristics. There shall be no obstructions that significantly influence the engine noise field. The weather shall be free of precipitation with relative humidity between 20 and 95 percent, ambient temperature between -10°C (14°F) and 35°C (95°F), wind velocity less 6 miles per hour and no temperature inversions or anomalous wind conditions. Acceptable atmospheric absorption corrections to standard acoustic day conditions of 25°C (77°F) and 70 percent relative humidity shall be made. No post-test engine recalibration shall be required. The signal level shall be at least 10 dB greater than the background noise level in each third octave band in the frequency range of interest.

a. Data presentation. The engine noise level data shall be presented as follows andin accordance with the format shown on the referenced figures:

1. Equal overall sound pressure level contours (dB reference of 0.0002microbars) as shown on figures 30-1 and 30-2.

2. One-third octave frequency spectra from a center frequency of 20 Hz to 10kHz at the positions A through G specified on Figure 30-1 shall be tabulated. All near field sound pressure levels shall be based on measurements at head level where personnel are located during engine maintenance or other operation. In addition, one-third octave band sound pressure levels from a center frequency of 20 Hz to 10 kHz in the form of uncorrected data (e.g., uncorrected for weather, terrain, etc.) shall be tabulated for all Table IA, IB, IC, IIA, and IIB performance points and for 19 positions. The 19 positions shall be based on measurements located 100 feet radially from the center of the engine’s exhaust plane at ten degree increments starting at zero degrees directly in front of the engine and ending at 180 degrees directly behind the engine.

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3. Equal sound pressure level contours plotted for octave bands with center frequencies of 250 Hz, 500 Hz, and 1000 Hz shall be shown on Figure 30-1 for the maximum and maximum continuous power of Table IA, IB, IC, IIA, and IIB only.

4. Equal perceived noise level (tone-corrected) contours in 5 PNdB incrementsshall be shown on Figure 30-2 for all Table IA, IB, IC, IIA, and IIB performance points.

5. Three narrow band spectrum plots from 20 Hz to 10 kHz for the inlet and exhaust noise shall be shown in figures. These narrow band spectrum plots shall be presented in a spectral density format (e.g., sound pressure spectrum level) which normalizes the data to account for changing analysis bandwidth. (Position and power setting for the narrow band data shall be determined for an examination of the third-octave spectra).

6. The total acoustic power (dB ref 10-12 watts) generated at the maximum power condition.

7. There shall be sufficient data in the report to enable prediction of the engine noise signatures at any power setting. The required data shall include rotor speed, number of blades, hub to tip ratio, diameter, discharge total temperature, pressure ratio, exit velocity, exhaust mass flow, and ambient temperature and pressure.

b. Measurement system. The acoustical measurement system shall consist of approved equipment equivalent to the following:

1. A microphone system with frequency response compatible with measurement and analysis system accuracy as stated in c. below.

2. Tripods or similar microphone mountings that minimize interference with the sound being measured.

3. Recording and reproducing equipment with characteristics, frequency response, and dynamic range compatible with the response and accuracy requirements of c. below.

4. Acoustic calibrators using sine wave or broadband of known sound pressure level. If broadband noise is used, the signal shall be described in terms of its average and maximum RMS value for a non-overload signal level.

5. Analysis equipment with the response and accuracy requirements of d. below.c. Sensing, recording and reproducing equipment. The sound produced by the

engine shall be recorded in such a way that the complete data history is retained. A magnetic tape recorder is acceptable.

1. The characteristics of the system must comply with recommendations givenin International Electromechanical Commission (IEC) 179 with regard to the microphone and amplifier characteristics.

2. The response of the complete system to a sensibly plane progressivesinusoidal wave of constant amplitude must lie within the tolerance limits specified in IEC 1672-2 over the frequency range of interest.

3. The equipment must be acoustically calibrated using facilities for acoustic freefield calibration and electronically calibrated as stated in paragraph d. below.

d. Analysis equipment. A frequency analysis of the acoustical signal shall beperformed using one-third octave filters complying with the recommendations given in IEC 225.

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1. The analyzer-indicating device must be analog, digital, or a combination of both. The preferred sequence of signal processing is:

(a) Squaring the one-third octave filter outputs.(b) Averaging of integrating and linear to logarithmic conversion.

2. The indicating device must have a minimum crest factor capacity of threeand shall measure, within a tolerance of +1.0 dB, the true root-mean-square (rms) level of the signal in each one-third octave band. If other than a true rms device is utilized, it must be calibrated for nonsinusoidal signals. The calibration must provide means for converting the output levels to true rms values.

3. The amplitude resolution of the analyzer must be at least 0.25 dB.4. Each output level from the analyzer must be accurate within +/-1.0 dB with

respect to the input signal, after all systematic errors have been eliminated. The total systematic errors for each of the output levels must not exceed +/-3 dB. For continuous filter systems, the systematic correction between adjacent one-third octave channels may not exceed 4 dB.

5. The dynamic range of the analyzer must be at least 55 dB in terms of thedifference between full-scale output level and the maximum noise level of the analyzer equipment.

6. The complete electronic system must be subjected to a frequency andamplitude electrical calibration by the use of sinusoidal or broadband signals at frequencies covering the range of interest of known amplitudes covering the range of signal level furnished by the microphone. If broadband signals are used, they must be described in terms of their average and maximum rms values for a non-overload signal level.

7. Narrowband analysis shall be conducted with a maximum bandwidth of 5 Hzin the 20 Hz to 500 Hz frequency range, 20 Hz in the 500 Hz to 5000 Hz frequency range, and 100 Hz in the 5000 Hz to 10000 Hz frequency range.

e. Noise measurement procedures. In order to ensure uniform practices relative toacoustic testing of engines, the following procedures shall be required:

1. Immediately prior to and after each test, a recorded acoustic calibration of thesystem shall be made in the field with an acoustic calibrator to check system sensitivity and provide an acoustic reference level for the analysis of the sound level data.

2. For the purpose of minimizing equipment or operator error, field calibrationsshall be supplemented with the use of an insert voltage device to place a known signal at the input of the microphone, just prior to and after recording engine acoustic test data.

3. The ambient noise, including both acoustical background and electrical noiseof the measurement system, shall be recorded and determined in the test area with the system gain set at levels which will be used for aircraft engine noise measurements.

f. Reporting and correcting measured data. Data representing physicalmeasurements or corrections to measured data shall be recorded in permanent form and included in the final report. Estimates must be made of the individual errors inherent in each of the operations employed in obtaining the final data.

1. Measured and corrected sound pressure levels must be presented in one-thirdoctave band levels obtained with equipment conforming to the standards described in previous paragraphs.

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2. The type of equipment used for measurement and analysis of all acoustic, meteorological, and engine performance data shall be reported.

3. The following atmospheric environmental data shall be measured at representative microphone positions.

(a) Air temperature in degree Celsius and relative humidity in percent.(b) Maximum, minimum, and average wind velocity in knots and wind direction

relative to engine centerline. (c) Atmospheric pressure in psi.(d) Comments on local topography, ground cover, and events that might

interfere with sound recordings shall be reported. 4. All data shall be corrected to the following:

(a) Sea level pressure of 14.7 psi.(b) Ambient temperature of 77°F.(c) Relative humidity of 70 percent.(d) Zero wind.

4.6.4.11 Exhaust Gas Emission Test.

4.6.4.11.1 Exhaust Smoke Emission. The engine shall be subjected to an exhaust smoke emission test to demonstrate compliance with 3.2.5.8.1. Engine exhaust smoke measurements shall be taken using the equipment, instrumentation, and test procedures set forth in ARP 1179D. Smoke level will be determined at six engine power settings: maximum, intermediate, maximum continuous, 75 percent maximum continuous, 25 percent maximum continuous, and idle. Prior to sampling at a particular power setting, the engine shall become fully stabilized by running a minimum of a ten minutes stabilization run at that power setting. A performance check need not be made after this test.

4.6.4.11.2 Invisible Exhaust Mass Emissions. During the test of 4.6.4.11.1, the engine exhaust gases shall be analyzed for nonvisible contamination using the equipment, instrumentation, and procedures in ARP 1256D and the data reduction of ARP 1533A. The tests shall be conducted at maximum, intermediate, maximum continuous, 75 percent maximum continuous, 25 percent maximum continuous, and idle power settings to demonstrate the levels of invisible exhaust mass emissions specified in 3.2.5.8.2. Prior to sampling at a particular power setting, the engine shall become fully stabilized by running a minimum of a ten minutes stabilization run at that power setting. A performance check need not be made after this test.

4.6.4.12 Attitude Test. The engine shall be subjected to an attitude test to demonstrate compliance with 3.2.1.5.1 and 3.7.7.3. Engine capability to operate for the time specified in 3.2.1.5.1 for both negative “g” and zero “g” conditions shall be verified by analysis or by a rig test of the lubrication oil system. The engine shall be started, and then operated at the maximum rated rotor speeds specified in Table IA, IB and IC for at least 30 minutes at each of the test points shown in the clear area of Figure 9. The engine shall also be operated at maximum rated rotor speeds for at least 30 seconds ateach of the six test points shown in the shaded area of Figure 9. This test shall be considered

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satisfactorily completed when, in the judgment of the Using Service, the engine starts satisfactorily, remains within all operating limits, and there is no evidence of mechanical damage. This test may be waived by the Using Service if significant components are substantially identical to those successfully completing the test in 4.5.4.4.

4.6.4.13 Loss of Oil Test. A loss of oil test shall be performed IAW 4.5.4.5 if significant components have been changed from the configuration of PFR.

4.6.4.14 Electromagnetic Environmental Effects (E3). The requirement of 3.2.8.4 shall be verified by analysis and test.

4.6.4.14.1 Electromagnetic Interference (EMI). The requirements of 3.2.8.4.1 shall be verified by analysis and testing IAW ADS-37-PRF.

4.6.4.14.2 Not used.

4.6.4.14.3 Not used.

4.6.4.14.4 Lightning. The requirement of 3.2.8.4.4 shall be verified by testing IAW RTCA/DO-160F, Section 22.

4.6.5 Engine Characteristics and Fuel Tests. The tests in the following subparagraphs shall be conducted on engines having the same parts list and configuration as the endurance test engine of 4.6.1. Unless otherwise specified in the individual test, the engine need not be calibrated before and recalibrated after each test. Unless otherwise specified for a particular test, the test shall be conducted at the ambient conditions which prevail at the test site, using the fuel specified in following subparagraphs.

4.6.5.1 Starting Torque. The engine shall be subjected to a test to demonstrate the starting torque and speed requirements of 3.7.9.1.1. The measurement procedures and the calibration and use of test equipment shall be defined in the pretest data. The test shall be conducted using MIL-PRF-5624 fuel, grade JP-4, JP-5 and MIL-DTL-83133H fuel, grade JP-8.

4.6.5.2 Not used.

4.6.5.3 Not used.

4.6.5.4 Not used.

4.6.5.5 Emergency Fuel Test. The engine shall be subjected to emergency fuel tests consisting of two six hour cycles in accordance with 4.6.1.3 using each emergency fuel specified in 3.7.3.2.3. Separate tests shall be conducted for each fuel specified. All tests can be conducted on a single engine with appropriate inspections of the hot section after each fuel tested. Changes in performance due to the use of each emergency fuel shall be determined by calibrations with primary and emergency fuel prior to each test and recalibration with both fuels after each test. Calibrations shall be made to the extent required to establish performance levels

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and as defined in the pretest data. At test completion, the engine shall be disassembled to the extent necessary to perform a hot section inspection. This test will be considered satisfactorily completed when, in the judgment of the Using Service, engine performance meets the requirements specified in 3.7.3.2.3, and results of the hot section inspection do not reveal abnormal hot section distress, parts failures, or impending failures.

4.6.6 Structural Tests. Engines or components conforming to the parts list and configuration of the qualification endurance test engine shall be used for the following tests.

4.6.6.1 Emergency Power Demonstration An Emergency Power Demonstration (EPD) test shall be performed on the engine hardware that successfully completes either the endurance test of 4.6.1 or the Low Cycle Fatigue (LCF) test of 4.6.6.2.2. Following successful completion of exit criteria from the endurance or LCF test, the used rotating hardware shall be inspected and condition documented as a start point for the EPD test. With acceptance of the Using Service, some hardware displaying damage beyond service limits may be replaced prior to this test. For the purposes of this test a loss of more than 20% shaft horsepower while on point constitutes imminent engine failure and the test may be terminated. The test rotor speeds shall be at maximum allowable steady state or higher. The starting point for this test shall be a target T4.1 temperature 40°F higher than that temperature demonstrated in 4.6.6.4.2. After 2 minutes at this thermal condition, T4.1 is to be increased an additional 40°F for another 2 minute segment. Continue to elevate turbine inlet temperature in this manner for 2 minute time segments until imminent engine failure as defined above occurs. The EPD test will demonstrate the non-catastrophic failure of the engine at higher than contingency power. Successful completion of the test will show:

a. The engine does not experience an uncontained fire.b. There is no catastrophic rotor, support or engine mount failure.c. The engine successfully demonstrates the containment requirements of 3.3.8.9.1.d. The engine does not lose the ability to be shut down.

4.6.6.2 Low Cycle Fatigue Tests.

4.6.6.2.1 Low Cycle Fatigue Component Tests. Selected critical parts shall be subjected to testing to show compliance with 3.3.8.3 life requirements. The selected parts shall be cycled at engine operating conditions to a level sufficient to demonstrate the -3σ required life with a 95% confidence level. This testing procedure, reviewed and approved by the Using Service, shall provide for a component “cleared life” for the life requirements specified in 3.3.8.3. The critical engine components listed in Table IX shall be subjected to individual component spin pit tests until the required LCF lives are cleared. Dummy blades can be used if necessary to duplicate engine stress ranges as closely as possible. Finite element models shall be used to determine the proper cyclic speeds, feature stress/strain ranges, rig test configuration and to ensure the LCF critical site in each spin test component remains the same as in the engine. The spin pit critical LCF site location for each component shall be maintained as close as possible to that in the actual engine to avoid a potential failure at a different location induced by the spin pit test condition. If such a failure away from the critical site occurs, the life of the component before

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failure can be credited in a statistical analysis. Additionally, the spin rig cyclic stress ratio (R ratio) shall remain as close as possible to the engine R ratio. With Using Service approval, the contractor may use tested parts from the engine LCF test (4.6.6.2.2) to shorten required component test time, if the engine test parts clear all necessary inspections. The test procedures shall be as specified in the pretest data. During these tests, no repair shall be permitted for the first equivalent life time. Permissible component repair intervals after the first equivalent life time shall be as specified in the pretest data. All component fatigue tests shall be performed with operating temperatures and loads appropriate for realistically simulating engine operating conditions. All repairs and parts replacement requests shall be reviewed by the Using Service for approval and shall be recorded and reported. Test conduct of critical parts may require other parts to receive cyclic damage greater than their required/predicted LCF life. Part replacement or repair of these pre-test identified parts may then be accomplished in order to continue the test, with approval of the Using Service. The LCF component test shall be considered successful and the minimum lives of 3.3.8.3 verified, if no components under test fail within the prescribed test. Failure is defined as generation of a crack size per 3.3.8.11.2 at any time during or at the conclusion of the component test.

4.6.6.2.2 Low Cycle Fatigue Engine Test. A low cycle fatigue test of 3750 cycles shall be performed. No manual adjustment of the engine control system components is required. The accessory pads shall be loaded to provide maximum continuous loads. The engine components shall be calibrated prior to the test and recalibrated after the test in accordance with paragraph 4.6.1.2.1.

For this engine test, the test cell shall conform to the power absorption device requirements of 4.3.3.3.2.

The engine shall be calibrated prior to the test and recalibrated after the test. The calibration shall be conducted in accordance with paragraphs 4.6.1.2.2a, 4.6.1.2.2b, and 4.6.1.2.3 except compliance with the starting torque requirement (paragraph 3.7.9) need not be shown. The recalibration shall be in accordance with paragraphs 4.6.1.2.2a, 4.6.1.2.2b, and 4.6.1.2.3 except that compliance with the starting torque requirement need not be shown. The recalibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature condition, the values of output shaft power obtained during the initial calibration. The recalibration may be preceded by a specified run during which the cleaning procedure of paragraph 3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration. The test will be considered to be satisfactorily completed when the engine has completed the test and during recalibration, the steady state first stage turbine rotor inlet gas temperature does not exceed a value of the gas temperature obtained for the initial calibration plus 30 percent of the difference between the maximum allowable (minimum endurance) steady state first stage turbine rotor inlet gas temperature and the rated temperatures specified in Table IV for all ratings; the corrected specific fuel consumption does not exceed 105 percent of the calibration values; the engine meets all other specified performance requirements which can be checked by the calibration procedure; and, in the judgment of the Using Service, the test engine and components are operating satisfactorily at the end of the tests and teardown inspections do not disclose parts failure or impending failures. The final corrected specific fuel consumption shall be reported. First-stage turbine rotor inlet gas temperature deterioration and SFC increase shall be determined at the same rated output shaft power.

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The LCF test shall be run at constant power at a power turbine speed of 105%. The establishment of the maximum power settings at the start of the test shall be as follows:

For operation at maximum power this shall be at the power corresponding to the pre-test calibration at a gas temperature at or above the rated gas generator first stage turbine rotor inlet temperature specified in Table VI for the maximum rated power setting at 105% power turbine speed.

Following completion of the test, the engine shall be disassembled and inspected for evidence of cracking and shall be within allowable limits. The low cycle fatigue test cycle shall be as follows:

Approximate Approximate total time schedule time (minute) (minute) Event 0.5 0.5 Start engine 2.5 2.0 Run at idle 2.6 0.1 Accelerate to maximum power 10.6 8.0 Run at maximum power 10.7 0.1 Decelerate to idle 12.7 2.0 Run at idle 15.0 2.3 Shutdown and cool down

4.6.6.3 Containment. A blade containment analysis shall be performed to determine the critical stage of each module with the lowest containment margin. This containment margin relates the released kinetic energy of a blade fragment (as defined in 3.3.8.9.1) to the energy absorption capability of the engine structures for containment. For the purposes of the critical stage containment assessment, modules shall be defined as a compressor, gas generator turbine or power turbine rotating group at a common speed. The containment analysis shall be provided at the initiation of QT to define which stage of each module shall be tested. The analysis shall be substantiated and correlated with actual containment tests of similar engine hardware. Prior failures at different energy levels on identical structures shall be presented and discussed in the analysis report and used as supplemental support for meeting containment requirements. The requirements of 3.3.8.9.1 shall be demonstrated by either engine or spin pit tests. These tests shall be conducted at or above the maximum allowable transient rotor speeds specified in 3.2.1.4.6 and maximum component operating temperatures per 3.2.1.4.5. All containment demonstrations (engine and spin pit) shall include increased speeds to account for minimum material properties and case thicknesses. If a spin pit demonstration is selected, fragment release speed shall be increased to account for a lower spin pit test temperature. Selected compressor and turbine blades determined to be most critical by analysis shall be undercut to provide the fragment size required by 3.3.8.9.1 and to fail at the required demonstration speed. Engine tests to demonstrate containment of turbine rotor blades shall be conducted with an exhaust pipe conforming to 3.3.8.9.1. Prior to each containment test, a test plan including all details for the calculation of the required test speed shall be provided to the Using Service for review and approval. The test will be considered completed when, in the judgment of the Using Service all requirements of paragraph 3.3.8.9.1 are met. Component containment analyses and testing shall be performed on other critical rotating parts including the inlet particle separator rotor. The analyses and tests shall be conducted to ensure all engine

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components utilizing rotating parts shall contain any rotating part failure at maximum transient speed.

4.6.6.4 Rotor Integrity

4.6.6.4.1 Overspeed. The most critical stage of each rotor system including the power turbine, gas generator turbine, compressor, and inlet particle separator, as defined by analysis and agreed to by the Using Service, shall be subjected to engine or spin pit operation for a stabilized period of at least five minutes duration at component temperatures predicted to occur during operation at stabilized maximum allowable first stage turbine rotor inlet temperature (3.2.1.4.5) at one of the following speed conditions. If an ambient temperature spin pit is used, additional demonstration speed shall be added to accommodate material property scatter (average to -3σ yield strength values), yield strength differences resulting from the temperature difference between engine operating conditions and spin pit test conditions, and actual versus average yield strength material properties in accordance with the guidance of the Using Service. If a heated spin pit or engine test is used, additional demonstration speed shall be added to accommodate material property scatter (average to -3σ yield strength values) and actual versus average yield strength material properties in accordance with the guidance of the Using Service. Prior to any testing, critical disk dimensions (such as rotor bores and other critical pilots) agreed to by the Using Service shall be recorded for comparison to post test measurements. Gas generator turbine and compressor rotors shall be subjected to 115 percent of the maximum allowable steady state speed limit specified in 3.2.1.4.6. Power turbine rotors shall be subjected to the greater of the following two methods:

a. 115 percent of the transient speed limit specified in 3.2.1.4.6.b. 105 percent of the predicted peak speed attained in a loss of load analysis per the

requirements stated in 3.7.2.3.3.1, conducted throughout the engine operating envelope at worst case ambient temperature and altitude conditions. For calculation of this speed, the analysis shall assume loss of load at maximum rated power redline conditions (min endurance test temperature per Table VI), 100% power turbine design speed per 3.2.1.4.6, engine performance model approved by the Using Service and worst stackup of overspeed control system tolerances from the loss of load event until the point at which the overspeed device actuates. Following these overspeed tests the power turbine, compressor and gas generator turbine parts and assemblies shall be measured at the same critical locations as pre-test to determine residual growths. Analysis shall demonstrate why/how these residual growths are acceptable or unacceptable for continued disk operation after an overspeed event. The parts and assemblies shall show no evidence of impending failure, in the judgment of the Using Service. This test may be waived for components identical to those successfully completing the test in 4.5.4.2.1.

4.6.6.4.2 Overtemperature. Using the same engine rotors that the completed the overspeed test, the engine shall be operated at a first-stage turbine rotor inlet temperature of at least 45°C (81°F) in excess of the maximum allowable first stage turbine rotor inlet temperature as specified in 3.2.1.4.5a, at no less than maximum allowable steady state speed specified in 3.2.1.4.6 for all rotors, for five minutes.

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Following the test, parts and assemblies shall comply with 3.3.8.8 and there shall be no evidence of impending failure. This test may be waived for components identical to those successfully completing the test in 4.5.4.2.2.

4.6.6.4.3 Disk Burst Disk burst testing shall be conducted to evaluate whether the burst margin requirement of 3.3.8.9.4 can be met with a minimum tensile strength disk (based on the minimum properties specified in 3.3.8.5). Disk burst testing shall be conducted on all engine disks. As a minimum, with the approval of the Using Service, disk burst tests may be conducted on the most limiting disk (disk with the minimum burst capability) of each rotor spool.

Disks shall be operated at burst speeds no less than those of 3.3.8.9.4 while exposed to the maximum temperature gradient and maximum material temperature that would occur for that part. Maximum test speed shall be sufficient to demonstrate that a minimum tensile strength component (–3 Sigma) can meet the burst margin requirement based on the specific ultimate strength capability of the test component. These conditions shall be maintained for a minimum of 45 seconds. The test shall be considered successfully completed if there is no evidence of imminent failure.

Substitute blades (dummy blades) may be used in lieu of actual disk blades during the disk speed evaluation if analysis is provided and the Using Service agrees that at the test speed, the centrifugal load of the dummy blades is equal to or greater than the actual blade load.

4.6.6.5 Static Load Tests.

4.6.6.5.1 Engine Static Load Test. The engine cases and mounts conforming to the same parts list and configuration of the endurance test engine shall be subjected to a static test to verify the requirements of 3.1.2.5 and 3.1.2.6. A static rig test utilizing a full engine carcass shall be conducted to demonstrate the capability of the engine and its supports to meet the requirements of this plan during and after exposure to maximum externally applied forces specified in Figure 1 without permanent deformation of any component and 1.5 times those forces without failure of any component. In this test, maximum thrust loads, acceleration loads, gyroscopic moments, torque, and engine reaction loads shall be applied separately and then in combination for all mount systems listed in paragraph 3.1.2.1 (a.), “Engine Configuration and Envelope Figure”. Stress and deflection data shall be obtained at critical locations as determined by analysis and preliminary stress coating tests. The limit loads shall be based on a mass factor consisting of the dry mass of the engine, increased by the specific mass allowed for all engine mounted accessories and operating fluids.

4.6.6.6 Vibration and Stress Tests. The vibration and stress tests consist of the following separate and distinct tests:

a. Compressor Strain Testb. Gas Generator Turbine Strain Testc. Power Turbine Strain Testd, Gas Generator Rotor Bearing Evaluation e. Power Turbine Rotor Bearing Evaluation

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f. Engine Vibration Surveyg. Stinger Rig Vibration Test

4.6.6.6.1 Compressor Strain Test. A compressor strain test shall be conducted on an engine to obtain data to substantiate the vibration and stress analysis report of 3.3.8.10.2 and the report 3.3.8.6. The test shall be performed on an engine with the compressor components sufficiently instrumented with strain gages to obtain continuous strain gage data throughout the engine operating speed range. Each strain gage shall be mounted on critical compressor components as close as possible to the location where the minimum high cycle fatigue (HCF) margin occurs, as determined from the vibration and stress analysis.

4.6.6.6.2 Gas Generator Turbine Strain Test. A gas generator turbine strain test shall be conducted on an engine to obtain data to substantiate the vibration and stress analysis report of 3.3.8.10.2 and the report 3.3.8.6. The test shall be performed on an engine with the gas generator turbine components sufficiently instrumented with strain gages to obtain continuous strain gage data throughout the engine operating speed range. Each strain gage shall be mounted on critical gas generator turbine components as close as possible to the location where the minimum high cycle fatigue (HCF) margin occurs, as determined from the vibration and stress analysis.

4.6.6.6.3 Power Turbine Strain Test. A power turbine strain test shall be conducted on an engine to obtain data to substantiate the vibration and stress analysis report of 3.3.8.10.2 and the report 3.3.8.6. The test shall be performed on an engine with the power turbine components sufficiently instrumented with strain gages to obtain continuous strain gage data throughout the engine operating speed range. Each strain gage shall be mounted on critical power turbine components as close as possible to the location where the minimum high cycle fatigue (HCF) margin occurs, as determined from the vibration and stress analysis.

4.6.6.6.4 Gas Generator Rotor Bearing Evaluation. A bearing evaluation test shall be conducted on an engine to document gas generator rotor bearing thrust loading, bearing cage speeds, and bearing outer race temperatures throughout the engine operating envelope. Sufficient instrumentation shall be installed at appropriate locations on gas generator rotor bearings to permit measurement of the required bearing parameters.

4.6.6.6.5 Power Turbine Rotor Bearing Evaluation. A bearing evaluation test shall be conducted on an engine to document power turbine rotor bearing thrust loading, bearing cage speeds, and bearing outer race temperatures throughout the engine operating envelope. Sufficient instrumentation shall be installed at appropriate locations on power turbine rotor bearings to permit measurement of the required bearing parameters.

4.6.6.6.6 Engine Vibration Survey. A vibration survey which demonstrates compliance with 3.2.1.4.10 shall be conducted. The vibration survey shall include, but not necessarily be limited to, data showing true RMS velocity spectrograms and peak acceleration spectrograms for each designated sensor location at the highest vibration point in the operating envelope (which shall be identified) and at designated

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engine rating points. The spectrograms shall cover the frequency range of 3 Hz to 5 kHz. The method used for determining the overall true RMS velocity from the spectrogram and the maximum permissible overall true RMS velocity limit shall be described. The engine to be used for the vibration survey shall be assembled with at least the maximum allowable imbalances specified for the gas generator spool and power turbine module. The test stand shall comply with 4.3.3.3.

4.6.6.6.7 ”Stinger” Rig Vibration Test. A "stinger" rig vibration test shall be conducted on an engine to verify that external component resonances excited by airframe induced loads are not harmful to the engine during normal operation. The test shall be performed on an engine carcass with all external and internal components. External components including tubes, fuel control, valves, main oil tank, electronic control unit, inlet particle separator blower, etc. shall be instrumented with accelerometers or strain gages as appropriate. Low amplitude sweeps between 5 and 2000 Hz shall be conducted on the engine carcass in a free-free boundary condition. Vibration energy shall be transmitted from a portable vibration exciter, through a stinger rod attached to the engine, exciting the natural frequencies of the external components. Upon test completion, a test report shall be submitted. 4.6.6.7 Gyroscopic Test. An engine shall be subjected to a gyroscopic test to demonstrate compliance with 3.1.2.5.1. Prior to the test, the engine shall be assembled with special emphasis placed on measuring and recording clearances between blades and cases and radial and axial rotor clearances. Rub probes shall be installed around compressor and turbine cases at symmetrical locations and at blade tip locations as designated in the pretest data. Instrumentation shall be sufficient to permit measurement of rotor deflection and shift under gyroscopic loads. Strain gage instrumentation shall be provided to measure stresses at critical locations. Sufficient instrumentation of the oil system shall be provided to evaluate the oil system’s ability to scavenge and function properly during the test. Test data to be taken during the test, in addition to the data required above, shall include vibration measurements at locations as specified in 3.7.6.5. The engine shall be installed on a gyroscopic test stand with an inlet configuration and exhaust nozzle as designated in the pretest data. Prior to the test, the engine shall be subjected to a performance check. The test shall be conducted with the gyroscopic rig operated in incremental steps of 0.5 radians per second up to and including 2.5 radians per second. At each step, the engine shall be operated as follows:

a. Idle for one minute.b. Accelerate from idle to maximum allowable steady state rotor speed.c. Dwell at maximum allowable steady state gas generator speed and maximum

allowable power turbine speed for 10 seconds or time sufficient to record data. d. Decelerate from maximum allowable steady state gas generator speed and

maximum allowable power turbine speed to idle in 30 seconds. e. Stop rig and engine to visually check for indications of rub.

At the 2.5 radian per second point, the engine shall be operated as follows: a. Start engine and stabilize at idle for 1 minute.b. Accelerate the gyroscopic rig to 2.5 radians per second.c. Immediately accelerate the engine to maximum allowable steady state gas

generator speed and maximum governing power turbine speed in approximately 7 seconds.

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d. Stabilize at maximum allowable steady state gas generator speed and maximum governing power turbine speed until total engine exposure time at 2.5 radians per second angular velocity reaches 30 seconds. Reduce rotation to 2.0 radians per second at that point.

e. Decelerate the engine to idle.f. Stop the rig and engine to visually check for indications rub.

NOTE: At gyroscopic loads above 1.5 radians per second snap accelerations and decelerations may be made to reduce time exposure. The total time at 2.5 radians per second gyro load shall be IAW 3.1.2.5.1. The above test shall be conducted with the gyroscopic rig rotating in one direction and then the test shall be repeated with the rig rotating in the opposite direction. At the completion of the test, the engine shall be subjected to a post test performance check and then disassembled for inspection. The test shall be satisfactorily completed when, in the judgment of the Using Service: the post test calibration reveals no significant loss in performance, the engine and its systems operated properly during the test, structural loads were within acceptable limits, and teardown inspection reveals no evidence of excessive blade tip or seal rubbing or evidence of impending failure.

4.6.6.8 Engine Overtemperature Control System Test. An engine, conforming to the same parts list and configuration as the endurance test engine, shall be tested to verify all capabilities of the engine overtemperature system which consists of transient, steady state and start temperature limiting. This will be demonstrated by showing the control’s temperature limiting systems can be activated and that the systems perform as designed. Furthermore the capability of the overtemperature system to prevent the engine from exceeding the maximum measured gas temperature value specified in 3.2.1.4.5b shall be demonstrated. Other control limiting functions may be altered or disabled as required to ensure the overtemperature system is engaged and the measured gas temperature does not exceed the limits specified in 3.2.1.4.5b.

4.6.6.9 Engine Overspeed Control System Test. An engine, conforming to the same parts list and configuration as the endurance test engine, shall be tested to verify the adequacy of the power turbine overspeed control system described in 3.7.2, 3.7.2.2, and 3.7.2.3.3.1. The engine shall be set up in a test cell with a shaft power absorber that can be instantaneously uncoupled from the engine. The power turbine shaft speed shall be adjusted to 100% speed per 3.2.1.4.6 with the engine operating at maximum rated power. After 10 minutes at this operating condition, the shaft power absorber shall be instantaneously disconnected from the engine. Power turbine shaft acceleration rates and terminal speed data shall be recorded during this test and used to calibrate/verify the analytical speed prediction program used in the 3.7.2.3.3.1 envelope analysis. This test will be considered satisfactorily completed if the power turbine overspeed control system prevents destructive overspeed of the power turbine.

4.6.6.10 Main Shaft Bearings Assurance Testing. Bearing endurance testing shall be conducted for each main shaft bearing. A total of at least 1000 hours of endurance, AMT, and/or LCF testing shall be accumulated on each of three sets of main shaft bearing design with at least 500 hours accumulated on a single bearing of each design. Main shaft bearings shall be instrumented in an engine to measure temperatures (outer race), cage speeds, and thrust loads for the engine operating range which includes hot shutdown soakback and restart conditions. Demonstration methods to acquire inner race

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temperature data include engine, gas generator, or rig tests. This test or portions thereof may be waived by the Using Service if significant components are substantially identical to those successfully completing the test in 4.5.4.9.

4.6.6.11 Gear Resonance test. Gear resonance testing shall be conducted on all gears to demonstrate freedom from damaging resonance (see 3.3.8.10).

4.7 Acceptance Test (AT). An acceptance test, as specified herein, shall be conducted on each engine submitted for delivery. A detailed Acceptance Test Procedure shall be submitted to the Using Service for approval. Engines submitted for EIT or QT testing need not be subjected to acceptance tests. Acceptance Test Procedures for all controls and other external components shall be prepared by the contractor and submitted to the Using Service for approval. Any component utilized for bench qualification or engine EIT and QT shall be tested IAW the approved ATP. At the Using Service’s option, formal approval of any individual ATP may be waived until completion of the qualification program. However, ATPs shall be maintained under configuration control so that all pre- and post-test component performance checks are conducted using the same ATP baseline.

4.7.1 Test Apparatus.

4.7.1.1 Automatic Recording Equipment. Automatic continuous recording equipment shall be used to record data during the execution of those parts of the engine test requiring the evaluation of engine variables versus time.

4.7.1.2 Vibration Measuring Equipment and Response Characteristics. The engine vibration shall be measured with sensors. The vibration measurement and analysis equipment shall operate over a frequency band of at least 3 Hz to 5000 Hz and produce acceleration spectrograms having a demonstrated accuracy with confidence level of 95 percent. The maximum allowable effective filter bandwidth of the spectrum analysis equipment shall be 3 Hz up to 1000 Hz and 6 Hz above 1000 Hz. The vibration measuring equipment shall be calibrated as a complete system. The frequency response of the system, when calibrated by applying a known sinusoidal motion to the sensor, shall not deviate by more than ±3.0 dB from the known sinusoidal input at frequencies from 3 Hz to 5000 Hz. If filters are required when measuring overall velocity levels, they shall be not more than 3.0 dB down at the cutoff frequencies, which shall be 30, 70, or 100 Hz high pass, as appropriate, and they shall have a roll-off rate of at least 18 dB per octave. High pass filters shall not be used to produce velocity and acceleration spectrograms.

4.7.1.3 Test Stand and Test Equipment. Engines shall be subjected to acceptance testing in a test cell and with test equipment that is acceptable to the Using Service.

4.7.1.3.1 Dynamic Characteristics. Vibration shall be measured with the engine operating on a test stand which has natural frequencies with the engine installed no higher than 50 percent of the idle rotor speed in all modes of motion which can be excited by residual rotor unbalances. The test stand shall not induce damaging or detrimental resonance into the engine at any test or operating condition.

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4.7.1.3.2 Power Absorption Characteristics. To provide realistic engine accelerations and decelerations, the engine power absorber shall have a polar mass moment of inertia (referred to the output shaft speed) within ±5% of intended field application for all engine acceptance tests.

4.7.1.4 Starter. Starting shall be performed with a starter that has torque characteristics similar to the minimum required torque shown in Figure 17 with the following requirements.

a. Starting torque shall not exceed the minimum required torque of Figure 17 by more than 5%, and

b. Overall starting torque when integrated over the start speed range shall not exceed the minimum starter torque by more than 5%.

4.7.2 Test Conditions.

4.7.2.1 Servicing.

4.7.2.1.1 Oil Servicing. The oil used for acceptance testing shall conform to MIL-PRF-23699. All oil filter inspection results shall be recorded in the engine records.

4.7.2.1.2 Fuel Servicing. The fuel used for acceptance testing shall be any primary fuel IAW 3.7.3.2.1. The fuel reference heating value used in acceptance testing will have a lower heating value of 18,300 BTU/lb.

4.7.2.2 Electrical and Electronic Interference and Susceptibility Check. Electrical and electronic system or components of production engines shall have the same EMI characteristics as the configuration which demonstrated compliance with paragraph 4.5.5.2 (PFR) or 4.6.4.14.1 (QT). If a change is made which, in the judgment of the Using Service, might affect the EMI characteristics, then the systems or components incorporating the change shall be tested in accordance with applicable parts of paragraph 4.5.5.2 (PFR) or 4.6.4.14.1 (QT) until the systems or components have passed the tests without reworking. The Using Service under future and separate production contracts may require an EMI sampling plan.

4.7.2.3 Component Calibration. Functional bench calibration/checks and/or ATPs shall be conducted on each of the components specified in Table XVIII. All fuel nozzles and fuel carrying components of the engine control system shall undergo bench calibrations using fluid in accordance with 3.7.3.3.1 or pass approved ATPs. Components shall conform to the design tolerance range required by the applicable design specifications. All control system sensors shall be calibrated and/or pass an ATP. Temperature sensing system performance shall meet the tolerance and thermal response characteristics specified in 3.7.6.4 of this AQP.

4.7.2.4 Not Used.

4.7.2.5 Environmental Stress Screen. Each component shall undergo an environmental stress screen (ESS) during, or in addition to, the calibration or ATP performed under paragraph 4.7.2.3. The ESS shall conform to HASS (Highly Accelerated Stress Screen) standards and methodologies. The HASS shall be approved

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by the Using Service. For simple components, the requirement to perform an ESS may be waived at the discretion of the Using Service.

4.7.3 Test Records.

4.7.3.1 Acceptance Test Log Sheet. The following information shall be legibly recorded on the test log sheets for each engine run:

a. Dateb. Engine Type and Modelc. Engine Serial Numberd. Cell Numbere. Bellmouth Serial Number and Area, square centimetersf. Exhaust Nozzle Serial Number and Area, square centimetersg. Type of Fuel Usedh. Type of Oil Usedi. Fuel Lower Heating value and specific gravityj. Total Running Time and Total Number of Startsk. Vibration (Max Recorded in 4.7.4)l. All Data Required in 4.7.4, except 4.7.4.3.

The contractor shall retain legible copies of acceptance test log sheets for each engine for two years. Copies of test sheets shall be furnished to the Using Service upon request.

4.7.4 Test Data. The data in the following subparagraphs shall be taken during the acceptance test and recorded on the acceptance test log sheets or engine data forms.

4.7.4.1 Preliminary Data. The engine mass shall be determined and recorded. If the engine mass is measured after the engine has been serviced with fuel and oil, and subsequently drained, the dry mass may be calculated by subtracting the mass of residual fluids specified in 3.2.2.2 from the measured engine mass. The effective flow area of each turbine nozzle shall be determined and recorded.

4.7.4.2 Steady-State Data. During operation at each specified steady state condition and after performance stabilization, the following minimum data shall be recorded once during each test period.

a. Time of day.b. Total running time.c. Power demand signal.d. Engine rotor speeds, rpm.e. Data for calculating output shaft power.

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f. Fuel consumption, lb/h.g. Data for determining airflow, lb/s.h. Engine inlet total pressure, psi. (Also compressor inlet total pressure average for

engines with IPS). i Engine inlet total temperature, °C (°F). (Also compressor inlet total temperature

average for engines with IPS). j. Customer bleed air extraction airflow, lb/s.k. Exhaust static pressure, psi.l. Not used.m. Oil temperature at the point shown on the Engine Configuration and Envelope

Figure, °C (°F) n. Oil pressure at the point shown on the Engine Configuration and Envelope Figure,

psi.o. Oil outlet temperaturep. Fuel pressure at the fuel system inlet, psi.q. Fuel pressure at the point shown on the Engine Configuration and Envelope

Figure. r. Measured steady state gas temperature, °C (°F).s. Data for calculating gas generator first stage turbine rotor inlet gas temperature,

°C (°F). t. Not used.u. Engine case vibration at points shown on the Engine Configuration and Envelope

Figure, inches per second, true RMS velocity. v. Oil leakage at accessory pads and drains.w. Additional data as required by the Using Service.

4.7.4.3 Transient Data. For each transient performed during the power transient runs the parameters of power demand, measured temperature, engine speed (output and gas generator), compressor discharge pressure, fuel flow, and output shaft power shall be continuously recorded versus time.

4.7.4.4 Starting Data. For each start performed, the time required from initiation of the start to: ignition, starter cutout, and stabilized engine idle speed shall be recorded as well the engine speed where each of event occurs. The maximum measured temperature shall be recorded for each start.

4.7.4.5 Miscellaneous Data. All stops and coast down times shall be measured and recorded. At least once during each test, readings shall be taken of barometric pressure, ambient air temperature, water vapor pressure, and fuel specific gravity. Oil consumption for the entire test run shall be measured and recorded. Notes shall be placed on the log sheets of all incidents of the run, such as leaks, unusual vibrations, and other irregular functioning of the engine together with corrective measures taken.

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4.7.4.6 Accuracy of Data. For all engine acceptance testing, reported data shall have a steady state accuracy within the tolerances shown below. The accuracy of transient data and the corresponding instrument calibration methods shall be subject to the approval of the Using Service and shall be described in the test procedure. All instruments and equipment shall be calibrated as necessary to insure that the required degree of accuracy is maintained.

DATA ITEM TOLERANCE Rotor speed(s) +0.2 percent of the value obtained at maximum ratingFuel Flow +0. 5 percent of the value measured for maximum continuous

rating and above+0. 5 percent of the value measured at maximum continuousrating for all values below intermediate rating

Torque +0. 5 percent of the value measured for maximum continuousrating and above+0. 5 percent of the value measured maximum continuous ratingfor all values below intermediate rating

Airflow +1.0 percent of the value measured for maximum continuousrating and above+1.0 percent of the value measured at maximum continuousrating for all values below intermediate rating

Temperatures ±1.0°C (1.8°F) up to 200°C (392°F) ±3.0°C (5.4°F) above 200°C (392°F) ±4.0°C (7.2°F) above 800°C (1472°F)

Engine Weight +1.0 lb or +0.1 percent of the weight being determined,whichever is greater

Vibration Velocity +10.0 percent of specified engine limitAll other data "±2.0 percent of the value obtained at maximum rating

4.7.5 Test Procedure. The acceptance test shall consist of the initial and final runs specified below at ambient inlet conditions. (See 4.7.7 if an engine sampling plan is selected.) When special engine features which would not function under the following test schedules are provided, these features shall be tested in a manner approved by the Using Service. Recorded time at each power condition shall start after reaching steady state measured gas temperature. During the final run, for each rating condition, the engine shall produce rated power, or higher, and rated specific fuel consumption, or lower, within the Table IA limits of gas temperature and rotor speed. The engine shall not produce a first stage turbine rotor inlet gas temperature (contractor shall specify designation) higher than the first stage turbine rotor inlet temperature for any corresponding rating Table IA gas temperature limit for the respective rating condition. If any adjustment to the engine or its components becomes necessary, after starting the final run, there shall be a rerun of those portions of the test run already completed. No load shall be considered as being 20 SHP or less. The contractor may tailor the acceptance test procedure with the approval of the Using Service. If no-load conditions cannot be obtained with the power

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absorber, then the transient time requirement will be decremented by the time equivalent to the min load as obtained from Figure 10-2. The first 30 engines shall be delivered based on first stage turbine rotor inlet gas temperature. The contractor shall monitor the first stage turbine rotor inlet gas temperature/measured gas temperature relationship during these first 30 engines in order to get a representative relationship. The remaining production engines shall be sold on a measured gas temperature basis.

4.7.5.1 Initial Run. The engine shall be subjected to an initial run in accordance with the following schedule. The steady state power times may be increased as an option if needed to obtain stable performance data required to verify sea level static rated performance. All power demand requests shall be accomplished in one half second or less. The nature and extent of checks, adjustments, and running prior to the initial run shall be specified. Gas temperature, as used during the runs described below, is defined as the first stage turbine rotor inlet gas total temperature, averaged over the gas path area.

a. Intermediate and Maximum Power Transients Run. This run shall consist of 24 minutes in the sequence of power conditions and time durations as follows:

1. Six minutes at idle.2. Seven minutes at the rated gas temperature and the rated output shaft

speed, all as specified for the intermediate rating. 3. Six minutes at the rated gas temperature and the rated output shaft speed,

all as specified for the maximum rating. 4. Repeat 2 above for five minutes

b. Transients Run. This run shall consist of approximately 19 minutes of powertransients as follows:

1. Three minutes at no load condition followed by one minute at the rated gastemperature, and the rated output shaft speed, all as specified for the maximum rating followed by a decrease to no load.

2. Three minutes of operating at no load followed by one minute at the ratedgas temperature, and the rated output shaft speed, all as specified for the intermediate rating followed by a decrease to no load, and retain this power condition for 3 minutes.

3. Three minutes at the rated gas temperature and the rated output shaftspeed, all as specified for intermediate rating followed by an immediate decrease to no load condition. As soon as the engine gas generator speed reaches the no load condition, the engine power shall be increased to the measured temperature and output shaft speed associated with the previous intermediate condition and maintained at this power level for one minute. There shall be no surge during this transient sequence.

4. Repeat 3 above, except that intermediate rating shall be replaced bymaximum rating.

c. Short Transient Run. This run shall consist of 12 minutes of operation in thesequence of power conditions and time durations as follows:

1. Three minutes at no load.2. Three minutes at the rated gas temperature and the rated output shaft

speed for the maximum rated power condition.

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3. Three minutes at the rated gas temperature and the rated output shaft speed for the 50 percent maximum continuous rated power condition.

4. Three minutes at the rated gas temperature and the rated output shaft speed for the intermediate rated power condition.

d. Intermediate and Maximum Power Transient Run. This run shall consist of 10 minutes of operation in the sequence of power conditions and time durations as follows:

1, Four minutes at idle. 2. Two minutes at the rated gas temperature and the rated output shaft speed

for the intermediate power condition. 3. Two minutes at rated gas temperature and the rated output shaft speed

for the maximum power condition. 4. Repeat 2 above.

e. Performance Run. This run shall consist of 28 minutes of operation in thesequence of power conditions and time durations as follows:

1. Three minutes at idle.2. Five minutes at the rated gas temperature and rated output shaft speed

for the 75 percent maximum continuous power condition. 3. Five minutes at the rated gas temperature and rated output shaft speed

for the 90 percent maximum continuous power condition. 4. Five minutes at the rated gas temperature and rated output shaft speed

for the maximum continuous power conditions. 5. Five minutes at the rated gas temperature and rated output shaft speed

for intermediate power condition. 6. Five minutes at the rated gas temperature and rated output shaft speed

for maximum power condition. f. Slow Transient Run. This run shall consist of seven minutes of operation as

follows: 1. After three minutes at idle, the engine shall be accelerated at a uniform

rate in 90 seconds from idle to the rated gas temperature and rated output shaft speed, as specified for maximum power.

2. After one minute of operation at this condition the engine shall bedecelerated at a uniform rate in 90 seconds to idle. Throughout this run the engine vibration shall be continuously recorded using vibration equipment in accordance with 4.7.1.2. The engine vibration shall not exceed the levels specified in 3.2.1.4.10. The peak vibration and the speed at which it occurs shall be recorded.

4.7.5.1.1 Inspection After Initial Run. Upon completion of the initial run, the engine shall be disassembled sufficiently to allow a detailed inspection of all vital working parts. The extent of disassembly shall be decided by the Government representative. If any part is found to be defective, an approved part shall be supplied to replace it, and at the discretion of the Government representative, a penalty run of suitable duration shall be made.

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4.7.5.1.2 Penalty Run. The duration of the penalty run shall be at the discretion of the Government representative. The maximum penalty run shall be a complete repetition of the initial run. Additional running-in prior to the penalty run may, as an option, be performed for the accommodation of replaced parts.

4.7.5.1.3 Inspection After Penalty Run. Upon completion of the penalty run, the engine shall, at the discretion of the Government representative, be disassembled to allow for inspection.

4.7.5.2 Final Run. The final run shall consist of a repeat of the initial run of 4.7.5.1.

4.7.5.2.1 Special Control System Features. At the end of the final run, special engine control system features (e.g. channel switching, backup modes, overspeed, etc.) shall be sufficiently tested to assure functionality. Features to be tested shall be coordinated with, and approved by, the Using Service.

4.7.5.3 Rejection and Retest. Whenever there is evidence that the engine is malfunctioning or is not meeting specification requirements, the problem shall be investigated and its cause corrected to the satisfaction of the Government representative before the test is continued. If such investigation requires disassembly of the engine or any of its components, this shall be considered a rejection. At the option of the Government representative, a complete rerun or a repetition of the portion of the test prior to encountering the difficulty shall be made.

4.7.5.3.1 Engine Vibration. It shall be considered a malfunction when the vibration limits as specified in Figure 29 are exceeded.

4.7.5.3.2 Overtemperature. If at any time the temperature exceeds the maximum allowable measured temperature (transient or steady-state limit as applicable) as specified herein, this shall be considered a malfunction.

4.7.5.3.3 Stoppage. Interruptions or stoppage from any cause other than an engine malfunction shall require a repetition of the particular period during which the interruption or stoppage occurred.

4.7.5.3.4 Fluid Leakage. If fluid leaks beyond the limits specified in 3.3.6.4 are discovered, a check run or a complete rerun after correction of the leak shall be made at the discretion of the Government representative.

4.7.5.3.5 Maximum Hours of Running. Any engine which has more than a total of 10 hours of operation or more than a total of 20 starts, including all runs, checks, and adjustments of 4.7.5 and any runs, checks, and adjustments prior to the initial run, shall stand rejected. Parts and components from these rejected engines may be used in other engines being built, provided the parts and components from the rejected engines are resubmitted for inspection required for new parts and components, with full particulars being given the Government representative concerning previous rejection of the engine.

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4.7.6 Test Completion. The acceptance test shall be considered to be satisfactorily completed when the conditions of 4.7.5 have been met and the data demonstrates compliance with applicable portions of the following engine specification requirements.

a. 3.2.1 - Performance Characteristicsb. 3.2.1.1 - Performance Ratings. NOTE: All data used for comparative purposes

shall be based on rated gas temperatures in Table IA. c. 3.2.1.4.8 - Oil Pressure and Temperature Limitsd. 3.2.1.4.9 - Oil Consumptione. 3.2.1.4.10 - Vibration Limitsf. 3.2.1.5.3 - Stoppingg. 3.2.1.5.5 - Stabilityh. 3.2.1.5.6 - Engine Power Transientsi. 3.2.2.1 - Dry Weight of Complete Enginej. 3.3.6.4 - Fluid Leakagek. 3.7.7.4.2 - Oil Drainsl. 3.7.9 - Starting System

4.7.7 Sampling Plan. When a sampling plan is invoked, the selected sample engines shall be subjected to the runs of 4.7.5.1 and 4.7.5.2. Any sampling plan shall be provided to the Using Service for review and approval before implementation. All remaining engines shall be subjected to the run of 4.7.5.2.

4.8 Operational Capability Release (OCR) Test. The OCR program shall include an accelerated endurance test and two accelerated mission tests and permit correction of engine failure modes prior to field occurrence. An OCR program test report shall be submitted IAW the CDRL.

4.8.1 Accelerated Endurance Test (AET).

4.8.1.1 Endurance Test. The endurance test shall consist of 600 hours operation on an engine IAW the following test schedules, calibrations, and procedures. The engine shall be tested using one of the oils and fuels as specified in 3.7.3.2.1 and 3.7.7.2.1.

4.8.1.1.1 Pretest Verification.

4.8.1.1.1.1 Engine Dry Mass. Prior to initiation of the calibration, the dry mass of the engine as specified in 3.2.2.1 shall be verified in accordance with 4.3.5.2.

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4.8.1.1.2 Calibrations (AET).

4.8.1.1.2.1 Component Calibration. Functional bench calibration/checks shall be conducted on each of the components specified in Table XVIII. All fuel nozzles and fuel carrying components of the engine control system shall undergo bench calibrations using fluid in accordance with 3.7.3.3.1. Components shall conform to the design tolerance range required by the applicable design specifications. All control system sensors shall be calibrated. Temperature sensing system performance shall meet the tolerance and thermal response characteristics specified in 3.7.6.4 of this AQP.

4.8.1.1.2.2 Engine Calibration. The procedure during the engine calibration shall be such as to establish the performance characteristics of the complete engine. Prior to the beginning of the calibration, the engine shall be cleaned using the wash procedure specified in 3.7.13, and all engine controls shall be adjusted and shall not be readjusted throughout the calibration. Calibrations shall be made initially with no customer accessory power extraction and no bleed air extraction other than that required for continuous engine operation. During calibration, engine inlet air shall be controlled to the temperature specified in Table IA. The fuel and oil used shall be the same as those used during 4.8.1.1.3. Data indicated for calibration in 4.3.5.4 and 4.3.5.5 shall be recorded. During calibration, conformance with 3.3.6.4 and 3.3.6.5 shall be demonstrated. The following data shall be obtained:

a. Data required to establish compliance with the sea level performance ratings inTable IA, IB and IC and 3.2.1 of this AQP and to establish the accuracy of the torque sensor signal.

b. Data required establishing compliance with 3.2.1.5.5, 3.2.1.5.6, and 3.7.9 at sealevel, static conditions.

c. Repeat items “a” and “b” with maximum permissible bleed air extraction. Themaximum permissible bleed air extraction includes customer and anti-icing bleed air flow.

d. Repeat “c” with accessory power extraction as specified by the Using Service.

4.8.1.1.3 Endurance Test Procedure. Following the calibration run, the engine shall be adjusted to permit operation at the maximum allowable steady state gas temperature at the first stage turbine rotor inlet as defined in 3.2.1.4.5. This value shall be reestablished at the beginning of each cycle. The number of adjustments required shall be recorded. The engine inlet air shall be controlled where necessary during a test cycle to ensure engine operation at the specified test cycle conditions. For the purpose of the endurance test, operation at maximum continuous through contingency test temperature settings shall be defined as operation at a gas temperature at or above two thirds the difference between the maximum allowable (min endurance temperature) gas generator first stage turbine rotor inlet temperature and the minimum rated gas generator first stage turbine rotor inlet temperature specified in Table VI for the maximum continuous through contingency power settings. The engine shall operate at the MGT that corresponds to the first stage turbine rotor inlet temperature (TRIT). The relationship between MGT and TRIT for these settings shall be established via a 4-point performance calibration, prior to the test and checked at the conclusion of every 10th cycle (±2 cycles).

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For the purpose of the endurance test, minimum output shaft speed shall be interpreted as that output shaft speed at which the output shaft torque is not less than the torque limit for the applicable power condition or the minimum output shaft governing speed, whichever occurs first. Maximum output shaft speed shall be interpreted as the output shaft maximum speed limit or the output shaft maximum governing speed, whichever occurs first. The engine shall be subjected to a 600 hour endurance test consisting of 100 cycles of six hours each in accordance with the schedule listed below using the oil and fuel specified in 4.8.1.1. Each cycle shall be preceded by a 2 hour shutdown. The test runs in each cycle shall be conducted in the order given. The time for changing power shall be charged to the duration of the lower setting. Inlet fuel pressure and temperature shall be maintained within the operating limits of this AQP throughout the test. During five successive cycles prior to the tenth cycle and five successive cycles after the 90th cycle of the endurance test, the temperature of the fuel shall be maintained at the maximum temperature specified in 3.7.3.1.3.3. The fourth cycle and every fifth cycle thereafter shall be accomplished with the anti-icing bleed air system in operation. The amount of bleed air required for continuous engine operation and the maximum permissible customer and anti-icing bleed air flow during the above cycles, including bleed air temperatures and pressures, shall be stated in the test report. For all operations during runs “e” through “k”, the oil temperature shall be maintained at no less than the maximum oil temperature specified in 3.2.1.4.8. The oil pressure shall be adjusted at the beginning of the test to the minimum steady state value specified in 3.2.1.4.8 for the test. No further adjustments shall be permitted during the test except when authorized by the Using Service. Oil consumption shall be determined and reported after each cycle. Samples of oil shall be taken and spectrometric oil analysis performed, after the calibration run and at the completion of each endurance cycle. Analysis and reporting of any one sample shall not lag the actual sampling by more than three cycles. Oil drained for analysis shall not be charged to engine oil consumption and shall be replaced by an equivalent amount of new oil. The engine shall be run with the control system in a control failure mode during runs “a” and “e” of every 5th cycle, for the times designated, to verify engine control system reliability as specified in 3.7.2.2.3. The control failure modes to be evaluated shall be as specified in the pretest data and shall be subject to approval by the Using Service. Accessory pads shall be subjected to rated loads and overhung moments. The actual torque loading and overhung moments imposed during the endurance test shall be stated in the test report. During the test, the exhaust duct, intake duct, and bleed air duct connections shall be loaded as specified in this AQP. If the engine is supplied with an infrared suppression system, it shall operate continuously during runs “e” and “f”. If the engine provides special features such as fuel heaters, indicator lights, and switching functions, these items shall be actuated during selected test runs as specified in the detailed test procedures as approved by the Using Service. If the engine incorporates an engine condition monitoring system, the system shall be in operation throughout the endurance test. The angular misalignment of the power absorber drive shaft to the engine output shaft shall not be less than the maximum allowable angular misalignment specified in 3.1.2.14.1. The engine internal washing provisions shall be demonstrated once every 10 cycles. The procedure for demonstration of the internal washing provisions shall be as specified in the pretest data. At the completion of the endurance runs, the engine oil drain provisions specified in 3.7.7.4.2 shall be demonstrated. This demonstration shall also verify the adequacy of the port provided for obtaining oil samples for spectrometric analysis, and the locations for the magnetic chip

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detectors. At the end of each cycle, special engine control system features (e.g., overspeed, overtemperature, torque matching) shall be cycled throughout their functions. Each cycle shall consist of the following runs:

a. Maximum-Idle Run. This run shall consist of six successive periods of 10 minutes each. Each period shall include 5 minutes at maximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, followed by 5.0 minutes at idle operation condition. If the engine provides for anti-icing, at the end of each period at maximum test temperature setting, anti-icing controls shall be operated for one minute with the maximum anti-icing bleed air, before the power setting is changed. During the 5th cycle and every fifth cycle thereafter the first 3.0 minutes of each 5.0 minute period at maximum test temperature shall be run with the control in the failure mode. Transient data recording systems are to be on when switching the control from the automatic to the failure mode, and also when switching back to the automatic mode from the failure mode.

b. Incremental Torque Run. This run shall consist of 96 minutes including:1. Eight minutes at the maximum output shaft speed and the output shaft

torque available at maximum test temperature setting or at the maximum continuous torque limit, whichever is less.

2. Eight minutes at the minimum output speed and the maximum continuousoutput shaft torque limit.

3. Eight minutes at the maximum output shaft speed and 90 percent of themaximum continuous output shaft torque limit.

4. Eight minutes at the minimum output shaft speed and 90 percent of themaximum continuous output shaft torque limit.

5. Eight minutes at the maximum output shaft speed and 80 percent of themaximum continuous output shaft torque limit.

6. Eight minutes at the minimum output shaft speed and 80 percent of themaximum continuous output shaft torque limit.

7. Four minutes at the maximum output shaft speed and 60 percent of themaximum continuous output shaft torque limit.

8. Four minutes at the minimum output shaft speed and 60 percent of themaximum continuous output shaft torque limit.

9. Four minutes at the maximum output shaft speed and 40 percent of themaximum continuous output shaft torque limit.

10. Four minutes at the minimum output shaft speed and 40 percent of themaximum continuous output shaft torque limit.

11. Four minutes at the maximum output shaft speed and 20 percent of themaximum continuous output shaft torque limit.

12. Four minutes at the minimum output shaft speed and 20 percent of themaximum continuous output shaft torque limit.

13. Four minutes at the maximum output shaft speed and 10 percent of themaximum continuous output shaft torque limit.

14. Four minutes at the minimum output shaft speed and 10 percent of themaximum continuous output shaft torque limit.

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15. Four minutes at the maximum output shaft speed and at zero output shafttorque.

16. Four minutes at the minimum output shaft speed and at zero output shafttorque.

17. Four minutes at idle and at zero output shaft torque.18. Four minutes at idle and at zero output shaft speed.

If the test data indicates the existence of critical compressor or turbine vibration conditions within the operating speed range of the engine between idle and maximum test temperature setting, at the option of the Government, the following shall be substituted for 48 minutes of the incremental torque run of each cycle to be chosen by the Government.

1. Sixteen minutes at gas generator rotor and output shaft speeds whichproduce the critical vibration conditions.

2. Eight minutes at gas generator rotor and output shaft speeds 2.0 percentmore than those which produce the critical vibration conditions.

3. Eight minutes at a gas generator rotor speed 2.0 percent more and anoutput shaft speed 2.0 percent less than those which produce the critical vibration conditions.

4. Eight minutes at gas generator rotor and output shaft speeds 2.0 percentless than those which produce the critical vibration conditions.

5. Eight minutes at gas generator rotor speed 2.0 percent less and outputshaft speed 2.0 percent more than those which produce the critical vibration conditions.

c. Power Transient Run. This run shall consist of 39 minutes of power transients.1. Four minutes at no load condition followed by one minute at maximum test

temperature setting at the output shaft speed specified for the maximum rating in Table IA. Repeat the above for a total of 20 minutes.

2. The remaining 19 minutes of the run shall consist of one minute atmaximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, followed by an immediate decrease to no load condition. As soon as the engine reaches the no load condition the engine power shall be increased to the temperature and output shaft speed associated with the previous condition and maintained at this power level for a period of one minute before repeating the cycle.

d. Intermediate Test Temperature Setting Run. This run shall consist of 9 minutes ofoperation in the sequence of condition and time duration as follows:

One minute idle, 3 minutes intermediate test temperature setting, one minute at maximum test temperature setting at the output shaft speed specified for the maximum rating in Table IA, 3 minutes intermediate test temperature setting, and one minute idle. Because the engine does not provide reverse, intermediate test temperature setting has been substituted for maximum reverse in this run.

e. Intermediate Run. This run shall consist of 30 minutes:1. Fifteen minutes at intermediate test temperature setting at the maximum

allowable output shaft speed. 2. Fifteen minutes at intermediate test temperature setting at the maximum

allowable torque.

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During the 5th cycle and every fifth cycle thereafter the first 26 minutes of this run shall be with the control in the failure mode. Transient data recording systems are to be on when switching the control from the automatic to the failure mode and also when switching back to the automatic mode from the failure mode.

f. Maximum Continuous Run. This run shall consist of 20 minutes:1. Ten minutes maximum continuous test temperature setting at maximum

allowable output shaft speed. 1. Ten minutes maximum continuous test temperature setting at the

maximum allowable torque. g. Intermediate Run. This run shall consist of 15 minutes at intermediate test

temperature setting at intermediate rated output shaft speed. h. Maximum Continuous Run. This run shall consist of 15 minutes at maximum

continuous test temperature setting at maximum continuous rated output shaft speed. i. Intermediate-Maximum Run. This run shall consist of six periods of 5 minutes each

alternating between: 1. Five minutes at intermediate test temperature setting at intermediate rated

output shaft speed. 2. Five minutes at maximum test temperature setting at maximum rated

output shaft speed. j. Maximum Continuous Run. This run shall consist of 16 minutes at maximum

continuous test temperature setting at maximum continuous rated output shaft speed. k. Intermediate-Maximum Run. This run shall consist of 30 minutes:

1. Fifteen minutes at intermediate test temperature setting at intermediaterated output shaft speed.

2. Fifteen minutes at maximum test temperature setting at maximum ratedoutput shaft speed. When maximum temperature setting is limited to less than 15 minutes duration, maximum test temperature setting shall be run for that time duration and the remaining time shall be at intermediate test temperature setting. At 5 minute intervals during the run, the anti-icing controls shall be operated for one minute with maximum anti-icing bleed air.

4.8.1.1.3.1 Starts. A minimum of 1200 starts, commencing with the first test cycle of 4.8.1.1.3 shall be made on each endurance test engine. In addition to the 1200 endurance test starts, there shall be 40 false starts (a starting sequence without benefit of light-off followed immediately after the permissible engine draining time by a successful start) and 40 restarts (a start within a maximum of 14 minutes time from shutdown). Starts shall be performed with a starter that is acceptable to the Using Service. The engine shall be started and shutdown not less than six times each endurance test cycle. Of the at least 1200 endurance starts, 440 shall be accomplished following varied regulated shutdown periods. Those starts at the beginning of each endurance cycle shall follow a shutdown period of at least two hours. The shutdown period for 72 starts shall be regulated to provide intervals between starts of 2.5 minutes, 5.0 minutes, 7.5 minutes, and increasing, thereafter, by 2.5 minute increments up to and including 90 minutes for the 72nd start. Each of the 72 regulated shutdown periods shall be preceded by immediate engine shutdown, without being held at idle, after engine operation for a duration of not less than two minutes at intermediate test temperature setting. For 268 starts, the shutdown period shall be regulated to

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provide an interval between starts of no less than 45 minutes. The shutdown period for the remaining at least 760 starts need not be controlled. During all endurance starts, immediately after the engine has reached idle speed, an acceleration to the next scheduled endurance test condition shall be accomplished by selecting the control power signal for the appropriate test temperature setting.

4.8.1.1.3.2 Contingency Power Run. The engine contingency power capability shall be qualified by four periods of operation at contingency power test temperature for each 300 hours of operation. Each operation at contingency power shall be preceded by and followed by a period of operation at maximum continuous power. The duration of each period of operation shall be 3t0.677 where "t" is the rating time for contingency power. All times are in minutes. This will be demonstrated on the endurance test engine after completion of the endurance test.

4.8.1.1.4 Recalibrations.

4.8.1.1.4.1 Engine Recalibration. After completion of the test specified in 4.8.1.1.3 through 4.8.1.1.3.2, recalibration in accordance with 4.8.1.1.2.2 shall be conducted on the endurance test engine. The recalibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature conditions, the values of output shaft power obtained during the initial calibration. The recalibration may be preceded by a specified run during which the cleaning procedure of 3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration. The performance degradation at the completion of 600 hours of endurance testing shall not exceed test plan identified field removal limits.

4.8.1.1.4.2 Component Recalibration. Functional bench calibrations/checks shall be in accordance with 4.8.1.1.2.1.

4.8.1.1.5 Engine Disassembly and Inspection. Each engine completing the endurance test shall be completely disassembled for examination of all parts. Prior to cleaning, the engine parts shall be given a “dirty inspection” for evidence of leakage, oil coking, unusual heat patterns, and abnormal conditions. The “dirty inspection” shall be completed before any parts are cleaned. The engine parts shall then be cleaned and a “clean inspection” shall then be performed. Engine part measurements shall be taken as necessary to determine excessive wear and distortion. These measurements shall be compared with the engine manufacturer’s drawing dimensions and tolerances and with similar measurements made prior to the test. Inspection techniques may also include but not be limited to: magnetic particle, fluorescent penetrate, X-ray, and ultrasonic. During the “clean inspection” a visual examination and condition assessment shall be conducted. Upon completion of “clean inspection,” the Using Service shall be provided all results of nondestructive tests and recommendations for modification or redesign of deficient parts. The Using Service shall be notified of the inspection commencement date prior to each inspection. The following data shall be made available to the Using Service during both inspections:

a. Inspection forms filled out by the contractor listing all observed deficiencies.b. Tabulation of all parts found deficient.c. Detailed configuration list of the component or system tested.d. Test logs and list of test events.

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e. Spectrometric oil analysis report.f. Other data as required.

4.8.1.1.6 Endurance Test Completion. The endurance test will be considered to be satisfactorily completed when the engine has completed the endurance test of 4.8.1 and during the final engine recalibration, performance degradation at the completion of 600 hours of endurance testing shall not exceed test plan identified field removal limits; the engine meets all other specified performance requirements which can be checked by the calibration procedure; the test engine and components are operating satisfactorily at the end of the tests, and teardown inspections do not disclose part failures or impending failures. The Using Service shall be provided all results of nondestructive tests and recommendations for modification or redesign of deficient parts shall be presented by the engine contractor.

4.8.2 Accelerated Mission Test (AMT). The accelerated mission test shall consist of two (2) 1200 hour accelerated mission tests. Each AMT shall simulate an engine life equivalent of 6000 hours defined in paragraph 3.3.8.1. During AMT, the engine shall be operated, maintained, and inspected IAW approved field maintenance procedures. Engine performance deterioration shall not exceed the difference between the minimum production engine performance and the approved field removal performance limit identified in the pretest data.

4.8.2.1 Accelerated Mission Test. The accelerated mission test shall consist of 1200 hours operation on an engine IAW the following test schedules, calibrations, and procedures. The engine shall be tested using one of the fuels and oils as specified in 3.7.3.2.1 and 3.7.7.2.1.

4.8.2.1.1 Pretest Verification.

4.8.2.1.1.1 Engine Dry Weight. Prior to initiation of the calibration, the dry weight of the engine as specified in 3.2.2.1 shall be verified in accordance with 4.3.5.2.

4.8.2.1.2 Calibrations (AMT).

4.8.2.1.2.1 Component Calibration. Functional bench calibration/checks shall be conducted on each of the components specified in Table XVIII. Components requiring fuel shall undergo bench calibrations using fluid in accordance with 3.7.3.3.1. Components shall conform to the design tolerance range required by the applicable design specifications. Temperature sensing system performance shall meet the tolerance and thermal response characteristics specified in 3.7.6.4 of this AQP.

4.8.2.1.2.2 Engine Calibration. The procedure during the engine calibration shall be such as to establish the performance characteristics of the complete engine. Prior to the beginning of the calibration, the engine shall be cleaned using the wash procedure specified in 3.7.13, and all engine controls shall be adjusted and shall not be readjusted throughout the calibration. Calibrations shall be made initially with no customer accessory power extraction and no bleed air extraction other than that required for continuous engine operation. During calibration, engine inlet air shall be controlled to the

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temperature specified for the Table IA, IB and IC engine rating. The fuel and oil used shall be the same as those used during 4.8.2.1.3. Data indicated for calibration in 4.3.5.4 and 4.3.5.5 shall be recorded. During calibration, conformance with 3.3.6.4 and 3.3.6.5 shall be demonstrated. The following data shall be obtained:

a. Data required to establish compliance with the sea level performance ratings in Table IA, IB and IC of the engine specification and to establish the accuracy of the torque sensor signal.

b. Data required establishing compliance with 3.2.1.5.6 and 3.7.9.

4.8.2.1.3 Accelerated Mission Test Procedure.Following the calibration run, the engine shall be adjusted to permit operation at the maximum allowable steady state gas temperature at the first stage turbine rotor inlet as defined in 3.2.1.4.5. This value shall be reestablished at the beginning of each 100 cycles. The number of adjustments required shall be recorded. The engine inlet air shall be at the prevailing ambient conditions of the test site. For the purpose of the accelerated mission test, operation at maximum continuous through maximum test temperature settings shall be defined as operation at a gas temperature at or above the maximum allowable (min endurance temperature) gas generator first stage turbine rotor inlet temperature specified in Table VI for the maximum continuous through maximum power settings. The engine shall operate at the MGT that corresponds to the first stage turbine rotor inlet temperature (TRIT). The relationship between MGT and TRIT for these settings shall be established via a 4-point performance calibration, prior to the test and checked at the conclusion of every 50th cycle (±2 cycles). For the purpose of the accelerated mission test, minimum output shaft speed shall be interpreted as that output shaft speed at which the output shaft torque is not less than the torque limit for the applicable power condition or the minimum output shaft governing speed, whichever occurs first. Maximum output shaft speed shall be interpreted as the output shaft maximum speed limit or the output shaft maximum governing speed, whichever occurs first. The engine shall be subjected to a 1200 hour accelerated mission test consisting of 3000 cycles of 24 minutes each in accordance with the schedule listed below using the oil and fuel specified in 4.8.2.1. The test runs in each cycle shall be conducted in the order given. The time for changing power shall be charged to the duration of the lower setting.

ACCELERATED MISSION TESTING DEFINITION (AMT) HELICOPTER Power Level TIME _____ (min) Start + Idle 1.5 MRP 1.0 IRP 2.0 MCP 1.0 35% IRP 0.5 MCP 1.0 35% IRP 0.5 MCP 1.0 No Load 0.5 IRP 1.0 Idle 1.0

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Shutdown 1.0 Start and Idle 1.5 MRP 1.0 IRP 2.0 MCP 1.0 35% IRP 0.5 MCP 1.0 35% IRP 0.5 MCP 1.0 27% IRP 0.5 IRP 1.0 Idle 1.0 Shutdown 1.0

Total 24.0 min

Inlet fuel temperature and pressure shall be maintained within the operating limits of 3.7.3.3.4 throughout the test. For all operations the oil temperature shall be maintained within the operating limits of 3.2.1.4.8. Oil consumption shall be determined and reported after each 100 cycles. Samples of oil shall be taken and spectrometric oil analysis performed, after the calibration run and at the completion of each 500 accelerated mission test cycles. Oil drained for analysis shall not be charged to engine oil consumption and shall be replaced by an equivalent amount of new oil. The actual accessory pads torque loading and overhung moments imposed during the accelerated mission test shall be stated in the pretest report. During the test, the exhaust duct, intake duct, and bleed air duct connections shall be loaded as specified in the pretest data.

4.8.2.1.4 Recalibrations.

4.8.2.1.4.1 Engine Recalibration. After completion of the tests specified in 4.8.2.1.3 recalibration in accordance with 4.8.2.1.2.2 shall be conducted on each accelerated mission test engine. The recalibration shall be conducted with the engine adjusted to produce, under the rated inlet temperature conditions, the values of output shaft power obtained during the initial calibration. The recalibration may be preceded by a specified run during which the cleaning procedure of 3.7.13 may be applied. The fuel and oil used shall be the same as those used during the initial calibration.

4.8.2.1.4.2 Component Recalibration. Functional bench calibrations/checks shall be in accordance with 4.8.2.1.2.1.

4.8.2.1.5 Engine Disassembly and Inspection. Each engine completing the accelerated mission test shall be completely disassembled for examination of all parts. Prior to cleaning, the engine parts shall be given a “dirty inspection” for evidence of leakage, oil coking, unusual heat patterns, and abnormal conditions. The “dirty inspection” shall be completed before any parts are cleaned. The engine parts shall then be cleaned and a “clean inspection” shall then be performed. Engine part measurements shall be taken as necessary to determine excessive wear and distortion. These measurements shall be compared with the engine manufacturer’s drawing dimensions and tolerances and with similar measurements made prior to the test. Inspection techniques may also include but not be limited

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to: magnetic particle, fluorescent penetrate, X-ray, and ultrasonic. During the “clean inspection” a visual examination and condition assessment shall be conducted. Upon completion of “clean inspection,” the Using Service shall be provided all results of nondestructive tests and recommendations for modification or redesign of deficient parts. The Using Service shall be notified of the inspection commencement date prior to each inspection. The following data shall be made available to the Using Service during both inspections:

a. Inspection forms filled out by the contractor listing all observed deficiencies.b. Tabulation of all parts found deficient.c. Detailed configuration list of the component or system tested.d. Test logs and list of test events.e. Spectrometric oil analysis report.f. Other data as required.

4.8.2.1.6 Accelerated Mission Test Completion. The accelerated mission test will be considered to be satisfactorily completed when the engine has completed the accelerated mission test of 4.8.2.1 and during the final engine recalibration, performance degradation at the completion of 1200 hours of accelerated mission testing shall not exceed test plan identified field removal limits; the engine meets all other specified performance requirements which can be checked by the calibration procedure; the test engine and components are operating satisfactorily at the end of the tests, and teardown inspections do not disclose part failures or impending failures. The Using Service shall be provided all results of nondestructive tests and recommendations for modification or redesign of deficient parts shall be presented by the engine contractor.

4.8.3 Reliability Evaluation Testing. The requirements a paragraph 3.1.4 shall be demonstrated using paragraphs 5.4.1 through 5.4.2.5 of MIL-HDBK-781A as a guide.

4.9 Engine Development Special Tests. The following special tests shall be accomplished during the PFR and QT phase of the development program.

4.9.1 PFR Phase.

4.9.1.1 1000 Cycle Engine Low Cycle Fatigue Test. A low cycle fatigue test of 1000 cycles shall be performed in accordance with paragraph 4.6.6.2.2 of this AQP.

4.9.1.2 Preliminary Icing Test. Anti-icing tests shall be conducted with the inlet particle separator installed. The tests shall be conducted within the guidelines of 4.6.4.2 of this AQP, but shall be conducted as a gas generator test.

4.9.1.3 150 Hour Benchmark Endurance. An endurance test of 150 hours shall be performed within the guidelines of 4.5.1 of this AQP.

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4.9.1.4 Fine Sand Ingestion Tests. A fine sand ingestion test shall be performed in accordance with paragraph 4.6.4.7.1 of this AQP.

4.9.1.5 Not Used.

4.9.1.6 Highly Accelerated Life Test (HALT). All components identified in Table XVIII, unless specifically exempted by the Using Service, shall undergo HALT. The component shall be exposed to increasing stresses, applied in step increments, of temperature, vibration, and voltage, individually and combined, until the operating and destruct limits of the component (i.e., the fundamental limit of the technology) are determined. Design weaknesses shall be identified and corrected, in concurrence with the Using Service, to the maximum extent practicable. Efficacy of all design changes shall be verified in a repeat HALT.

4.9.1.7 Inlet Thermal Distortion Test. An engine inlet thermal distortion test shall be conducted to establish distortion sensitivity. A rapid symmetrical temperature rise shall be imposed on the engine inlet covering a 60o segment. The Temperature ramp rate shall be IAW 3.1.2.10.4 for 0.15 seconds while operating at idle, 50 percent maximum continuous power, maximum continuous power, and intermediate power to identify surge, sustained power performance and any detrimental effects.

4.9.1.8 Inflow Bleed Test. The relationship between the compressor surge line and operating line in steady state conditions shall be established using an inflow bleed technique. The following steady state conditions shall be performed; four engine corrected speeds corresponding to idle, 50 percent maximum continuous power, maximum continuous power, and intermediate power at standard conditions for sea level, one half absolute altitude and absolute altitude.

4.9.1.9 Engine Back Pressure Test. An engine back pressure test shall be performed to determine sensitivity to exhaust back pressure. The test shall be performed by restricting the engine exhaust flow and evaluating the impact at various back pressures while operating at the steady state conditions of 4.9.1.7.

4.9.1.11 Not used.

4.9.1.12 Not used.

4.9.1.13 Fuel System Suction Test. Prior to completion of EIT testing, a fuel system suction test shall be conducted. This test shall include a simulation of temperature soak back on engine starting, and operation at the extremes of fuel and air temperature and altitude environment of the engine fuel system. Fuel system performance parameters shall be demonstrated with the designated primary fuel that is most critical for that parameter. Performance parameters that are critical due to vapor and air evolution shall be demonstrated with MIL-PRF-5624 grade JP-8 fuel and MIL-PRF-23699 oil. The test shall be conducted within the guidelines established in (4.5.3). Demonstration test points shall include: (1) sea level, 68°C (154°F) fuel temperature, and 70°C (158°F) surrounding air temperature; (2) 4000 ft, 68°C (154°F) fuel temperature, and 70°C (158°F) surrounding air temperature; and 20000 ft, 55°C (131°F) fuel temperature, and 38°C (100°F) surrounding air temperature. Soakback testing shall be accomplished on the first two test conditions. For

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soakback testing, the engine shall be operated for 30 minutes before shutdown. A restart shall be attempted after soak periods of 15 and 30 minutes.

4.9.2 QT Phase.

4.9.2.1 300 Hr Preliminary Endurance Test. Benchmark 300 hour endurance shall be accomplished, on a single engine build, prior to the initiation of the official QT endurance test. This test shall be performed within the guidelines of 4.6.1 of this AQP.

4.9.2.2 Preliminary Low Cycle Fatigue (LCF) Test. An LCF test of 3750 cycles shall be performed on at least one engine prior to the initiation of the official QT LCF test. This test shall be performed within the guidelines of 4.6.6.2.2 of this AQP.

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5. PACKAGING.

5.1 Packaging. For acquisition purposes, the packaging requirements shall be as specified in the contract or order (see 6.2). When packaging of materiel is to be performed by DoD or in-house contractor personnel, these personnel need to contact the responsible packaging activity to ascertain packaging requirements. Packaging requirements are maintained by the Inventory Control Point’s packaging activities within the Military Service or Defense Agency, or within the military service’s system commands. Packaging data retrieval is available from the managing Military Department’s or Defense Agency’s automated packaging files, CD-ROM products, or by contacting the responsible packaging activity.

6. NOTES.

6.1 Intended Use. Engines covered by this AQP are intended for military unique air vehicle propulsion.

6.2 Acquisition Requirements. Acquisition documents shall specify the following; the Title, number, and date of this AQP.

6.3 Definitions. Absolute Altitude. The maximum altitude at which the engine will function satisfactorily for the range of Mach numbers specified in the engine specification. Acceptance Test (AT). The acceptance tests are those tests conducted on engines submitted for acceptance under contract to demonstrate correct assembly and performance to the extent specified in the engine specification. Acceptance Test Procedure (ATP). The procedure through which engine components demonstrate correct assembly and performance in accordance with their purchase or design specifications. Accessories. Items of engine- or airframe-mounted equipment not furnished by the engine contractor, which are required for aircraft operation or as auxiliaries for engine operation. Additional Equipment. Any item shipped with the engine which is neither an accessory nor a component. Air Starting. Air starting is starting in flight under a specific range of airspeed, altitude, air temperature and temperature soak conditions, and is obtained using the starting method (windmilling and/or starter assist) and procedure specified in the engine specification. Analysis. To study or review the nature and relationship of the component, environment, strength or any other aspect of the design. Article. An individual item or particular unit that may be a component (e.g., fuel pump) or a system when considered as a whole (e.g., engine). Bodie Transient. A transient in which the engine power is commanded from maximum to idle, then as the engine is decelerating, the engine power is commanded back to maximum. Capture Area. Capture area is the projected physical area of the aircraft inlet. Case. A case is any static member of the engine that carries a load or acts as a pressure vessel.

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Catastrophic Engine Failure. A failure which results in engine stoppage and extensive damage to the engine. This is distinguished from those failures which cause only a partial degradation of capability or a gradual degradation over an extended period of time. Center of Gravity. A point at which the dry engine mass may be assumed to be concentrated. Class A Incident. An accident that involves $1 million in damages, a fatality, or a destroyed aircraft. Cold Atmosphere Conditions. Cold atmosphere conditions are specified on Figure 24. Cold Engine (for starting only). A ‘‘cold” engine is defined as one which has been allowed to windmill or “motor” (rollover on the starter) until the low pressure turbine exit gas temperature is within 38°C (100°F) of the engine compressor inlet temperature before a start is attempted. Cold Parts. Those parts not listed as hot parts. Combined Environmental Reliability Test (CERT). A test that is the simultaneous application of temperature, vibration, humidity and altitude cycling per defined aircraft mission profiles Combustor Inlet. The exit plane of the diffuser. Combustor Pattern Factor. T4max – T4avg

T4avg – T3avg Combustor Total Pressure Loss. (P3 – P4) 100 percent

P3 Component Test. A test of engine subsystem components. This includes components such as fuel or lubrication pumps as well as cases, gearboxes, disks, and blades. Compressor Efficiency. The ratio of the work required in an isentropic compression process to that required in the actual process. Compressor Inlet. The inlet plane of the first stage compressor rotor. Compressor Inlet Area. The flowpath area at the station defined by the first stage rotor airfoil root leading edge (or defined by the contractor herein). Contingency Power. An operating condition at which the engine is capable of operating for at least an incremental time duration of 2.5 minutes. This operating condition is available for use only under one engine inoperative (OEI) conditions. Contingency Test Temperature Setting. A condition at which the engine is operated during the EIT and QT contingency test to assure an operational capability for at least an incremental time duration specified in this AQP for contingency power. Continuous Ignition. A low energy level, low spark rate (relative to the system used to ignite the combustor during an engine start) ignition system that is constantly fired regardless of operating condition or power setting. No pilot action is needed to activate the system. Control Limiting Temperature. The maximum value built into the engine control to limit the operating value at a commanded power setting and flight condition. Critical Design Review (CDR). The transition point between detailed design and fabrication of a configuration item or aggregate of configuration items. The primary focus is on the completed detailed design documentation and draft production specifications. The detail design on the hardware drawings is reviewed before the contractor manufactures actual test items. Critical Size. The crack size where unstable growth occurs.

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Customer Bleed Air. Air which may be bled from the engine for use in any system external to the engine and is that quantity available over and above the bleed air needed for engine acceleration, engine anti-icing and any other engine system requirements. Damage Tolerance. The ability of the engine to resist failure due to the presence of flaws, cracks or other damage for a specified period of uninspected usage. Damage Tolerant. Resistant to failure due to the presence of flaws, cracks or other damage for a specified period of uninspected usage. Output Shaft Power. The power output at the output shaft(s) as measured by output torque and output shaft speed. Demonstration. The functional capabilities of a component or procedure are reviewed. Depot, Intermediate, or Base Level Inspectable Structure. A structure is depot, intermediate, or base level inspectable if the nature and extent of damage can be detected utilizing one or more selected NDI procedures. Derating. Using an item in such a manner that the stresses or temperatures applied during operation are lower than the stresses and temperatures the item was designed to withstand. Design Service Life. The life in cycles or hours defined by the Using Service (used for design purposes) that the engine or engine component is expected to attain during operational service before becoming uneconomic to repair. Development Test. A non-qualification test used to show risk reduction or proof of concept. Diagram. A diagram is a sketch or outline giving only interface or mounting information for a part or part of an assembly without all the details of the item shown (i.e., diagrams show bolt hole patterns, flanges, electrical plug sockets, etc.). A diagram shows the relationship between the parts of a whole assembly. It does not show many details. Dissimilar Control Technology. Other than digital electronic technology, e.g. hydromechanical, fluidic, analog. Documentation. Any media that provides a record of the design or design process that can be reviewed. Domestic Object Damage. Domestic object damage is caused by engine parts coming loose and striking the engine. See foreign object damage (FOD). Drawing. A drawing is a document that delineates a part or assembly showing major dimensions, clearances, and details of a part or assembly. It can show every detail needed to manufacture the part, assembly, or engine. Materials of composition are shown on the drawing as a related document. Dry Weight. Dry weight is the combined physical weight of the engine and its components with no liquids in the system. It does not include the weight of accessories or additional equipment. Durable. Resistant to cracking, corrosion, deterioration, thermal degradation, delamination, wear, etc., for a specified period of time. Durability. The ability of the engine to resist cracking, corrosion, deterioration, thermal degradation, delamination, wear, etc., for a specified period of time. Durability Critical. Failure results in a significant economic impact to the system but will not necessarily impair flight safety or mission capability.

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Appendix D

Durability Non-Critical. Failure results in a minor economic impact to the system but will not impair flight safety or mission capability. Dynamic Response. Response due to forced functions. Forced functions include: surge, stall, flutter, and resonance. Emergency Fuel. Fuel which significantly limits the overhaul life of the engine and imposes operational restrictions on the aircraft as well. Emergency Power Capability. The engine capability to run higher than contingency power where the engine may fail but the failure will be in a non-catastrophic manner. Any use of this capability shall result in engine removal. Engine. The complete propulsion system (including all parts, components, and expendables) delivered by the engine contractor. Engine Axes. Engine axes are geometric reference lines passing through the engine in the longitudinal, lateral, and vertical directions. (See Engine Coordinate System) Engine Component. Items of equipment, furnished as part of and qualified with the engine, whose size, conformation, and dynamic and static characteristics are essential to attain the engine performance specified in the engine specification. Fuel pumps, engine controls, variable guide vane actuators, anti-icing valves, and the temperature sensing system or devices are included in this category. Components may require separate qualification, calibration, acceptance test or adjustment. Engine Coordinate System. The engine coordinate system is a position indicating system employing the X, Y, and Z-axes. The origin of the system shall be defined in the engine specification. When looking from the rear of the engine, the positive directions shall be as follows: X-axis – forward from the origin, Y-axis – to the left of the origin, and Z-axis – above the origin.Engine Frontal Area. The area based on the front view projection of the engine including engine mounted components. Engine Inlet Area. The area in a plane perpendicular to the airflow path at the junction of the aircraft intake duct and the engine front face. Engine Integrity Testing (EIT). This milestone establishes the acceptability of the engine to continue on into full Preliminary Flight Rating testing. The engine is not required to meet full verification requirements for durability and reliability but is required to be the final production configuration. This milestone is insufficient to authorize production or material release. For reference, the analyses, demonstrations and tests required for EIT are listed in Appendix A. Engine Part. A piece, or two or more pieces joined together, that are not normally subject to disassembly without destruction of the designed use. Engine System Level Ground Test. A complete engine test in either a sea level open air cell or a test cell that can be conditioned to simulate altitude conditions. Environmental Conditions and Operating Envelope of the Engine. Environmental conditions and operating envelope of the engine includes all extremes and limits such as externally applied loads, attitudes, and environmental extremes independently and concurrently in all combinations within the scope of the engine specification. Exhaust Nozzle Effective Area. That area at the exit plane of the nozzle required to pass the engine flow with the ideal velocity and density calculated at the exit plane. This definition accounts for the boundary layer along the outer wall of the nozzle, the very low flow or ineffective

207

Appendix D

area behind the center body, the expansion of metal parts when raised to operating temperature and the actual velocity and density profiles that exist at the exit plane. Fail Degraded. The ability to continue to operate the engine after any single or multiple failure(s) and retain a defined minimum operational. Fail Operational. The ability to continue to operate the engine after any single or multiple failure(s) and still be capable of meeting all defined mission and performance requirements. Fail Safe. The ability to continue to operate the engine safely or safely shut down following the failure or partial failure of one or more elements. False Start. A starting sequence without benefit of light-off, followed immediately after the permissible engine draining time by a successful start. Fatigue Crack Growth Life. The predicted typical propagation life required to grow an assumed initial flaw to its critical size. Flaw Growth Interval. The predicted typical propagation life required to grow an assumed initial flaw to its critical size. Flight Safety Part. Any part, assembly, or installation containing a critical characteristic whose failure, malfunction, or absence could cause an uncommanded engine shutdown, and/or a catastrophic engine failure resulting in loss or serious damage to an aircraft and/or serious injury or death to the occupants. Flight Test. The engine is installed in an aircraft that is then tested in flight to acquire verification data not obtainable through engine system level ground test. Foreign Object Damage. Foreign object damage (FOD) is caused by non-engine parts striking the engine. See domestic object damage. Fracture Critical. Failure will result in probable loss of the aircraft or degradation in mission capability. There are two categories under fracture critical: safety critical and mission critical. Frequency Response. The steady state output of the system to input sinusoids of varying frequency. The output for a linear system can be completely described in terms of the amplitude ratio of the output sinusoid to the input sinusoid. The amplitude ratio, gain, and phase are functions of the frequency of the input sinusoid. Fully Loaded Accessory Pads. Fully loaded accessory pads occur when each accessory pad is subject to the maximum rated torque for that pad. Gas Temperature. The actual steady state temperature of the gas averaged over the gas path at the engine station designated. Gas Temperature Margin. Gas temperature margin is the difference between the maximum allowable temperature defined in 3.2.1.4.5 and the highest first stage turbine rotor inlet temperature which can occur within the engine operating envelope at the highest engine rating. Gulping Volume. Gulping volume is the difference between oil reservoir volumes: (1) with the oil at 15°C (59°F) and the engine at zero speed; and (2) with the engine at stabilized maximum continuous speed. The gulping volume represents the initial amount of oil required to fill the lubrication system lines, pumps, sumps, bearing cavities, etc., each time the engine is started. Highly Accelerated Life Test (HALT). A development test that increasingly stresses a component in step increments of temperature, vibration, and voltage, individually and combined, until the operating and destruct limits of the component (i.e., the fundamental limit of the technology) are

208

Appendix D

uncovered. Design weaknesses are then rectified and the test repeated until all weaknesses are eliminated, to the maximum extent practicable. Highly Accelerated Stress Screen (HASS). A production unit test that increasingly stresses a component in step increments of temperature, vibration, and voltage, individually and combined, in order to precipitate workmanship and component infant mortality failures. The applied stress levels are determined based on data gathered during the development HALT Hot Atmospheric Conditions. Hot atmospheric conditions are specified on Figure 24. Hot engine (for starting only). A ‘‘hot” engine is defined as one where a start is attempted within 10 seconds after a flameout or shutdown. Hot Parts. Parts exposed to the hot gas stream. Idle. See 3.2.1.5.4.1. Impending Failure. The condition of a part where loss of performance or function would take place if the part was continued to be used for its part life required by the engine specification. In-Service Noninspectable Structure. Structure is in-service noninspectable if either damage size, accessibility, or maintenance requirements preclude inspection. Inspection. The physical examination of an article, item, drawing, or part. Installed Configuration. When the engine is tested in the air vehicle, this is known as the installed configuration. Intermediate Power. An operating condition at which the engine is capable of operating for at least an incremental time duration of 30 minutes. Intermediate Test Temperature Setting. A condition at which the engine is operated during EIT, PFR and QT testing to assure an operational capability for at least an incremental time duration of 30 minutes. Latent Failure. Any undetected condition which prevents the part or component from performing its intended function. This includes undetected conditions between missions and during missions. Limit Load. The maximum load expected to be encountered when operated for the design service life at design usage conditions. The factor of safety associated with this load is defined as the limit load. Line Replaceable Unit. Engine components or accessories which are designed for replacement at the lowest level of maintenance without removal of other components or accessories. Load Burst. A rapid application of load to the output shaft. Load Chop. A rapid unloading of load to the output shaft Maximum Allowable Speed. As specified for either steady state or transient conditions, the limit beyond which operation of the engine is not allowed. Maximum Allowable Gas Temperature. The maximum allowable gas temperature, as specified for either steady state or transient condition, is the limit beyond which operation of the engine is not allowed. Maximum Allowable Torque. As specified for either steady state or transient conditions, the torque limit beyond which operation of the engine is not allowed.

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Appendix D

Maximum Continuous Power. An operating condition at which the engine is capable of operating continuously. Maximum Continuous Temperature Setting. A condition at which the engine is operated during EIT, PFR and QT testing to assure a continuous operational capability. Maximum Fuel Consumption Guaranteed. The fuel consumption which has been calculated by a thermodynamic cycle analysis and increased by a production margin factor. Maximum Operating Temperature. The highest temperature the engine can obtain in normal operation at a commanded power setting considering all effects due to engine-to-engine variation, deterioration, installation factors, control limits and tolerances, environment, and operating point. Maximum Power. An operating condition at which the engine is capable of operating for at least an incremental time duration of 10 minutes. Maximum Test Temperature Setting. A condition at which the engine is operated during EIT, PFR and QT testing to assure an operational capability for at least an incremental time duration specified in this AQP for maximum power. Measured Gas Temperature. The signal after the temperature sensor electrical harness plug of the gas path total temperature at the designated engine station. Minimum Output Shaft Power Guaranteed. The output shaft power which has been calculated by a thermodynamic cycle analysis and decreased by a production margin factor. Minimum Engine. A guaranteed performance level having the lowest engine performance over a time period and environment specified by the Using Service without exceeding the highest SFC and considering all effects due to control system variations, engine-to-engine build variations, deterioration and flight condition, without exceedance of engine pressure, temperature, torque, and speed operating limits. Mission. Period beginning with the start of engine prior to flight and ending with the engine shutdown at the completion of the flight. Specific missions include: familiarization, air combat maneuvers, navigation, and air-to-ground weapons. Mission Critical. Failure will generate a significant operational impact by degrading mission capabilities to the extent of creating an indirect safety impact on the weapon system or results in less than Level II handling qualities. Mission Mix. The frequency each mission is flown in a finite period of time (e.g., 5% for familiarization and 25% for air combat maneuvers). Mission Profile. A representation of a specific mission in terms of flight conditions and usage parameters. New Minimum Engine. Minimum engine at the time of government acceptance. No Load. See 3.2.1.5.4.2. Nominal Performance. The performance associated with the statistically average production engine. Non-Operating Environment. The environment to which the engine is exposed during logistics, storage, maintenance, and transportation activities. Normal temperature. The normal or typical operating temperature of a component. One of the temperature environments used during component missionized testing. (4.11.2.1.3)

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Operating Fluids. The total mass of operating fluids within the engine shall equal the summation of the mass of residual and drained fluids. Operational Capability Release (OCR). The OCR milestone establishes the production engine configuration for unlimited production release. For reference, the analyses, demonstrations required for PFR are listed in Appendix D. OCR Test. The OCR tests are those tests conducted on engines to demonstrate by a simulated engine cycle an aircraft mission that is equivalent to one engine life. Operational Life. The life of the engine when exposed to the operational usage. Operational Usage. The usage that the engine is exposed to during actual service operation. Preliminary Design Review (PDR). This review represents the approval to begin detailed design. The primary focus is on the adequacy of top level design documentation for hardware and software configuration items (CI). The PDR is a check to verify that the allocated baseline requirements for the engine, including interfaces, have been addressed and that they can most likely be met or exceeded by the contractor’s proposed functional design.. Preliminary Flight Rating (PFR). This milestone establishes the acceptability of the engine to power the aircraft throughout its full envelope. The engine is not required to meet full verification requirements for durability and reliability and is not required to be the final production configuration. For reference, the analyses, demonstrations required for PFR are listed in Appendix B. Primary Fuels. Fuels on which the engine is designed to operate continuously without restrictions. Process Quality Control. The contractor has established manufacturing procedures that ensure repeatability within specific process limits. Production Margin. A factor applied to fuel consumption, output shaft power and gas generator speed which accounts for performance variations associated with the production of a large number of engines. Qualification Test (QT) Rating. This milestone establishes the acceptability of the engine for low rate production release. It is the sum of analysis, demonstration and test activity in accordance with 4.6 accomplished on engines and components submitted for qualification to demonstrate the suitability of an engine model for production and service use. For reference the analyses, demonstrations and tests required for QT are listed in Appendix C. Rated Output Shaft Power. The value of power output for a given rating when the engine rating temperature, fuel consumption, and speed are not exceeded. The power is that specified in Tables IA, IB, IC, IIA, IIB, and IIC of the engine specification for each of the various rating points. Rated Output Shaft Speed. The speed of the output shaft when the engine is delivering rated torque while operating at a Table IA, IB, IC, IIA, IIB, or IIC power rating. Rated Gas Generator Speed. The maximum allowable speed permitted to attain a given Table IA, IB, IC IIA, IIB, or IIC rating at or below rated gas temperature. Rated Gas Temperature. The maximum allowable turbine rotor inlet temperature (TRIT) permitted to attain a given Table IA, IB, IC IIA, IIB, or IIC rating, at or below the rated gas generator speed. Rated Torque. Rated torque is the output shaft torque at rated output shaft speed when the engine is operating at a Table IA, IB, IC, IIA, IIB or IIC power rating.

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Rating. A rating is a performance characteristic value specified in Table IA, IB, IC, IIA, IIB or IIC of the engine specification. Recovery time. The period it takes the engine to return to stable operation after a dynamic event. Residual Strength. The minimum load carrying capability of the part, minus the maximum applied load, at any time during the design service life. Residual strength accounts for the presence of damage and the growth of damage as a function of design usage and must maintain limit load capability. Restart. A start within a maximum of 14 minutes time from shutdown. Restricted Fuel. Fuel that imposes operational restrictions on the aircraft when used. Safety Critical. Failure will result in the probable loss of the aircraft or hazard to personnel due to direct part failure or by causing other progressive part failures. Satisfactorily. The words ‘‘satisfactorily” or ‘‘satisfactory” as used in this AQP in conjunction with words or terms relating to operation or performance of the engine described in this model specification shall mean: Under the condition specified, throughout the entire operating envelope, the engine operating characteristics and performance are not affected, and the operating and physical limits shown in the specification are not exceeded and no permanent deformation or other damage to the engine occurs. Shaft Power Absorber. Any device or load transferring mechanism which absorbs engine output shaft torque and speed. Dynamometers, propellers and helicopter rotor systems are included in this category. Similarity Analysis. A study or review an existing part or component (or a component which can be shown to be similar) to determine if a new application or environment for the component or part is within the design limits. Small Engines. A small engine is an engine with fan inlet corrected flow requirements of 100 pps or less or compressor inlet corrected flow of 20 pps or less. Smoke Puff. A smoke puff is visually detectable unburned fuel droplets that are produced when a combustor or afterburner is initiated (lightoff) or terminated (shutdown). Specific Fuel Consumption. The mass of fuel consumed per unit of time per unit of power. Stall. Stall is an engine compression system flow instability that does not result in the loss of engine control. A condition in the engine compression system where one or more blades or stages is operating with separated flow. Stall Margin. Stall Margin (SM) is defined as follows:

SM = (Pr/Wa) stall – (Pr/Wa) operating line x 100 percent (Pr/Wa) operating line

Pr is the compressor pressure ratio. Wa is the compressor airflow.

Standard Condition. Standard conditions are the values of air temperature and pressure given in the US Standard Atmosphere, 1976 (NOAA-S/T 76-1562). The standard humidity, for the purpose of this AQP, is zero vapor pressure at all altitudes. All heights noted in this AQP shall be geopotential altitudes. Steady state Stability. Steady state stability is the condition at a constant power setting in which power fluctuations do not exceed a stated margin.

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Strength. The ability of an engine or component to withstand limit loading conditions while maintaining operational capability. Surge. Surge is the loss of engine control resulting from an engine compression system flow instability. Surge Margin. Surge margin is the calculated and/or demonstrated tolerance of the engine to adverse operating conditions while maintaining the required steady state and transient performance capability, calculated by:

SM = (Pr/Wa) surge – (Pr/Wa) operating line x 100 percent (Pr/Wa) operating line

Pr is the compressor pressure ratio. Wa is the compressor airflow.

Throughput. The time required for a computer circuit to process the input signal. Transient. Power demand increase and/or decrease. Transient Stability. Transient stability is the condition following a transient in which power fluctuations do not exceed a stated margin. Transparent to the Pilot. Any event the pilot cannot detect. This includes any significant difference of aircraft handling ability or aircraft stability. Turbine Efficiency. The ratio of the work available in the actual expansion process to that available in an isentropic expansion process. Turbine Inlet. The inlet plane at the first stage nozzle vane row of the turbine. Ultimate Load. The maximum load that a material, part, component, or engine must withstand without catastrophic failure. Unusable Oil. The maximum quantity of oil in the engine lubrication system which is not available to meet engine lubrication requirements throughout the operating envelope under the maneuver forces and attitudes specified in engine specification. Unweathered fuel. Fuel with chemical and physical properties that have not been degraded by environmental effects such as temperature and pressure cycles. Usable Life. Life without repair or replacement. Using Service. The Using Service is the service whose model dash number has been assigned to the engine. Windmilling. Rotation of any or all shafts by any external means other than normal starter mechanism. 6.4 Symbols, Subscripts, Abbreviations, and Acronyms. The symbols, subscripts, abbreviations and acronyms used in this AQP are listed below and shall be in accordance with AS 681J and AS 755D.

AGI Armament Gas Ingestion

AMCOM Aviation and Missile Command

AMS Aerospace Material Specification

Appendix D

213

AMT Accelerated Mission Test

ARD Aerospace Resource Document

ARP Aerospace Recommended Practice

AS Aerospace Standard

ASCII American Standard Code for Information Interchange

ASIP Aircraft Structural Integrity Program

ASME American Society of Mechanical Engineers

ASMET Accelerated Simulated Mission Engine Test

ASTM American Society for Testing and Material

BIT Built-in Test

BOCLE Ball-on-Cylinder Lubricity Evaluator

C&E Controls and Externals

CDRL Contractor Data Requirements List

CERT Combined Environment Reliability Test

CG Center of Gravity

CIT Compressor Inlet Temperature

CL Confidence Level

CONUS Continental United States

CSCI Computer Software Configuration Item

cSt Centistokes

dBsm Decibels per sq. meter

DEM/EVAL Demonstration/Evaluation

ECS Environmental Control System

ECU Electronic Control Unit

EEPROM Electrically Erasable Programmable Read Only Memory

EFH Engine Flight Hours

EIA Electronic Industries Association

EIT EM

Engine Integrity Testing Electromagnetic

EMC Electromagnetic Compatibility

EMD Engineering and Manufacturing Development

EME Electromagnetic Environment

Appendix D

214

EMI Electromagnetic Interference

EMP Electromagnetic Pulse

EMS Engine Monitoring System

EMT Elapsed Maintenance Time

EMV Electromagnetic Vulnerability

ENSIP Engine Structural Integrity Program

EOH Engine Operating Hours

EPR Engine Pressure Ratio

ESS Environmental Stress Screening

FADEC Full Authority Digital Engine Control

PFR Preliminary Flight Rating

FMECA Failure Modes, Effects, and Criticality Analysis

FOD Foreign Object Damage

FSD Full Scale Development

FSII Fuel System Icing Inhibitor

GFE Government Furnished Equipment

GGT HALT HASS

Gas Generator Turbine Highly Accelerated Life Test Highly Accelerated Stress Screen

HC Hydrocarbon

HCF High Cycle Fatigue

HEMP High Energy Electromagnetic Pulse

HFE Human Factors Engineering

HPT High Pressure Turbine

IAW In Accordance With

ICD Interface Control Document

IEC International Electrotechnical Commission

IECMS In-Flight Engine Condition Monitoring System

IFSD In-Flight Shut Down

IGV Inlet Guide Vanes

ILS Integrated Logistics Support

IPS Inlet Particle Separator

215

IR

IRP

IRTP

ISA

QT

JOAP

LCC

LCF

LHV

LI

LPT

LRU

LWC

MA

MAU

MECSIP

MEFH

MMH/EFH

MTBF

MTBMA

MTTR

NACA

NAVAIR

NBC

NDI

NOX

OCR

OEI

Appendix D

Infrared

Intermediate Rated Power

Integrated Reliability Test Program

International Standard Atmosphere

Qualification Test

Joint Oil Analysis Program

Life Cycle Cost

Low Cycle Fatigue

Lower Heating Value

Lubricity Improver

Low Pressure Turbine

Line Replaceable Unit

Liquid Water Content

Maintenance Action

Maintenance Actions Unscheduled

Mechanical Equipment Structural Integrity Program

Mean Engine Flight Hours

Maintenance Man Hours/Engine Flight Hour

Mean Time Between Failure

Mean Time Between Maintenance Action

Mean Time to Repair

Nation Advisory Committee for Aeronautics

Naval Air Systems Command

Nuclear, Biological and Chemical

Non-Destructive Inspection

Oxides of Nitrogen including Nitric Oxide plus Nitrogen dioxide

Operational Capability Release

One engine inoperative

OPNAVINST Naval Operations Instruction

PFR Preliminary Flight Rating

PLA Power Lever Angle

PNdB Perceived Noise Level

Appendix D

216

POD Probability of Detection

PPM Parts Per Million

PTO Power Takeoff

QT R&M

Qualification Test Reliability and Maintainability

RAM Random Access Memory

RCS Radar Cross Section

RFP Request for Proposal

RMS Root Mean Square

RVP Reid Vapor Pressure

SAE Society of Automotive Engineers

SFC Specific Fuel Consumption

SHP Shaft Horsepower

SI International System of Units

SLS Sea Level, Static

SOW Statement of Work

SRA Service Replaceable Assembly

SRU Service Replaceable Unit

TAC Total Accumulated Cycles

TBO Time Between Overhauls

TS Turboshaft

USA United States Army

USAF United States Air Force

USC United States Code

USN United States Navy

UUT Unit Under Test

UV Ultraviolet

V/L Vapor-Liquid Ratio

WSD Wear Scar Diameter

6.5 Material Safety Data Sheet. Contracting officers shall identify those activities requiring copies of completed Material Safety Data Sheets prepared in accordance with FED-STD-313. The pertinent Government mailing addresses for submission of data are listed in FED-STD-313.

Appendix D

217

6.6 Subject Term (Key Word) Listing. Engine, aircraft Gas turbine engine, aircraft Propulsion Turboshaft

6.7 Metrication. Not used.

6.8 Engine Specification Preparation. A complete engine specification conforming to the instructions for preparation contained herein shall be prepared and submitted by the contractor in Adobe Acrobat 4.0 or later format for approval by the Using Service.

6.8.1 Instruction for Preparation. The engine specification shall be prepared as follows:

a. The headings and numbering of sections and paragraphs shall correspond tothose of this specification.

b. Paragraphs herein that are applicable as written shall be copied into thespecification. Any change, addition, or deletion shall be identified by placing an identifying mark or symbol following the appropriate paragraph number and headings.

c. Paragraphs herein that are not applicable to the particular engine design shallhave the words "not applicable" entered following the appropriate paragraph number and headings.

d. Paragraphs requiring modification to define a particular design shall be modifiedto the extent necessary to describe the characteristics of that particular engine and model.

e. New requirements or additions shall be added as additional subparagraphs or asnew paragraphs in logical sequence and location.

f. Items such as tables, figures, drawings, diagrams, and appendices, shall bepresented in complete form in the specification. Complete statements calling out these items shall be included in the text of the specification.

g. The prepared specification shall have hypertext links between all major and minorsections, figures, tables, and appendices.

Appendix D

218

TABLE IA. Performance 15°C (59°F), sea level static conditions with nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUMCONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

Appendix D

219

TABLE IB. Performance 32.1°C. (90°F) sea level static conditions with nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUM CONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

Appendix D

220

TABLE IC. Performance 50°C. (122°F) sea level static conditions with nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUM CONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

221

Appendix D

TABLE IIA. Performance 0.61 km, (2000 ft.) 21.1°C (70°F) static conditions with ________nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUM CONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

222

Appendix D

TABLE IIB. Performance 1.22 km, (4000 ft.) 35°C (95°F) static conditions with _______ nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUM CONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

223

Appendix D

TABLE IIC. Performance 1.83 km, (6000 ft.) 35°C (95°F) static conditions with _______ nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUM CONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

224

Appendix D

TABLE IID. Performance 6.1 km (20,014 ft.), 0.4 Mach number standard day with _______ nozzle.

ENGINE RATINGS,

TRIT °C (°F) (1)

MINIMUM OUTPUT SHAFT

POWER, (shp)

OUTPUT TORQUE,

NM (lb-ft)

OUTPUT SHAFT SPEED,

rpm

MAXIMUM SPECIFIC

FUEL CON-SUMPTION,

kg/hr-kW (lb/hr-shp) (6)

MAXIMUM GAS

GENERATOR SPEED

rpm

MAXIMUM MEASURED

TEMP., °C

(°F) (3)

ENGINE AIR

FLOW kg/s

(lb/s) (3)

CONTINGENCY

MAXIMUM

INTERMEDIATE

MAXIMUM CONTINUOUS

75% MAX CONT

50% MAX CONT

25% MAX CONT

NO LOAD (2)

IDLE (4) (2)

Notes: (1) Additional ratings and columns shall be added as required by the Using Service. (2) Fuel consumption, kg/hr (lb/hr) (Max.).(3) Supplied for reference.(4) Speed producing highest output shaft power.(5) Not Used.

(6) Fuel LHV is 42,576 KJ/kg (18,300 BTU/lb).

225

Appendix D

TABLE III. Low Cycle Fatigue Damage Fraction Table (Ref. 3.3.8.6, item E)

Appendix D

226

TABLE IV. Component test report documentation requirements (Ref. 4.3.6.1.2). a. Front end matter in accordance with Table VII, items a through g.b. 1.0 Introduction.

Include text similar to:This report describes the environmental qualification testing of P/N XYZ. A suite comprising six individualenvironmental tests was performed over the period of May through December, 2006. In addition to the sixtests, analyses were prepared as the basis for waiving the performance of two other required tests. The sixindividual test reports are included herein, in sections 1 through 6, and the analyses are referenced bydocument number.

Testing was conducted on a single serial number, 0346PLD12024, in order to demonstrate that thecumulative damage of the qualification environments had no deleterious effects on the UUT’s (unit under test)structural integrity or performance to specified requirements.

A separate unit, S/N 0346PLD12026, underwent a 500-hour SOT (simulated operational test) over the periodJune through August, 2006. The results of this test are also included herein….

c. 2.0 Component Description.2.1 System Description. 2.2 Unit Construction and Assembly. 2.3 Component Non-Conformances.

Any non-conformances shall be described here, with reference to Appendix A for the actual documentation of such. Using Service approval of the non-conformances shall also be mentioned and included in Appendix A.

d. 3.0 Test Plan.Explanation that there is a single test plan, that it specifies the order of testing, includesthe individual test procedures, and the pass/fail criteria, etc.

3.1 Test Plan Summary. Summarize the test plan in a table, including the order of testing, the name of the test, the test procedure, and the test plan paragraph number.

3.2 Plan Approval. Specify when/how Using Service approval of the test plan was provided, and Include in Appendix B the documentation of such. Specify when/how Contractor approval (if required) of the test plan was provided, and include in Appendix B the documentation of such.

3.3 Deviations to the Test Plan - Global. Explain any global deviations (applicable to all tests) that had to be made to the test plan and reference Appendix B for the documentation of the deviations (IFM, CM, etc.), along with Using Service approval.

Appendix D

227

TABLE IV. Component test report documentation requirements - Continued. e. 4.0 Test Readiness Review.

Text indicating that a TRR was held. Provide information on who, when and where, andthe outcome of the TRR.

4.1 Using Service Approval to Test. Include documentation of such.

4.2 Test Readiness Review. Include TRR materials in Appendix C.

f. 5.0 Testing Summary.Include table similar to the one illustrated below.

Test Name UUT S/N

Test Order

Test Plan # Paragraph

Date of Test Result Pre-Test ATP Test Report # Section # Post-Test ATP

Date Appendix Date Appendix SOT

043PDL124 1 Pass

Vibration 2 Pass

Operational Shock 3

Sustained Acceleration 4

Explosive Atmosphere 5

Humidity 6

Sand & Dust N/A By Analysis N/A N/A N/A

Fungus N/A By Analysis N/A N/A N/A

g. Individual test summaries. For each of the tests listed in the table of paragraph f, providein separate test subparagraphs, each beginning at the top of a new page, the followinginformation. These summaries are typically 1 to 3 pages in length each, but may beexceeded as necessary.5.1 Simulated Operational Test.

5.1.1 Significant events that occurred or relevant observations. 5.1.2 Test procedure errors made or deviations taken. 5.1.3 The effect, if any, of 5.1.1 and 5.1.2 on the test outcome or results. 5.1.4 Pass/fail criteria that were not met and the significance of the failure. 5.1.5 Proposed remedies or design changes as a result of failures. 5.1.6 Any other relevant information. 5.1.7 Statement on whether or not the UUT met its qualification test objectives.

Appendix D

228

TABLE IV. Component test report documentation requirements - Continued.

5.2 Vibration. 5.2.1 Significant events that occurred or relevant observations. 5.2.2 Etc.

5.3 Shock. 5.4 Sustained acceleration. 5.5 Explosive atmosphere. 5.6 Humidity. 5.X Next Test.

h. 6.0 Disassembly and Inspection.(1) Each disassembled part or subcomponent shall be discussed in detail as to:

condition with respect to a new part, continued usability, and likelihood ofimpending failure.

(2) Photos of each part or subcomponent must be clear and in color; photocopiedpictures are not acceptable, as they do not provide adequate detail.

i. 7.0 Conclusion.j. Appendix A – Component Non-Conformance.

(1) Includes all material discrepancy actions, plus any other documentation of unitnon-conformance.

(2) Documentation of Using Service approval.k. Appendix B – Test Plan Documentation.

(1) Documentation of Using Service approval.(2) Documentation of contractor approval.

l. Appendix C – Test Readiness Documentation.(1) Documentation of Using Service approval.(2) TRR presentation or other materials.

m. Appendix D – ATPs or Calibrations Conducted During Testing.Appendix D-1 Appendix D-2 Appendix D-n

Appendix D

229

TABLE IV. Component test report documentation requirements - Continued. n. Individual test reports and results. Individual reports shall be included in separate ‘tabbed’

sections. Note that these reports shall not duplicate any of the information provided in athrough m, above, unless absolutely necessary.(1) The report shall be organized to generally follow the approved test procedure.(2) Test equipment and all test setups shall be photographed.(3) All environmental test conditions shall be documented.(4) Detail the conduct of the test, including a chronological history of events and

incidents connected with the test.(5) A complete analysis of results shall be included, as should any pertinent

performance/calibration data for illustration purposes.(6) Clearly highlight (including reference to applicable data sheets) and discuss in

detail any test anomalies, and propose any remedies, if necessary.(7) Tables and charts shall be utilized to clearly present and/or summarize the test

data. Note that an ATP top sheet with the word ‘pass’, or one hour’s worth ofdata from a 300-hour test, or only providing a few of the parameters monitoredas a ‘sample’ is insufficient documentary evidence of test passage and will berejected. Data presented in the example tabular format below has been deemedacceptable in most cases.

Parameter Setpoint Limit Actual Min Value Max Value Mean Std.

Deviation Low Limit High Limit Pass/ Fail

T1 -40C ±1.1C 39.93 -39.74 -40.16 -39.46 0.098 -41.03 -38.83 P 85C ±1.1C 84.53 84.71 84.32 85.23 0.112 83.43 85.63 P

N1 69% ±0.1% 69.11 69.11 69.11 69.11 0.001 69.01 69.21 P 110% ±0.1% 109.76 108.85 108.85 108.86 0.001 108.74 108.94 P

NR 67% ±0.1% 66.94 66.95 66.95 66.95 0.001 66.84 67.04 P 104% ±0.1% 104.02 104.02 104.02 104.03 0.001 103.92 104.12 P

WF 0pph ±11.4pph 0.16 0.38 0.23 0.63 0.076 -11.24 11.56 P 2700pph ±11.4pph 2694.69 2700.52 2700.36 2700.82 0.087 2683.29 2706.09 P

T45 1250F ±10.8F 1282.33 1281.43 1269.99 1296.44 5.938 1271.53 1293.13 F

(8) Component condition at test completion shall be detailed/discussed, andconclusions/recommendations rendered.

Section 1 – Simulated Operational Test. Section 2 – Vibration Test. Section 3 – Operational Shock Test. Section 4 – Sustained Acceleration Test. Section 5 – Explosive Atmosphere Test.

Section 6 – Humidity Test.

Appendix D

230

TABLE V. Gearbox pads and drives.

Accessory or Component

Type of Drive 1/

Ratio Pad to Rotor rpm 2/

Rotation Direction Facing Pad

Torque N-m (lbf-in) Overhung Moment N-m (lbf-in) 4/

Maximum Spline Mis- alignment mm (in)

Steady State transient

Maximum Continuous

Over-load 3/

Static

NOTES: 1/ Give the type of drive including AND or MS number and type.

2/ Ratio of speeds based on 100 percent (of the driving gas generator) rotor speed of _______ RPM. 3/ Specify duration and frequency of overload. 4/ Overhung moment of accessory or component must include max allowable load permissible from attachments for example fuel pump must include aircraft fuel connection load.

Appendix D

231

TABLE VI. Engine Temperature Limits. Ref. 3.3.8.9.4, 4.5.1.3, 4.6.1.3, 4.6.6.2.2, 4.8.1.1.3, 4.8.2.1.3

RATED GAS TEMPERATURE °C (°F)

MAX ALLOWABLE TEMPERATURE

(MIN ENDURANCE TEST TEMPERATURE) °C (°F)

CONDITION MEASURED GAS TEMP

TURB. ROTOR INLET TEMP

MEASURED GAS TEMP

TURB. ROTOR INLET TEMP

CONTINGENCY MAXIMUM INTERMEDIATE MAXIMUM CONTINIOUS

TABLE VII. Not used.

TABLE VIII. Not used.

Appendix D

232

TABLE IX. Critical Safety Items Low Cycle Fatigue Lives. (Ref. 3.3.8.1.1)

Component Low Cycle Fatigue Life

Appendix D

233

TABLE X. Fuel contamination (Ref. 3.7.3.3.2,).

Contaminant Particle Size (microns)

Quantity (gm/1,000 gallons)

Ferrous-Ferric Iron oxide (Fe3O4, Black color) Magnetite

Ferric Iron Oxide (Fe2O3, Hematite)

Ferric Iron Oxide (Fe2O3, Hematite)

0-5

0-5

5-10

14.0

27.0

1.5

Crushed Quartz Crushed Quartz Crushed Quartz Crushed Quartz

1000-1500 420-1000 300-420 150-300

0.25 1.75 1.00 1.00

ISO 12 103, A4 Coarse Test Dust

8.00

Cotton Linters Staple below 7 (USDA grading standards SRA-AMS 180 and 251)

0.10

Naphthenic Acid (not required if service fuel is used for testing)

- 0.03% by volume.

Salt water prepared by dissolving salt in distilled water or other water containing not more than 200 parts per million of total solids

4 parts by weight of NaCl 96 parts by weight of H2O

0.01% volume (entrained)

Appendix D

234

TABLE XI. Fuel contamination (Ref. 4.6.2.2.3.3).

Contaminant Particle Size (microns)

Quantity (gm/1,000 gallons)

Ferrous-Ferric Iron Oxide (Fe3O4, black color Magnetite) Ferric Iron Oxide (Fe2O3, Hematite) Ferric Iron Oxide (Fe2O3, Hematite)

0-5

0-5

5-10

0.90

5.0

1.0 Prepared dirt IAW ISO 12103-1, A2 Fine Grade Test Dust

2.0

Cotton Linters Staple below 7 (USDA grading standards SRA-AMS 180 and 251)

0.02

Appendix D

235

TABLE XII. Data recording requirements. (Ref.4.3.5.4, 4.3.5.5, 4.3.5.6, 4.5.3, 4.6.3).

Data Recording Frequency

Taken During Tests

Start Once/30 min or once/cycle steady-state

Continuously during

transients

Cal/ recal

Power

1. Time of day X X X X 2. Total endurance time X X 3. Power setting X X X X X 4. Exhaust nozzle area sq. in. X X 5. Variable geometry position X X X 6. Engine rotor(s) speed(s) rpm X X X X X 7. Rotor speed(s) at idle rpm X X 8. Rotor speed(s) at ignition rpm X X 9. Rotor speed(s) at starter cutoutrpm

X X

10. Engine fuel flow lbm/hr X X X X 11. Data for determining output shaftpower (shp)

X X X X

12. Customer power extraction hp X X X X 13. Data for determining airflowlbm/sec

X X

14. Engine inlet total pressureaverage in Hg abs (Also compressorinlet total pressure average forengines with IPS)

X X X

15. Engine inlet total temperatureaverage °F (Also compressor

inlet total temperature average for engines with IPS)

X X X

16. Compressor discharge totalpressure in Hg abs

X X X X

17. Customer bleed air total pressurepsia

X X X

18. Compressor bleed air totalpressure (acceleration) psia

X

19. Customer bleed air static pres-sure psia

X X X

TABLE XII. Data recording requirements - Continued.

Appendix D

236

Data Recording Frequency Taken During Tests Start Once/30 min

or once/cycle steady-state

Continuously during

transients

Cal/ recal

Power

20. Compressor bleed air staticpressure (acceleration) psia

X

21. Compressor bleed air totaltemperatures (all ports) °F

X X X

22. Compressor bleed airflow rate(all ports) lbm/sec

X X X

23. Turbine discharge total pressureaverage in Hg abs

X X X

24. Compressor discharge totaltemperature °F

X X X X

25. Oil breather static pressure psig X X X 26. Gas generator turbine discharge

total temperature °FX X X X X

27. Exhaust nozzle exit total pressurein Hg abs

X X

28. Exhaust nozzle static pressurein Hg abs

X X X

29. Exhaust nozzle total temperature°F

X X X

30. Oil flow lbm/min X X X 31. Oil inlet temperature at pressure

pump inlet °C (°F)X X X X

32. Oil pressure at pressure pumpinlet kPa (psig)

X X X X

33. Oil pressure at pressure pumpoutlet kPa (psig)

X X X X

34. Oil pressure at scavenging pumpoutlet kPa (psig)

X X X

35. Oil temperature at scavengingpump outlet °C (°F)

X X X

36. Oil temperature at outlet fromfuel/oil cooler °C (°F)

X X X

37. Oil consumption for each cycle X

Appendix D

237

TABLE XII. Data recording requirements - Continued.

Data Recording Frequency

Taken During Tests

Start Once/30 min or once/cycle steady-state

Continuously during

transients

Cal/ reca

l

Power

38. Fuel pressure at fuel system inlet kPA (psig) X X X X 39. Fuel pressure at point shown on

engine configuration and envelopefigure kPa (psig)

X X X X

40. Fuel temperature at fuel system inlet °C(°F)

X X X

41. Fuel temperature at outlet fromfuel/oil cooler °C (°F)

X X X

42. Measured temperature °C (°F)(specify station location)

X X X X X

43. Maximum measured temperature°C (°F) (specify station location)

X X

44. Not used.45. Engine vibration at points shown on

engine configuration and envelope figurevelocity cm/sec (in/sec)

X X X X X

46. Accessory compartment tempera-ture(s) where applicable °C (°F)

X X X

47. Cooling air inlet temperature °C (°F) X X X 48. Cooling air outlet temperature °C (°F) X X X 49. Engine condition monitoring

system data (list each parameter)X X X

50. Cell static pressure kPa(in Hg abs)

X X X X

51. Ignition source voltage and current(when external power is being used)

X X X

52. Oil leakage at accessory pads X X X 53. Start number X X 54. Time to ignition actuation X X

Appendix D

238

TABLE XII. Data recording requirements - Continued.

Data Recording Frequency

Taken During Tests

Start Once/30 min or once/cycle steady-state

Continuously during

transients

Cal/ recal

Power

55. Time to light-off X X X 56. Time to starter cutout X X X 57. Time to stabilized idle RPM X X X 58. Time to oil pressure indication

(at point shown on engineconfiguration and envelopefigure

X X X

59. Time to stabilize to normal oilpressure (at point shown onengine configuration andenvelope figure)

X X X

60. Engine life counter X X X X 61. Data for calculating gas

generator first stage turbine rotorinlet gas temperature

X X X X X

62. Additional data as required bythe Using Service

X X X X X

Appendix D

239

TABLE XIII. Sea level anti-icing conditions (Ref. 3.2.5.2, 4.6.4.2).

Part 1 Part 2 Ambient Temperature

-21 to -19°C (-6 to 2°F)

-6 to -4°C (22 to 24°F)

-6 to -4°C (22 to 24°F)

Velocity 0 to 31 m/sec (0 to 60 knots)

0 to 31 m/sec (0 to 60 knots)

0 to 31 m/sec (0 to 60 knots)

Altitude 0 to 150 m (0 to 500 ft)

0 to 150 m (0 to 500 ft)

0 to 150 m (0 to 500 ft)

Mean Effective Drop Diameter

20 +5 microns 20 +5 microns 30 +5 microns

Liquid Water Content (Continuous)

1.25 to 1.75 gm/m3

2.85 to 3.35 gm/m3

0.6 to 0.8 gm/m3

Appendix D

240

TABLE XIV. Schedule of salt spray injection endurance cycle (Ref. 4.6.4.3).

Phase No. 1/ Duration of Phase (hours)

Test Engine Operation

Salt Solution Parts per

Billion (PPB)

Engine Ambient Air

Temperature Relative Humidity

1 3 2/, 3/ Operating 200+40 4/ 10°C min. 50°F min.

73% min.

2 2 5/ Not operating 0 Atmospheric Atmospheric

3 7 5/ Not operating 200+40 4/, 6/ 10°C min. 50°F min.

73% min.

4 12 5/ Not operating 0 43+5°C 109+9°F

90% min.

NOTES: 1/ Engine inlet and exhaust openings shall remain open for all phases of the test cycle. 2/ During shutdown, while the engine is decelerating from idle, the salt solution shall continue to be sprayed into the engine until the rotor has come to rest. 3/ An engine operating cycle shall be selected from the mission mix which has at least six power levels with dwell times of at least 5 minutes per level. The power levels shall roughly be equally spaced throughout the cycle. If a suitable cycle cannot be identified, then the cycle of figure 23 shall be used. The cycle shall be constructed so that the metal temperature in the hot gas flow path is between 760°C and 980°C (1400°F and 1800°F). This shall be accomplished by varying the turbine inlet temperature from 760°C (1400°F) to the maximum temperature. 4/ Salt solution ingested by the engine shall conform to that specified in Note 7/ below, and shall be regulated to provide a concentration of 200+40 PPB by weight of salt in air during those phases of each cycle which requires salt ingestion. A salt sampling system shall be employed to determine the concentration level during each cycle of operation. 5/ The duration of phases 2, 3, and 4 can vary up to 10 percent of the time for the respective phase for any particular cycle. However, the total time for phases 1 - 4 (1 cycle) shall be 24 hours. Down time shall be considered as phase 2. Down time shall include time for engine facility maintenance, weekends, holidays, etc. The total down

Appendix D

241

TABLE XIV: Schedule of salt spray injection endurance cycles (continued). time shall not be less than 50 hours for 25 cycles. The maximum down time is negotiable. The minimum duration for phases 3 and 4 is 175 and 300 hours, respectively. 6/ The test facility shall provide the flow of salt-laden air through the engine gas flowpath(s) and over the external surfaces of the engine. The external flow velocity shall be between 4.6-6.1 m/sec (15-20 ft/sec). 7/ The basic salt formulation shall be composed of the following materials dissolved with sufficient distilled water to make one liter of salt solution. Additional distilled water may be added to the salt solution, as required, to provide a uniform salt aerosol profile across the face of the engine. The maximum dilution shall not exceed 40:1. The aerosol droplet size shall not exceed 25 microns.

Chemical Designation Quantity per liter of stock solution NaCl (c.p.) 23 grams

Na2SO4 10H2O 8 grams

Stock solution (see note 8/) 20 milliliters

8/ The stock solution shall be composed of the following materials dissolved with sufficient distilled water to make one liter of stock solution:

Chemical Designation Quantity per liter of stock solution KCl (c.p.) 10 grams

KBr 45 grams MgCl2 6H2O (c.p.) 550 grams

CaCl2 6H2O (c.p.) 110 grams

242

Appendix D

TABLE XV. Electromagnetic compatibility test methods.

(Contractor to provide.)

TABLE XVI. Provisions for engine condition indication. (Contractor to provide.)

TABLE XVII. EMS BIT code data. (Contractor to provide.)

TABLE XVIII. Components requiring functional bench calibration or ATP (Ref. 4.5.1.2.1, 4.6.1.2.1, 4.8.1.1.2.1). (Contractor to provide.)

TABLE XIX. Engine split-line flange leakage.

(Contractor to provide.)

Appendix D

243

TABLE XX-A. Component tests for EIT (Ref. 3.1.3). Assembly Missionized

/operational test

Environmental tests

shock vibration humidity fungus sand and dust

sustained acceleration

explosion proof

variable vanes

assembly load

control valve surge control valve

bleed flow sensor

assembly starter/

generator alternator ignition system

fuel control

assembly (T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

244

TABLE XX-A. Component tests for EIT - Continued. Assembly Individual Tests Reliability and Maintenance Test

fire cavitation low lubricity

sneak circuit

analysis

environmental stress

screening

reliability growth

test

built in test

variable vanes

assembly load

control valve surge control valve bleed flow

sensor assembly starter/

generator alternator ignition system

fuel control

assembly (T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

245

TABLE XX-A. Component tests for EIT (Continued). Assembly Missionize

d/operational test

Environmental tests

shock vibration humidity fungus sand and dust

sustained acceleration

explosion proof

fuel distribution

system oil system hydraulic system Engine history

recorder or counter

miscellaneous controls

and accessories Accessory drive/PTO Anti-icing system

Inlet Particle Separator plumbing

T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

246

TABLE XX-A. Component tests for EIT - Continued.

Assembly Individual Tests Reliability and Maintenance Test fire cavitation low

lubricity sneak circuit

analysis

environmental stress

screening

reliability growth

test

built in test

fuel distribution

system oil system hydraulic system Engine history

recorder or counter

miscellaneous controls

and accessories Accessory drive/PTO plumbing

(T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

247

TABLE XX-B. Component tests for QT (Ref. 3.1.3). Assembly Missionized

/operational test

Environmental tests

shock vibration humidity fungus sand and dust

sustained acceleration

explosion proof

variable vanes

assembly load

control valve surge control valve

bleed flow sensor

assembly starter/

generator alternator ignition system

fuel control

assembly (T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

248

TABLE XX-B. Component tests for QT - Continued. Assembly Individual Tests Reliability and Maintenance Test

fire cavitation low lubricity

sneak circuit

analysis

environmental stress

screening

reliability growth

test

built in

test variable vanes

assembly load

control valve surge control valve bleed flow

sensor assembly starter/

generator alternator ignition system

fuel control

assembly (T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

249

TABLE XX-B. Component tests for QT (Continued). Assembly Missionized/

operational test

Environmental tests

shock vibration humidity fungus sand and dust

sustained acceleration

explosion proof

fuel distribution

system oil system hydraulic system Engine history

recorder or counter

miscellaneous

controls and

accessories

Accessory drive/PTO Anti-icing system

Inlet Particle

Separator plumbing

T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

250

TABLE XX-B. Component tests for QT - Continued.

Assembly Individual Tests Reliability and Maintenance Test fire cavitation low

lubricity sneak circuit

analysis

environmental stress

screening

reliability growth test

built in test

fuel distribution

system oil system hydraulic system Engine history

recorder or counter

miscellaneous controls

and accessories Accessory drive/PTO plumbing

(T) = test, (TA) = test conducted as part of an assembly, (S) = similarity, (H) = hermeticallysealed, (N/A) = not applicableNote: Environmental tests must be conducted in sequence from left to right as presented on thetable. The same physical component must be used for the entire sequence of environmentaltests.

Appendix D

251

TABLE XXI. Test plan documentation requirements (Ref 4.3.5.1.1).

a. Cover.(1) Title, number and source of test procedure.(2) Date.(3) Names of the author(s).(4) Contract number.

b. Title page. Include same information as a.(1) through a.(4).c. Introduction.

(1) Identify item under test.(2) Government witnessing organization.

d. Applicable documents.(1) Test procedure related specifications, development plans, parts list documents,

inspection specifications, assembly instructions, detailed test directives, acceptancetest procedures, etc.,

(2) Any document without prior government approval must be attached to the procedure forUsing Service approval.

e. Test objective.(1) Test purpose,(2) Anticipated results.(3) Specification requirements to be verified (cite specific paragraph).(4) Success criteria, hardware condition, post test performance.

f. Test article description.(1) Conformance to official parts list, deviation from parts list with explanation/justification,

include schematic, drawings, pictures as required.(2) Identify specific build documents by number, revision, and date,(3) State relevant features (e.g., if engine performance test, state build clearances, nozzle

areas, etc.).(4) Identify specific test article by part number(s) and serial number(s), procedures for

approval of parts replacement.

Appendix D

252

TABLE XXI. Test plan documentation requirements - Continued.

g. Test requirements.(1) Description of test facilities, power absorption, inlet/exhaust, control/inputs, electrical

power required, starting system, test conditions/environment.(2) Description of special test equipment used.(3) Description of instrumentation placement and accuracy.(4) Description of data acquisition system, operating parameters (e.g., fuels, oils, test

limits).(5) All analysis or preliminary activity required to establish test conditions (e.g., spin pit

speeds or strain gage locations) shall be described or referenced.h. Test procedures.

(1) Pretest procedure (initial set-up, calibrations, pretest check-outs).(2) Conduct of test, description of specific actions required by specification (e.g., cycling,

stabilization, starting specifics, bleed air extraction and transient operations),(3) Deviations from specification with explanation or justification,(4) Interpretation of specification requirements (e.g., performance measurement is SHP, in

lieu of temperature).(5) Post test calibrations, inspections, teardowns.

i. Data.(1) Collection and reduction methodology.(2) Anticipated data points.

Appendix D

253

TABLE XXII. Maintainability Requirements

a. The contractor should design the engine to minimize required skills, tools, and support equipment.Required tooling shall be held to a minimum. The engine shall be capable of being maintained withstandard hand tools, as identified in accompanying table. Where provisions for standard hand toolsare not feasible, the design should provide, wherever possible and cost effective, for special toolsand test equipment that are available and in use on other in-service engines, as identified in theaccompanying table.

b. Parts of the engine requiring routine service, checking, adjustment, or frequent replacement shouldbe made readily accessible without removal of other parts or components, and without disassemblyof the engine.

c. Provisions for access to the location of any Line Replaceable Unit (LRU) should be based on reliabilitypredictions, i.e., the least reliable LRU on the engine should have the best accessibility.

d. LRUs should be no more than one deep. Removal of the LRU should not require the removal ofunaffected LRUs.

e. Shimming or measuring of LRUs should not be required for LRU replacement.

f. No adjustments, balancing, matching, or rigging should be required to install or replace LRUs on anassembled engine.

g. Plumbing or electrical removal should not be required for LRU access or removal.

h. Captive fasteners should be used for LRU mounts and connections. No fasteners smaller than 6-mm (1/4-inch) diameter should be used for LRU mounts or for organizational level maintenance.Individual LRUs should have the same style and size fastener for mounting and the same style andsize fasteners for mating.

i. Positive securing clamps, nuts, bolts, connectors, and keyed interfaces should be provided to preventmisalignment and improper installation of LRUs, plumbing, and cables. The design should permitremoval without exceeding capacity or capability of required tools after the engine has been inservice.

j. Fluid carrying systems subject to disassembly for LRU or engine replacement should be designed tominimize system drainage or spillage on the maintainer.

k. Features should be provided to prevent functionally different components with the same or similardesign from being improperly installed.

l. Safety wire, cotter pins, and bent tab washers should not be used in securing connectors, fasteners,etc. at the intermediate and organizational levels and should be kept to a minimum at the depot level.

m. The engine design should allow centrally located engine to air vehicle disconnects, subject tosurvivability and vulnerability requirements.

n. The design should not require engine-to-engine, engine to aircraft, or accessory and installationrigging. System alignment or complicated operational checks should not be required.

o. The design should allow for adequate access, space, and tool clearance around fasteners andconnectors to provide: adequate clearance for component removal and installation at all levels ofassembly; and to provide proper seating of tools and uniform torque application.

p. LRU modules and any SRU should be interchangeable with like items without match balancing.

q. The design should provide provisions for rapid repair such as external tubing repair versus whole linereplacement.

r. Sensor check or replacement should not require engine rig or trim. Any sensor that requires morethan 20 minutes to replace while the engine is installed in the aircraft should not be mission critical.

Appendix D

254

s. Aircraft leveling should not be required for installed engine servicing or quantity status reading.

t. Fluid level quantity indicators, if part of the engine design, should be clearly visible for accuratereading when using flashlights as approved by the Using Service or natural daylight illuminationsources.

u. Connecting devices should be selected to prevent improper installation while minimizing the numberof different parts. MIL-HDBK-965 or an equivalent commercial practice can be used for guidance.

v. No routine or scheduled maintenance should be needed for field washing, cleaning, or drying ofengine components. However, the engine should incorporate provisions for water wash compressorcleaning if required.

w. The design should consider including provisions for the incorporation and use of pyrometers to checktemperatures on all turbine blades during depot level testing. This should be verified during the initialdesign reviews.

x. Engine operation after installation in the aircraft should not be required except for leak check andcontrol.

y. Disconnect points for external plumbing should coincide with module or SRU flange and disconnectlocations.

z. The color red should not be used on the engine for any identification means without an adjacentsymbol readily visible under red-light conditions.

aa. Components subject to wear, requiring repair or rework to restore dimensioned features, should provide a dimensional baseline datum that is not destroyed or degraded in service or during repair or rework.

bb. The engine configuration should permit installation in the aircraft with as-received components and connectors mounted on the engine.

cc. Tubing installation should use ARP 994 as a guide. Plumbing and tubing lines should be identifiedby labels, markings, color coding, or a combination of these methods.

dd. The engine design shall permit maximum use of non-destructive inspection techniques andmulti-purpose test and inspection equipment. Inspection provisions, including access envelopes,shall be shown on the engine configuration and envelope figure.

ee. A positive means of slowly rotating the rotor system should be provided to facilitate inspection. The same tool should be used for removal and replacement of port covers, plugs, and associated fasteners. The number and location of access ports should be kept to a minimum, to minimize sources of leakage and maintenance discrepancies, while providing adequate inspection of critical areas. Access ports should be accessible without removing other components

ff. The contractor should propose the location of access ports or the areas to be seen through inspection. Typical areas of interest are:

1. Compressor inlet guide vane leading edge.

2. Compressor inlet guide vane trailing edge and compressor first stage rotor leading edge.

3. Compressor last stage rotor trailing edge and combustor inlet.

4. Combustor liner and fuel nozzle faces.

5. Combustor outlet and turbine first stage vane leading edge.

6. HPT first stage vane trailing edge and first stage rotor leading edge.

7. HPT last stage rotor trailing edge and low pressure turbine first stage vane leading edge.

8. LPT first stage vane trailing edge and first stage rotor leading edge.9. PT first stage vane trailing edge and first stage rotor leading edge.

Appendix D

255

TABLE XXIII. Turboshaft engine 1.22 km (4000 ft), 35°C (95°F) output shaft speed droop requirements.

Maneuver Endpoints

T - Transient Time (Seconds)

Q1 Q2 1 2 3 4 5 6 8 0 100 (d11) (d12) (d13) (d14) (d15) (d16) (d18) 40 100 (d21) (d22) (d23) (d24) (d25) (d26) (d28)

100 0 (d31) (d32) (d33) (d34) (d35) (d36) (d38) 100 40 (d41) (d42) (d43) (d44) (d45) (d46) (d48)

Percent Aircraft Torque

Percent Maximum Allowable Droop

Notes for Table XXIII: (1) Torque shall be applied linearly as in the accompanying figure(2) Torque is given as a percentage of AIRCRAFT torque determined using 100% engine outputshaft speed and the value of the power required for each engine as specified by the Using Servicethrough coordination with the airframe manufacturer as opposed to available ENGINE torque.(3) Droop requirement represents a maximum transient speed decrease (excursion from steadystate).(4) A negative droop represents a transient speed increase (excursion from steady state), neverto reach the engine’s overspeed shutoff threshold.(5) Zero percent torque condition is with the power turbine clutched to rotor drive system (zeroneedle split).

Values for the Table XXIII shall be as follows: d11) The value shall not exceed 12%.

Appendix D

256

(d12) The value shall not exceed 6%. (d13) The value shall not exceed 4%. (d14) The value shall not exceed 3%. (d15) The value shall not exceed 2%. (d16) The value shall not exceed 1%. (d18) The value shall not exceed 1%. (d21) The value shall not exceed 3%. (d22) The value shall not exceed 2%. (d23) The value shall not exceed 2%. (d24) The value shall not exceed 1%. (d25) The value shall not exceed 1%. (d26) The value shall not exceed 1%. (d28) The value shall not exceed 1%. (d31) The value shall not exceed -8% (d32) The value shall not exceed -4%. (d33) The value shall not exceed -3%. (d34) The value shall not exceed -2%. (d35) The value shall not exceed -1%. (d36) The value shall not exceed -1%. (d38) The value shall not exceed -1%. (d41) The value shall not exceed -1%. (d42) The value shall not exceed -2%. (d43) The value shall not exceed -1%. (d44) The value shall not exceed -1%. (d45) The value shall not exceed -1%. (d46) The value shall not exceed -1%. (d47) The value shall not exceed -1%.

Appendix D

257

TABLE XXIV. Parts classification (Ref. 3.3.8.1.1).

(Contractor to provide.)

Appendix D

258

TABLE XXV. Component Test Plan Documentation Requirements. (Referenced by: 4.3.5.1.2)

Part 1 – General 1. Front end matter in accordance with TABLE XXVI.2. Component description.

a. Functional description.(1) Component operational block diagram and/or schematic.(2) Relationship of the component to the overall control system, including

interfaces with other subsystems (airframe or engine).b. Physical description.

(1) 2-D and isometric drawings.(2) Photos (if available).

3. Component performance requirements.a. Required performance (i.e., pass/fail criteria) must be traceable to higher level

specification(s). Include a table that traces the performance requirements to thenext highest level specification (generally the ATP), or higher if necessary, tojustify the pass/fail values. Each parameter and its limits must reference a specificparagraph number.

b. If the parameter and its pass/fail limits cannot be easily translated from onespecification to the next (e.g., engineering units to electrical units), then there mustbe an explanation of how the conversion was made.

4. Testing to be performed.a. List of tests with a simplified description of each, the model specification or other

requirement paragraph, and any standard procedure reference (e.g.,MIL-STD-810G, Method 514.5, Procedure 1, etc.).

b. Order of testing and explanation if it deviates from the order specified.c. Any planned ATPs/calibrations, inspections, or other disassemblies planned

between the tests.d. For each test requirement, specify whether it will be satisfied via test, similarity

and/or analysis.e. Specify how many units will be tested and for which tests.

Appendix D

259

TABLE XXV. Test Plan Documentation Requirements - Continued. 5. Test summaries. These are meant to be 3 to 5 pages for each test, containing the

following information.a. Test purpose. Provide a clear, concise, easily understood description, for

example: “Demonstrate that the component meets its specified operational performance under a typical engine vibration environment as represented by Figure XXXX in MIL-STD-810G and 2) Demonstrate that the component meets the goal of XXXX hours of useful life by subjecting it to an accelerated vibration profile and then verifying that it still passes its ATP service limits.”

b. Test performance. Explain if the test will be performed exactly as described in thereferenced standard or procedure or if it will be some modified version. If asimilarity and/or analysis report will be prepared in lieu of testing (as identified in4d, then a summary of the rationale shall be included here, with the remainingitems c through g marked as N/A or left out completely.

c. Any planned ATPs/calibrations, inspections, or other disassemblies plannedduring the test sequence.

d. Data monitoring and recording. Explain what will be recorded and monitored andhow it will be captured (i.e., continuous vs. sampling).

e. Unit testing conditions. The environmental cycle/conditions, the unit functionalcycle/conditions, and the unit operating cycle/conditions must all be defined.

f. Pass/fail criteria. Specify, referencing the performance table at the front of thedocument. If different, then justify. List any other criteria such as passage of post-test ATPs or detailed inspections.

g. Test data presentation. Explain exactly what information will be provided to theAED as documentary evidence of test passage and in what form it will bepresented.

Part 2 - Detailed Test Procedures Separate tabbed sections or appendices shall be provided, one for each test procedure. The test procedure shall not duplicate information already provided in Part 1 of the test plan. These procedures may be submitted individually at the discretion of the contractor. However, Part 1 of the test plan shall have already been approved prior to the contractor submitting any detailed procedures for AED approval. In addition to providing detailed step-by-step instructions for test execution, the procedure shall also contain the following information, as a minimum. 1. Component mounting details and test orientation with orthogonal axis definition.2. Monitoring sensor placement locations.3. Drawing/sketches of the test setup showing all pertinent equipment by name and/or part

number.4. Planned test location.5. Serial number of the component undergoing test6. List of test equipment (including ranges and accuracies) and any special hardware or test

equipment requirements.

260

Appendix D

TABLE XXVI. Plans and Reports Front End Matter. (Referenced by: TABLE XXV)

1. Cover.a. Title, number, source of the report, and date.b. Contract number.

2. Title page.a. Title, number, source of the report, and date.b. Name(s) of the author(s).c. Contract number.

3. Abstract.a. Objective of the report .b. Brief statement of the contents of the report.

4. Table of contents.5. List of tables.

a. When used in a separate series, tables should be given Roman numerals.(Example: Table I, Table II, and Table III)

6. List of illustrations.a. Figure numbers and captions of all illustrations.b. Photographs, charts, and graphs should be treated as illustrations and given

figure numbers.7. Applicable documents.

AV-E-8593 AQP

APPENDIX A

REF. - 6.3

REQUIREMENTS FOR ENGINE INTEGRITY TESTING (EIT)

261

Para No. Requirements Analysis Demonstration Test Notes

3.1.2.5 Externally Applied Forces X A 3.1.2.8.2 Heat Rejection and Cooling Test

Report X A

3.2.1.3 Performance Verification X A 3.2.1.4.10 Vibration Limits X Q 3.3.1.1 Materials and Processes X C 3.3.8.1 3.3.8.1.2 3.3.8.2 3.3.8.3 3.3.8.4 3.3.8.4.1 3.3.8.5 3.3.8.6 3.3.8.6.1

Structural Life Parts Classification – Critical Safety Items High Cycle Fatigue Life (HCF) Low Cycle Fatigue Life (LCF) Strength Factors of Safety Material Properties Strength and Life Analysis Points of Life

X X

X X X X X X X

B B

U B B B B D B

3.3.8.8 Creep X B 3.3.8.9.1 3.3.8.9.4 3.3.8.10 3.3.8.10.1 3.3.8.10.2

Containment Disk Burst Vibration and Dynamic Response Critical Speeds Vibration and Stress Analysis

X X X X X

B B P P A

3.9.1 Engineering Evaluation Tests X F 4.3.5.1 Test Plans and Procedures X G 4.3.6.1 Test Reports X H 4.3.6.2 Summary Reports X I 4.4.1 Customer Bleed Analysis X J 4.4.2 Engine Heat Rejection and Oil

Cooling X K

4.5.1 Endurance Test X N 4.5.1.1.1 Engine Dry Weight X 4.5.1.2.1 Component Calibration X O 4.5.1.2.2.1. Customer Bleed Air Analysis X X 4.5.1.3 Endurance Test Procedure X 4.5.1.3.1 Starts X 4.5.1.3.2 Contingency Power Flight Rating X 4.5.1.4.1 Engine Recalibration X 4.5.1.4.2 Component Recalibration X 4.5.1.5 4.5.1.6

Engine Disassembly and Inspection Endurance Test Completion

X

Y

4.5.2 Engine Component Tests X 4.5.2.1 Previous Component Approval X 4.5.2.2 Component Acceptance Test X 4.5.2.3 Component Re-Test, Disassembly,

and Inspection X

AV-E-8593 AQP

APPENDIX A (continued)

REF. - 6.3

REQUIREMENTS FOR ENGINE INTEGRITY TESTING (EIT)

262

Para No. Requirements Analysis Demonstration Test Notes

4.5.2.4 Component Test Success Criteria X X 4.5.2.5.2 Vibration – Airframe and Engine X Y 4.5.2.5.9 Software Verification Tests X 4.5.3 Altitude Test (Engine) X 4.5.3.1 Altitude Calibration X 4.5.3.2 Altitude Test Procedure X 4.5.3.3 Altitude Test Completion X 4.5.4.8 Low Cycle Fatigue Engine Test X Y 4.5.4.9 Main Shaft, Seals and Bearing

Mechanical Tests. X

AV-E-8593 AQP

APPENDIX A (continued)

REF. - 6.3

REQUIREMENTS FOR ENGINE INTEGRITY TESTING (EIT)

263

NOTES:

A. A report shall be submitted to the Using Service prior to the initiation of the EIT. B. Analysis results shall be included in the strength and life analysis. C. When engine manufacturer's documents are used for materials and processes, such

documents shall be subject to review by the Using Service prior to the start of the EIT D. The strength and life analysis report shall be updated prior to EIT E. Matched or selected fit parts shall be identified and listing shall be provided to the Using

Service prior to EIT. F. Test reports shall be submitted prior to the completion of the EIT. G. Pretest data shall be submitted to the Using Service for approval prior to the initiation of each

test or demonstration required for EIT. H. Test reports shall be submitted to the Using Service for approval following completion of each

test or demonstration required for EIT. I. Summary reports shall be submitted to the Using Service following completion of EIT. J. A report shall be submitted prior to the start of EIT endurance test. K. A report verifying the analytical heat rejection and surface temperature analysis shown in the

engine system specification shall be submitted to the Using Service within 30 days after completion of the test.

L. A vibration survey report shall be submitted to the Using Service for approval within 30 days after completion of the test.

M. Procedure for accomplishing the starting torque demonstration shall be submitted to the Using Service for approval prior to initiation of the test.

N. Engine weight shall be measured, before the engine has been serviced with fuel or oil, prior to initiation of calibrations.

O. Shall be performed prior to the engine Calibration of 4.5.1.2.2. P. Analysis results shall be included in the vibration and stress analysis. Q. True RMS velocity limits shall be provided to the Using Service prior to initiation of the EIT

for inclusion in the engine specification. R. A list of adhesives and sealants shall be provided for Using Service approval 60 days after

contract. S. The near and far field engine noise signatures at the Table 1 ratings shall be provided to the

Using Service prior to initiation of the EIT. T. Correlation studies shall be presented to the Using Service prior to initiation of the EIT

endurance test. U. Specify in the pretest data the failure modes to be evaluated. The failure modes shall be

subject to approval by the Using Service. V. Module replacement limits shall be provided to the Using Service prior to the start of the EIT

and shall be incorporated into the ESS by SCN action. W. The results of the customer bleed air analysis and the method and test apparatus used

shall be detailed in the test report. Y. A “dirty inspection” and “clean inspection” IAW 4.5.1.5

AV-E-8593 AQP

APPENDIX B

REF. - 6.3

REQUIREMENTS FOR PRELIMINARY FLIGHT RATING (PFR)

264

Para No. Requirements Analysis Demonstration Test Notes

3.1.2.5 Externally Applied Forces X A 3.1.2.8.2 Heat Rejection and Cooling Test

Report X K

3.2.1.3 Performance Verification X A 3.2.1.4.10 Vibration Limits X Q 3.2.5.7 Noise levels X J,S 3.2.1.5.8 Droop X J 3.2.8.4.1 Electromagnetic Interference (EMI) A 3.3.1.1 Materials and Processes X C 3.3.1.1.1 Adhesive and Sealants R 3.3.5 Interchangeability A,E 3.3.8.1 3.3.8.1.2 3.3.8.2 3.3.8.3 3.3.8.4 3.3.8.4.1 3.3.8.5 3.3.8.6 3.3.8.6.1

Structural Life Parts Classification – Critical Safety Items High Cycle Fatigue Life (HCF) Low Cycle Fatigue Life (LCF) Strength Factors of Safety Material Properties Strength and Life Analysis Points of Life

X X

X X X X X X X

B B

U,Y B,Y B,Y B,Y B,Y D,Y B,Y

3.3.8.8 Creep X B,Y 3.3.8.9.1 3.3.8.9.4 3.3.8.10 3.3.8.10.1 3.3.8.10.2

Containment Disk Burst Vibration and Dynamic Response Critical Speeds Vibration and Stress Analysis

X X X X X

B,Y B,Y P,Y P,Y A

3.5.1.1 Modular Concept X A,V 3.7.13 Wash System X J 3.9.1 Engineering Evaluation Tests X F,H 3.9.2 Preliminary Flight Rating (PFR) X F,G,H 4.3.5.1 Test Plans and Procedures X G 4.3.6.1 Test Reports X H 4.3.6.2 Summary Reports X I 4.4.1 Customer Bleed Analysis X J,Y 4.4.2 Engine Heat Rejection and Oil

Cooling X K.W

4.4.3 Oil Flow Interruption Test X J,Z 4.4.4 Engine Electrical Power failure Tests X J 4.4.5 Engine Vibration Survey X L 4.4.6 Starting Torque X M,J 4.4.10 Verification of Correction Factors X A 4.5.1 Endurance Test X 4.5.1.1.1 Engine Dry Weight X N,X 4.5.1.2.1 Component Calibration X O,X 4.5.1.2.2` Engine Calibration X G,X

AV-E-8593 AQP

APPENDIX B (continued)

REF. - 6.3

REQUIREMENTS FOR PRELIMINARY FLIGHT RATIING (PFR)

265

Para No. Requirements Analysis Demonstration Test Notes

4.5.1.2.2.1. Customer Bleed Air Analysis X X W,X 4.5.1.3 Endurance Test Procedure X X 4.5.1.3.1 Starts X X 4.5.1.3.2 Contingency Power Flight Rating X X 4.5.1.4.1 Engine Recalibration X X 4.5.1.4.2 Component Recalibration X X 4.5.1.5 4.5.1.6

Engine Disassembly and Inspection Endurance Test Completion

X X

X,A1

4.5.2 Engine Component Tests X 4.5.2.1 Previous Component Approval X 4.5.2.2 Component Acceptance Test X G, 4.5.2.3 Component Re-Test, Disassembly,

and Inspection X

4.5.2.4 Component Test Success Criteria X X 4.5.2.5.1 Explosion Atmosphere X 4.5.2.5.2 Vibration – Airframe and Engine X X,A1 4.5.2.5.3 Fuel Pump Altitude Test X A1 4.5.2.5.4 Oil Reservoir Pressure Test X Z 4.5.2.5.5 Fire Test X A1 4.5.2.5.6 Overheat Test X Z 4.5.2.5.7 Output Shaft Static Torque Test X 4.5.2.5.8 Impact (Shock) Test X Z 4.5.2.5.9 Software Verification Tests X X 4.5.2.5.10 Electrical Power System Test X 4.5.2.5.10.1 Engine-Supplied Electrical Power X 4.5.2.5.10.2 Aircraft-Powered Electrical Power X 4.5.2.5.11 Helicopter Drive System Torsional

Stability XX

4.5.3 Altitude Test (Engine) X X 4.5.3.1 Altitude Calibration X G,X 4.5.3.2 Altitude Test Procedure X X 4.5.3.3 Altitude Test Completion X X 4.5.4.1 Component Vibration

Characterization X

4.5.4.2.1 Overspeed Test X 4.5.4.2.2 Overtemperature Test X A1 4.5.4.3 Engine Static Load test X 4.5.4.4 Attitude Test X 4.5.4.5 Loss of Oil Test X A1 4.5.4.6 Engine Crash Load test Z 4.5.4.9 Main Shaft Seals and Bearing

Mechanical Tests. X X

4.5.5.1 Electromagnetic Environmental Effects (E3)

X

4.5.5.2 Electromagnetic Interference (EMI) X

AV-E-8593 AQP

APPENDIX B (continued)

REF. - 6.3

REQUIREMENTS FOR PRELIMINARY FLIGHT RATIING (PFR)

266

4.9.1.1 1000 Cycle Low Cycle Fatigue Test X J,A1 4.9.1.2 Preliminary Environmental Icing Test X J 4.9.1.3 60 Hour Benchmark Endurance Test X J,A1 4.9.1.4 Fine Sand Ingestion Test X J,A1 4.9.1.6 Highly Accelerated Life Test (HALT) X A,Z 4.9.1.8 Inflow Bleed Test X J 4.9.1.9 Engine Backpressure Test X J 4.9.1.13 Fuel System Suction Test X J,A1

AV-E-8593 AQP

APPENDIX B (continued)

REF. - 6.3

REQUIREMENTS FOR PRELIMINARY FLIGHT RATIING (PFR)

267

NOTES: A. A report shall be submitted to the Using Service prior to the initiation of PFR. B. Analysis results shall be included in the strength and life analysis. C. When engine manufacturer's documents are used for materials and processes, such

documents shall be subject to review by the Using Service prior to the start of the PFR D. The strength and life analysis report shall be updated prior to PFR E. Matched or selected fit parts shall be identified and listing shall be provided to the Using

Service prior to PFR. F. Test reports shall be submitted prior to the completion of the PFR. G. Pretest data shall be submitted to the Using Service for approval prior to the initiation of each

test or demonstration required for PFR. H. Test reports shall be submitted to the Using Service for approval following completion of each

test or demonstration required for PFR. I. Summary reports shall be submitted to the Using Service following completion of PFR. J. A report shall be submitted prior to the start of PFR endurance test. K. A report verifying the analytical heat rejection and surface temperature analysis shown in the

engine system specification shall be submitted to the Using Service within 30 days after completion of the test.

L. A vibration survey report shall be submitted to the Using Service for approval within 30 days after completion of the test.

M. Procedure for accomplishing the starting torque demonstration shall be submitted to the Using Service for approval prior to initiation of the test.

N. Engine weight shall be measured, before the engine has been serviced with fuel or oil, prior to initiation of calibrations.

O. Shall be performed prior to the engine Calibration of 4.5.1.2.2. P. Analysis results shall be included in the vibration and stress analysis. Q. True RMS velocity limits shall be provided to the Using Service prior to initiation of the PFR

for inclusion in the engine specification. R. A list of adhesives and sealants shall be provided for Using Service approval 60 days after

contract. S. The near and far field engine noise signatures at the Table 1 ratings shall be provided to the

Using Service prior to initiation of the PFR. T. Correlation studies shall be presented to the Using Service prior to initiation of the PFR

endurance test. U. Specify in the pretest data the failure modes to be evaluated. The failure modes shall be

subject to approval by the Using Service. V. Module replacement limits shall be provided to the Using Service prior to the start of the

PFR and shall be incorporated into the ESS by SCN action. W. The results of the customer bleed air analysis and the method and test apparatus used

shall be detailed in the test report. X. This test shall be required only if significant components have been changed from the

configuration of EIT. Y. This analysis shall be required only if significant components have been changed from the

configuration of EIT. Z A “dirty inspection” IAW 4.5.1.5 A1 A “dirty inspection” and “clean inspection” IAW 4.5.1.5

268

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APPENDIX C

REF. - 6.3

REQUIREMENTS FOR QUALIFICATION TEST (QT) RATING

Para No. Requirements Analysis Demonstration Test Notes 3.2.1.3 Performance Verification X A

3.2.1.4.5 Engine Temperature Limits X Q 3.2.1.5.8 Droop X N 3.2.8.4.1 Electromagnetic Interference (EMI) X B 3.3.5 Interchangeability X C 3.3.8.1 3.3.8.1.2 3.3.8.2 3.3.8.3 3.3.8.4 3.3.8.4.1 3.3.8.5 3.3.8.6 3.3.8.6.1

Structural Life Parts Classification – Critical Safety Items High Cycle Fatigue Life (HCF) Low Cycle Fatigue Life (LCF) Strength Factors of Safety Material Properties Strength and Life Analysis Points of Life

X X

X X X X X X X

E E

U E E E E D E

3.3.8.8 Creep X N 3.3.8.10.2 Vibration and Stress Analysis X O 3.7.2.3.3.1 Power Turbine Overspeed Control

System X

3.9.1 Engineering Evaluation Tests X G 3.9.3 Qualification Rating (QT) X I 4.3.5.1.1 4.3.5.1.2

Engine & Engine Subassembles Components

X X

H H

4.3.6.1 Test Reports X I 4.3.6.2 Summary Reports X J 4.4.2 Engine Heat Rejection and Oil

Cooling Test X I

4.6.1 Endurance Test X 4.6.1.1.1 Engine Dry Weight X K 4.6.1.2.1 Component Calibration X L 4.6.1.2.2 Engine Calibration X 4.6.1.2.3 Customer Bleed Air Analysis X X T 4.6.1.3 Endurance Test Procedure X 4.6.1.3.1 Starts X 4.6.1.3.2 Contingency Power Qualification X 4.6.1.4.1 Engine Recalibration X 4.6.1.4.2 Component Recalibration X 4.6.1.5 Engine Dissassembly and Inspection X Z 4.6.1.7 Endurance Test Completion X 4.6.2 Engine Component Tests X

269

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APPENDIX C (continued)

REF. - 6.3

REQUIREMENTS FOR QUALIFICATION TEST (QT) RATING

Para No. Requirements Analysis Demonstration Test Notes 4.6.2.1 Previous Component Approval X

4.6.2.2 Simulated Operational Component Tests

X

4.6.2.2.1 Component Acceptance Test X X 4.6.2.2.2 4.6.2.2.2.1 4.6.2.2.2.2 4.6.2.2.3

Component Test Procedures Component Test Cycles (Excluding Ignition System) Ignition System Test Cycles Order of Testing

X X

X X

4.6.2.2.3.1 Low Lubricity Fuel Test X 4.6.2.2.3.2 High Temperature X 4.6.2.2.3.3 Room Temperature & Contamination X 4.6.2.2.3.4 4.6.2.2.3.5

Low Temperature Temperature & Vibration Cycling

X X

4.6.2.2.3.6 Engine Fuel System Cavitation Endurance

X

4.6.2.2.3.7 Fuel System Bubble Ingestion X 4.6.2.2.4 Component Acceptance Test or

Recalibration X X

4.6.2.2.5 Component Test Completion X Z 4.6.2.3 Environmental Component Tests X 4.6.2.3.1 4.6.2.3.2

Component Calibrations Order of Testing

X X

4.6.2.3.3 4.6.2.3.4

Component Test Success Criteria Component Test Procedures

X X X

Z

4.6.2.3.4.1 Temperature – High, Low, Transient X 4.6.2.3.4.2 Vibration – Airframe and Engine X 4.6.2.3.4.3 Impact (Shock) X 4.6.2.3.4.4 Gunfire Shock X 4.6.2.3.4.5 Sustained Acceleration X 4.6.2.3.4.6 Low Pressure (Altitude) X 4.6.2.3.4.7 Rain X 4.6.2.3.4.8 Explosive Atmosphere X 4.6.2.3.4.9 Fungus X 4.6.2.3.4.10 Humidity X 4.6.2.3.4.11 Salt Fog X 4.6.2.3.4.12 Sand and Dust X 4.6.2.3.4.13 Contaminated Fluids X 4.6.2.3.4.14 Proof Pressure X 4.6.2.4 Special Component Tests X 4.6.2.4.1 Oil Reservoir X Y 4.6.2.4.2 Accessory Drive X Z 4.6.2.4.3 Alternator Test X Y

270

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APPENDIX C (continued)

REF. - 6.3

REQUIREMENTS FOR QUALIFICATION TEST (QT) RATING

Para No. Requirements Analysis Demonstration Test Notes 4.6.2.4.4 Heat Exchanger, Proof Pressure X Z

4.6.2.4.5 Fire Test X Y 4.6.2.4.6 Output Shaft Static Torque Test X 4.6.2.4.7 Overheat Test X 4.6.2.4.8 PMA Neutral Short-to-Ground Test X 4.6.2.4.9 Overspeed and Containment X Z 4.6.2.4.10 Burst Pressure X 4.6.2.4.11 Pressure Cycling X 4.6.2.4.12 Proof Pressure X 4.6.2.4.13 Pressure Wash Test X Y 4.6.2.4.14 Ignition System Fouling X 4.6.2.4.14.1 Carbon Fouling X 4.6.2.4.14.2 Water Fouling X 4.6.2.4.15 Electrical Loads Analysis X 4.6.2.4.16 Short Circuit Protection X 4.6.2.4.17 Data Bus Specification Compliance X 4.6.2.5 Software Verification X 4.6.2.5.1 Engine Testing (Software FQT) X 4.6.2.5.2 Flight Testing (Software FQT) X 4.6.2.6 Complex Power Interrupt Test

(Closed Loop) X

4.6.2.7 EMS Lifing Algorithms X 4.6.2.8 Common Mode Multiple Signal

Failure Test. X

4.6.3 Altitude Test (Engine) X 4.6.3.1 Altitude Engine Calibration X 4.6.3.2 Altitude Test Procedure X 4.6.3.3 Altitude Test Completion X 4.6.4 Engine Environmental and Ingestion

Test X

4.6.4.1 Low and High Temperature Starting and Acceleration Test

X

4.6.4.2 Environmental Icing Test X Y 4.6.4.3 Corrosion Susceptibility Test X Z 4.6.4.4 Bird Ingestion Test X Z 4.6.4.5 Foreign Object Damage Test X 4.6.4.6 Ice Ingestion Test X 4.6.4.7.1 Fine Sand Ingestion Test X Z 4.6.4.7.2 Course Sand Ingestion Test X Z 4.6.4.8 Atmospheric Water Ingestion Test X 4.6.4.9 Engine Component Limiting

Temperature Test X

4.6.4.10 Noise Survey X

271

AV-E-8593A AQP

APPENDIX C (continued)

REF. - 6.3

REQUIREMENTS FOR QUALIFICATION TEST (QT) RATING

Para No. Requirements Analysis Demonstration Test Notes

4.6.4.11.1 Exhaust Smoke Emission X 4.6.4.11.2 Invisible Exhaust Mass Emissions X 4.6.4.12 Attitude Test X 4.6.4.13 Loss of Oil X S,Z 4.6.4.14 Electromagnetic Environmental

Effects (E3) X

4.6.4.14.1 Electromagnetic Interference (EMI) X 4.6.4.14.4 Lightning X 4.6.4.15 IR Signature X 4.6.5 Engine Characteristics and Fuel

Tests X X

4.6.5.1 Starting Torque X 4.6.5.5 Emergency Fuel Test X Y 4.6.6 Structural Tests X 4.6.6.1 Emergency Power Demonstration X Z 4.6.6.2.1 Low Cycle Fatigue Component Tests X Z 4.6.6.2.2 Low Cycle Fatigue Engine Tests X M,Z 4.6.6.3 Containment X Z 4.6.6.4.1 Overspeed X 4.6.6.4.2 4.6.6.4.3

Overtemperature Disk Burst

X X

Z

4.6.6.5.1 Engine Static Load Test X 4.6.6.6 Vibration and Stress Tests X 4.6.6.6.1 Compressor Strain Test X 4.6.6.6.2 Gas Generator Turbine Strain Test X 4.6.6.6.3 Power Turbine Strain Test X 4.6.6.6.4 Gas Generator Rotor Bearing

Evaluation X

4.6.6.6.5 Power Turbine Rotor Bearing Evaluation

X

4.6.6.6.6 Engine Vibration Survey X 4.6.6.6.7 ”Stinger” Rig Vibration Test X 4.6.6.7 Gyroscopic Test X Y 4.6.6.8 Engine Overtemperature Control X 4.6.6.9 Engine Overspeed Control System

Test X Z

4.6.6.10 Main Shaft Bearings Assurance Testing

X

4.6.6.11 Gear Resonance Test X 4.9.2.1 300 Hour Benchmark Endurance

Test X Z

272

AV-E-8593A AQP

APPENDIX C (continued)

REF. - 6.3

REQUIREMENTS FOR QUALIFICATION TEST (QT) RATING

Para No. Requirements Analysis Demonstration Test Notes

4.9.2.2 Preliminary 3750 Cycle Low Cycle Fatigue test

X Z

AV-E-8593 AQP

APPENDIX C (continued)

REF. - 6.3

REQUIREMENTS FOR QUALIFICATION TEST (QT) RATING

273

NOTES:

A. A report shall be submitted to the Using Service prior to the initiation of the Qualification Endurance Test.

B. An EMI control plan, in accordance with ADS-37-PRF, as required by MIL-STD-461F, shall be submitted to the Using Service prior to the start of the official EMI/EMC QT testing.

C. Matched or selected fit parts shall be identified and listing shall be provided to the Using Service prior to QT.

D. The strength and life analysis report shall be updated prior to QT. E. Analysis results shall be included in the strength and life analysis. F. A report verifying the heat balance analysis shown in the engine system specification shall be submitted to the Using Service prior to the initiation of the qualification endurance test. G. An engineering evaluation test report shall be submitted prior to the completion of the QT for

each applicable requirement or test paragraph for each qualification test. H. Pretest data shall be submitted to the Using Service for approval prior to the initiation of each

test or demonstration required for QT. I. Test reports shall be submitted to the Using Service for approval following completion of each

test or demonstration required for QT. J. Summary reports shall be submitted to the Using Service following completion of QT. K. Engine weight shall be measured, before the engine has been serviced with fuel or oil, prior

to initiation of calibrations. L. Shall be performed prior to the Engine Calibration of 4.6.1.2.2. M. The test cycle shall be defined by the contractor and submitted to the Using Service in the

pretest data. N. A report shall be provided to the Using Service 6 months prior to the endurance portion of the

QT. O. The vibration and stress analysis report shall be updated prior to QT. P. Not Used. Q. Revised correlation studies shall be presented to the Using Service for approval prior to the

initiation of the QT. R. Component drawings, as specified in 4.6.2.4.2, shall be submitted to the Using Service prior

to the initiation of the component QT. S. This test shall be required only if significant components have been changed from the

configuration of PFR. T. The results of the customer bleed air analysis and the method and test apparatus used

shall be detailed in the test report. U. The analysis results shall be included in the vibration and stress analysis. V. When engine manufacturer's documents are used for materials and processes, such

documents shall be subject to review by the Using Service prior to the start of the QT. W. Module replacement limits shall be provided to the Using Service prior to the start of the QT

and shall be incorporated into the ESS by SCN action. Y. A “dirty inspection” IAW 4.6.1.5 Z. A “dirty inspection” and “clean inspection” IAW 4.6.1.5

AV-E-8593 AQP

APPENDIX D

REF. - 6.3

REQUIREMENTS FOR OPERATIONAL CAPALILITY RELEASE (OCR)

274

Para No. Requirements Analysis Demonstration Test Notes 4.8.1 Accelerated Endurance Test X 4.8.1.1.1 Engine Dry Mass X A 4.8.1.1.2.1 Component Calibration X B 4.8.1.1.2.2 Engine Calibration X 4.8.1.3 Endurance Test Procedure X 4.8.1.1.3.1 Starts X 4.8.1.1.3.2 Contingency Power Run X 4.8.1.1.4.1 Engine Recalibration X 4.8.1.1.4.2 Component Recalibration X 4.8.1.1.5 Engine Disassembly and Inspection X D 4.8.1.1.6 Endurance Test Competition X 4.8.2 Accelerated Mission test X 4.8.2.1.1.1 Engine Dry Mass X A 4.8.2.1.2.1 Component Calibration X C 4.8.2.1.2.2 Engine Calibration X 4.8.2.1.3 Endurance Test Procedure X 4.8.2.1.4.1 Engine Recalibration X 4.8.2.1.4.2 Component Recalibration X 4.8.2.1.5 Engine Disassembly and Inspection X E 4.8.2.1.6 Endurance Test Competition X 4.8.3 Reliability Evaluation Testing X NOTES:

A. Engine weight shall be measured, before the engine has been serviced with fuel or oil, prior to initiation of calibrations.

B. Shall be performed prior to the Engine Calibration of 4.8.1.2.2. C. Shall be performed prior to the Engine Calibration of 4.8.2.2.2. D. A “dirty inspection” and “clean inspection” IAW 4.8.1.1.5 E. A “dirty inspection” and “clean inspection” IAW 4.8.2.1.5

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited. Other requests for this document shall be referred to FCDD-AE, 4488 Martin Road, Redstone Arsenal, AL 35898-5000

APPENDIX E

STRENGTH, LIFE, AND CREEP ANALYSIS REPORT

E.1 SCOPE

E.1.1 Scope.This appendix provides details regarding the information required in the Strength, Life, andCreep Analysis Report. This appendix is a mandatory part of this AQP. The informationcontained herein is intended for compliance.

E.2 APPLICABLE DOCUMENTSThe applicable documents in section 2 of this AQP apply to this appendix.

E.3 DEFINITIONS

E.3.1 General.Definitions used herein and in the strength, life, and creep analysis report shall be as specifiedbelow.

E.3.1.1 Acronyms used in this appendix.

The acronyms used in this appendix are defined as follows:

a. EMS - Engine Monitoring System

b. FEA - Finite Element Analysis

c. GGT - Gas Generator Turbine

d. LCF - Low Cycle Fatigue

e. MATS - Maximum Allowable Transient Rotor Speed

f. PT - Power Turbine

g. SOW - Statement of Work

E.3.1.2 Maximum Allowable Transient Rotor Speed.The MATS is as defined in the engine specification.

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276

E.4 GENERAL REQUIREMENTS.

E.4.1 General. This appendix outlines a list of requirements that shall be presented in the Strength, Life, and Creep Analysis Report. The Strength, Life, and Creep Analysis Report shall conform to the requirements this AQP, as well as the requirements in this appendix. If other parameters or methods not detailed in this appendix are necessary for independent verification of the design, these shall be provided and discussed as well. This report shall contain ample detail to allow for a thorough understanding of the engine structural design.

E.4.2 Content. The Strength, Life, and Creep Analysis Report shall contain, at a minimum, an introductory section, a configuration section, and an analyses methodology/results section. These sections are detailed in AC.5.

E.4.3 Material Property Curves. All material property curves (e.g., creep/stress rupture, corrosion, erosion, LCF, crack growth, tensile strength, etc.) used in strength and life predictions shall be presented so that accurate values can be obtained (to include headings, legends, and axes). The curves shall be labeled with all relevant information such as heat treatment, grain size, coating, and test temperature. Also, a description of the tests (to include number of tests) and test specimens from which these curves were derived shall be included. Any adjustments for potential differences in component behavior relative to test specimen data shall be explained.

E.4.4 Analysis Models. Finite element and other mathematical models used to calculate steady-state and transient component metal temperatures and stresses shall be presented along with plots of their results. Models and plots shall be clearly labeled so that all pertinent characteristics can be identified. All material properties, applied loads, boundary conditions, and cycle conditions used in each analysis shall be presented. An accuracy assessment of each model and plot shall also be included. All finite element analytical models, input files, and results files used to calculate temperatures and stresses in the strength and life analyses shall be provided to the Government using electronic media at the time of Strength, Life, and Creep Analysis Report submittal. If the Contractor utilizes their own internal codes for any of the necessary analyses, deliverables shall be negotiated prior to contract award. For example, if a Contractor uses an internal code for their thermal analysis work, the delivered stress models shall include a thermal results input file that is readable by the code used for the stress analyses.

E.4.5 Points of Life. Engine specific points of life shall be used in the strength, life, and creep analyses. See Paragraph 3.3.8.6.1 for details.

E.4.6 Delivery. The Strength, Life, and Creep Analysis Report shall be submitted in electronic format at the specified delivery times in the SOW.

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277

E.5 DETAILED REQUIREMENTS.

E.5.1 Introduction.

E.5.1.1 Purpose and Scope. The introduction shall contain the purpose and scope of the Strength, Life, and Creep Analysis Report.

E.5.1.2 Summary. A detailed summary shall be included in the introduction that shall contain the following items: a. An overview of the engine configuration including an engine cross-section with station numbers. b. An overview of the strength, life, and creep requirements, as stated in the engine specification. c. An overview of the results of the analyses to include a table of life summaries for each component analyzed. The table shall list each component along with its corresponding LCF life, creep/stress rupture life, oxidation/corrosion life, and burst margin. The table should be usable as a quick reference guide. d. An explanation of how margins of safety are determined.

E.5.2 Configuration.

E.5.2.1 Parts Listing. A parts listing shall be provided including part number, part name, material, and coating of each component analyzed for the report.

E.5.2.2 Mission Profile. A graphical and/or tabular representation of the structural design mission(s) (LCF cycle/duty cycle) to be used in the analyses shall be provided. The tabular representation shall clearly define all timepoints used in the analysis and shall define all transient temperatures, transient rotor speeds, and overshoots assumed.

E.5.2.3 Cycle Parameters. A table summarizing the critical engine cycle parameters corresponding to the operating conditions of the structural design mission(s) shall be provided. In addition, the engine performance digital computer program output files used for each of the operating conditions of the structural design mission(s) shall be provided.

E.5.2.4 Secondary Flow. Figures showing secondary flow and pressure/temperature maps shall be provided at the structural design point(s) identified in the engine specification. A description shall be presented

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278

that details how the secondary flow scheme is arranged in addition to cooling intentions. Secondary cavity flow diagrams shall be included. Secondary flow test substantiation shall be provided and shall include data from thermal survey tests. A figure shall be provided detailing the flow value predictions correlated with actual engine test data.

E.5.3 Analyses Methodology/Results. The following analyses shall be completed for each applicable component at the structural design point(s) identified in the engine specification. The Strength, Life and Creep Analysis Report shall contain details of the methodology and results of each analysis. Verification of these results through component or engine testing shall be presented including all details of the tests such as test performed, test conditions, facilities, hardware, instrumentation, and the results obtained.

E.5.3.1 Thermal Analysis. A detailed description of the thermal analysis methodology used to calculate metal temperatures shall be provided. The description shall include all assumptions, material properties, cycle conditions and boundary conditions used in the analysis. All 2D/3D FEA models and plots of results shall be presented (see C.4.4 for details). Detailed information regarding heat transfer coefficients and correlations used shall be provided. Transient metal temperatures and thermal gradient time history responses for each component analyzed shall be presented.

E.5.3.2 Stress Analysis. A detailed description of the methodology used to calculate component stresses, including major and minor cycle stresses calculated using the LCF test cycle shown in AV-E-8593, shall be provided. The description shall include thermal model mapping, 2D/3D FEA model descriptions, all material properties, stress concentration factor determinations, loading conditions, and any other assumptions used. All 2D/3D FEA models and plots of results shall be presented, in accordance with C.4.4. The stresses obtained from the analysis shall be presented in the form of tables C-I through C-VI for all critical locations. These stress tables include generalized stress component (tensor) tables and principal/equivalent stress tables for both major and minor cycle component stresses. The generalized stress tensor tables shall contain all components of both mechanical and thermal stress for major and minor cycles, the corresponding metal temperature, a description of the critical location(s), and the corresponding mission time point. The tables presented reference a cylindrical coordinate system. However, other coordinate systems may be used as necessary. These values shall then be used to determine principal and equivalent stresses, both maximum and minimum.

E.5.3.3 Strength Analysis. A detailed description of the ultimate and limit load analysis methodology and requirements shall be provided. The description shall include the operating conditions at which limit and ultimate loads are calculated. All 2D/3D FEA models and plots of results shall be presented, in accordance with C.4.4. Strength results shall present margins of safety for each component location considered, under both limit and ultimate loading conditions. Buckling/elastic stability

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279

shall be addressed. A table detailing how each component meets or exceeds engine specification strength requirements shall be provided.

E.5.3.4 Low Cycle Fatigue Analysis. A detailed description of the LCF lifing methodology shall be presented. The description shall include all assumptions, material properties, LCF material curves (see C.4.3), and equations used in life calculations, sensitivity studies, and damage fraction determinations. Sufficient detail shall be provided so that the lifing methodology can be fully understood. LCF life prediction shall be based on an allowable fatigue curve calculated from mean life to crack initiation reduced by a reliability factor of three standard deviations based on applicable material scatter and shall account for stress concentrations as applicable to the design. The LCF life and its sensitivities to increases in temperature and speed shall be presented (in the form of tables C-VII and C-VIII respectively) for all critical locations. Any life adjustment factor (based on test/experience, surface treatments, etc.), if used, shall be identified. Basis for life adjustment factors shall be discussed in detail. Details regarding all critical locations and how they were determined shall be presented. Cumulative LCF damage shall be calculated using an appropriate damage theory. Damage tables illustrating damage due to major and minor cycles shall be created and presented in the form of table C-IX. A table shall be provided detailing how the component meets or exceeds engine specification LCF life requirements.

E.5.3.5 Fracture Mechanics Analysis. A detailed description of any fracture mechanics analysis used shall be presented. A description of methods used to predict crack propagation from material voids or defects shall be presented. Data for material defect distribution shall be provided.

E.5.3.6 Creep/Stress Rupture Analysis. A detailed description of the creep/stress rupture analysis methodology shall be presented. The description shall include all assumptions and material properties and curves (see C.4.3) used in the analysis. The description shall include a specific definition of what defines creep failure. If components do not require creep/stress rupture analysis, rationale shall be provided detailing why it is not necessary. For those parts subject to creep/stress rupture, including all GGT and PT blades, all critical locations shall be presented along with the lifing estimates as detailed in table AC-X. Creep buckling shall be addressed for combustor liners. A table shall be provided detailing how the component meets or exceeds engine specification creep/stress rupture requirements.

E.5.3.7 Burst Margin Analysis. A detailed description of the burst margin analysis methodology shall be presented. Both hoop and radial burst shall be evaluated and presented. The description shall include all assumptions and material properties and curves (see C.4.3) used in the analysis. All critical/failure locations shall be presented. Results of the burst margin analysis shall be summarized and presented in the form of table C-XI. Other variables that are needed/used by the Contractor shall be presented in similar format to that of table C-XI to allow for independent Government verification. A table shall be provided detailing how the component meets or

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280

exceeds engine specification burst margin requirements. For those components not requiring burst margin analysis, rationale shall be provided.

E.5.3.8 Oxidation/Corrosion Analysis. A detailed description of the oxidation/corrosion analysis methodology shall be presented. The description shall include all assumptions and material properties and curves (see C.4.3) used in the analysis. All critical locations and lifing estimates shall be provided in the form of table C-XII along with a summary of the oxidation/corrosion analysis results. A table shall be provided detailing how each component meets or exceeds the engine specification oxidation/corrosion requirements. Requirements may be satisfied through demonstrated field experience with prior approval from the Government.

E.5.3.9 Containment Analysis. A detailed description of the containment analysis methodology shall be presented. The description shall include all assumptions and material properties used in the analysis. A summary of the containment analysis results shall be presented along with a table, in the form of table C-XIII, detailing the containment capacity of each applicable component. A table shall be provided detailing how engine specification containment requirements are met or exceeded.

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TABLE E-I. Major Cycle Maximum Component Stresses.

ID Description σrad σtan σax τrt τta τra σrad σtan σax τrt τta τra

Location Maximum Stress (Thermal) Maximum Stress (Mechanical) Metal Temp ºC (ºF)

Time Point (sec)

CommentsMPa (ksi) MPa (ksi)

TABLE E-II. Major Cycle Minimum Component Stresses.

TABLE E-III. Major Cycle Principal/Equivalent Stresses.

ID Description σ11 σ22 σ33 σequiv σ11 σ22 σ33 σequiv

LocationPrincipal/Equivalent Stresses

Maximum - MPa (ksi) Minimum - MPa (ksi)

ID Description σrad σtan σax τrt τta τra σrad σtan σax τrt τta τra

Location Minimum Stress (Thermal) Minimum Stress (Mechanical) Metal Temp ºC (ºF)

Time Point (sec)

CommentsMPa (ksi) MPa (ksi)

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TABLE E-IV. Minor Cycle Maximum Component Stresses.

ID Description σrad σtan σax τrt τta τra σrad σtan σax τrt τta τra

Location Maximum Stress (Thermal) Maximum Stress (Mechanical) Metal Temp ºC (ºF)

Time Point (sec)

CommentsMPa (ksi) MPa (ksi)

TABLE E-V. Minor Cycle Minimum Component Stresses.

ID Description σrad σtan σax τrt τta τra σrad σtan σax τrt τta τra

Location Minimum Stress (Thermal) Minimum Stress (Mechanical) Metal Temp ºC (ºF)

Time Point (sec)

CommentsMPa (ksi) MPa (ksi)

TABLE E-VI. Minor Cycle Principal/Equivalent Stresses.

ID Description σ11 σ22 σ33 σequiv σ11 σ22 σ33 σequiv

LocationPrincipal/Equivalent Stresses

Maximum - MPa (ksi) Minimum - MPa (ksi)

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TABLE E-VII. Low Cycle Fatigue Life.

ID Description - 3 σ Nominal - 3 σ Nominal - 3 σ Nominal

Stress Ratio

(Min/Max)

Low Cycle Fatigue Life(Cycles)

Location

Major Cycle (0-MRP-0) Minor Cycle (GI-MCP)Life

Adjustment Factor

Low Cycle Fatigue Life(Mission Cycles)Stress

Range MPa (ksi)

Stress Ratio

(Min/Max)

Low Cycle Fatigue Life(Cycles)

Stress Range

MPa (ksi)

TABLE E-VIII. Low Cycle Fatigue Life Sensitivities.

ID Description - 3 σ Nominal - 3 σ Nominal

LocationMajor Cycle Sensitivities

Life Sensitivity to Variation in Speed Life Sensitivity to Variation in Temperature+ 5% + 50 ºF

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TABLE E-IX. Low Cycle Fatigue Life Damage Fractions.

Condition 0 GI * FI * * * MCP IRP MRP CRP * *

Percent 0

Condition Percent RPM 0

MATS

*

*

CRP

MRP 1.0000E0

IRP

MCP

*

*

*

FI

*

GI

Minimum GGT Speed

Max

imum

GG

T Sp

eed

NOTES: 1. Damage values shall be presented using scientific notation with 5 significant digits (i.e., 0.0003764210 shall be presented as 3.7642E-4). 2. Unit damage of 1.0000E0 for the major cycle shall be defined at MRP. 3. Prudent scaling practices are acceptable. 4. (*) values shall be engine specific agreed upon condition points.

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TABLE E-X. Creep/Stress Rupture Life.

Overall Time (per Spec Mission)

Nominal Life -3 σ Life

ºC ºF MPa ksi hrs hrs hrsCRPMRPIRPMCP60% MCP40% MCPGI

Damage Fraction per Spec MissionNominal Creep/Stress Rupture Life, Mission Hours-3 σ Creep/Stress Rupture Life, Mission Hours

Power Condition Material Location

Metal Temperature Stress

Time %

Damage Fraction at

Power Condition

TABLE E-XI. Burst Margin.

Burst Margin

ºC ºF MPa ksi MPa ksi %

Disks

Spacers

Seal Plates

Com

pres

sor Disks

Spacers

Impellers

Turb

ine

GG

T &

PT

Component MaterialBulk Metal

Temperature-3σ Ultimate Tensile

StrengthAverage

Hoop/Radial Stress

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TABLE E-XII. Oxidation Life.

Oxidation Life at Power

Time at Power per Spec Mission

ºC ºF hrs hrsCRPMRPIRPMCP60% MCP40% MCPGI

Damage Fraction per Spec Mission

Oxidation Life, Mission Hours

Operating Condition

Material LocationMetal

Temperature

Damage Fraction at

Power Condition

TABLE E-XIII. Containment.

Typical Elongation

ºC ºF MPa ksi MPa ksi % mm in J in-lbf

Total Containment CapacityFragment Kinetic EnergyMargin of Safety (%)

Layer #

Component MaterialTemperature

-3σ Ultimate Tensile Strength

-3σ 0.2% Yield Strength

ThicknessContainment

Capacity

AV-E-8593 AQP

Appendix F

APPENDIX F

VIBRATION AND STRESS ANALYSIS REPORT

F.1 SCOPE

F.1.1 Scope.This appendix provides details regarding the information required in the Vibration and Stress Analysis Report. This appendix is a mandatory part of this AQP. The information contained herein is intended for compliance.

F.2 APPLICABLE DOCUMENTSThe applicable documents in section 2 of this AQP apply to this appendix.

F.3 DEFINITIONSThe definitions in section 6 of this AQP apply to this appendix.

F.4 GENERAL REQUIREMENTS.

F.4.1 General.This appendix outlines a list of requirements that shall be presented in the Vibration and Stress Analysis Report. The Vibration and Stress Analysis Report shall conform to the requirements in this AQP, as well as the requirements in this appendix. This report shall contain ample detail to allow for a thorough understanding of the engine structural design.

F.4.2 Content.The Vibration and Stress Analysis Report shall contain, at a minimum, an introductory section and a section detailing the vibration and stress characteristics of each part analyzed. These sections are detailed in D.5.

F.4.2.1 Stress/Thermal Analyses.Stress/Thermal Analyses from the Strength, Life and Creep Analysis Report may be referenced as needed.

F.4.3 Material Property Curves.Material property curves (e.g., HCF, etc.) used in life prediction shall be presented so that accurate values can be obtained (to include headings, legends, and axes). The curves shall be labeled with all relevant information such as heat treatment, grain size, coating, and test temperature. Also, a description of the tests (to include number of tests) and test specimens

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited. Other requests for this document shall be referred to FCDD-AE, 4488 Martin Road, Redstone Arsenal, AL 35898-5000

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from which these curves were derived shall be included. Any adjustments for potential differences in component behavior relative to test specimen data shall be explained.

F.4.4 Analysis Models. Finite element and other mathematical models used to calculate steady-state and transient component metal temperatures, stresses, resonant frequencies, potential excitation sources, and response characteristics shall be presented along with plots of their results. Models and plots shall be clearly labeled so that all pertinent characteristics can be identified. All material properties, applied loads, boundary conditions, and cycle conditions used in each analysis shall be presented. An accuracy assessment of each model and plot shall also be included. All finite element analytical models, input files, and results files used for the vibration and stress analyses shall be provided to the Government using electronic media at the time of Vibration and Stress Analysis Report submittal. If the Contractor utilizes their own internal codes for any of the necessary analyses, deliverables shall be negotiated prior to contract award. For example, if a Contractor uses an internal code for their thermal analysis work, the delivered stress models shall include a thermal results input file that is readable by the code used for the stress analyses.

F.4.5 Delivery. The Vibration and Stress Analysis Report shall be submitted in electronic format at the specified delivery times as stated in the SOW.

F.5 DETAILED REQUIREMENTS.

F.5.1 Introduction. The Vibration and Stress Analysis Report shall contain an introductory section that shall include the items discussed in this section.

F.5.1.1 Purpose and Scope. The introduction shall contain the purpose and scope of the Vibration and Stress Analysis Report.

F.5.1.2 Summary. A detailed summary shall be included in the introduction that shall contain the following items: a. A detailed description of vibration/stress analysis approach. b. An engine cross-section illustrating all applicable component sections and all blade/vane counts. c. An overview of the results of the analyses to include a table of all parts analyzed with their results. d. A summary table of engine strain gage testing in the form of table D-I, if applicable. e. An explanation of how margins of safety are determined (i.e., based on Goodman diagrams, etc.).

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F.5.2 General Component. This section presents an outline that shall be used for each component analyzed. Each component shall be analyzed at the structural design point(s) identified in the engine specification. Any additional requirements detailed in the engine specification shall also be presented.

F.5.2.1 Component Description. A detailed component description for each component analyzed shall be included. The component description shall contain the following: a. A detailed description of the component material and any coatings used. b. Major sources of component excitation.

F.5.2.2 Vibration and Stress Analysis. A section detailing the vibration and stress analysis shall be included. This section shall include the following: a. A description of the vibration analysis including the approach taken, all assumptions used, methods for calculation of natural frequencies, applicable engine specification requirements, and results/conclusions. HCF considerations shall be included where applicable (see D5.2.5). b. All 2D/3D FEA plots used in the modal analysis (see D.4.4). This shall include all assumptions, loading conditions, operating speeds, frequencies, and constraints used. c. All Campbell diagrams. These shall be presented in such a way that it can be determined if any of the engine excitation frequencies correspond to the fundamental natural frequencies. All mode shape curves on the diagrams shall be labeled properly. All stress levels shall be clearly defined. d. All excitation frequencies. These shall be clearly identified. Each mode shape label shall be described in detail (i.e., provide a key that physically describes each label). e. A graphical representation of strain gage locations used as determined by analyses.

F.5.2.3 Component Testing. A section detailing any component testing such as acoustic ring or holographic tests shall be included with results. A table that includes a comparison of the measured frequencies from component testing versus those determined in modal analyses shall be provided.

F.5.2.4 Engine Testing. A thorough description of any engine or rig testing that is used to verify the design shall be provided. It shall include all operating parameters and cycle conditions. The following shall also be provided:

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a. A graphical representation and discussion of strain gage locations used on the component as determined by analyses. Gages shall be clearly labeled and numbered. b. A table of individual strain gage results as detailed in table D-II. c. Information regarding any other instrumentation used, such as thermocouples, etc. This shall include graphical representation and discussion of locations and a table of results. d. Any modifications (such as gage factors) to test data used as input for HCF analysis.

F.5.2.5 High Cycle Fatigue Assessment. An assessment detailing the component HCF life shall be provided. This shall include a thorough discussion of HCF considerations, all fatigue property diagrams (Goodman, etc.), and HCF conclusions.

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Appendix F

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited. Other requests for this document shall be referred to FCDD-AE, 4488 Martin Road, Redstone Arsenal, AL 35898-5000

TABLE F-I. Summary of Component Vibration and Stress Analysis.

ºC ºF MPa ksi MPa ksi MPa ksi

Allowable Vibration Stress

Margin of Safety (%)

Component Location Material Temperature Critical Mode

Excitation Source

Measured Strain

(µmm/mm)

Vibration Stress Mean Stress

TABLE F-II. Vibratory Strain Survey.

ºC ºF MPa ksi MPa ksi MPa ksi MPa ksi MPa ksiGage # Gage

LocationFrequency

(Hz)Mode

Measured Strain

(µmm/mRPM Excitation

OrderTemperature Alternating

StressMean Stress Ultimate

Tensile Endurance

LimitAllowable

Alternating Margin

of Safety

AV-E-8593 AQP

Appendix G

Software/Programmable Hardware Development Requirements

G.1. Introduction.This appendix to the AV-E-8593 AQP describes the processes, products, and procedures required to ensure the development of safe aircraft propulsion system software which is capable of obtaining an Army Airworthiness Authority SAQ (statement of airworthiness qualification). For the purposes of this document, the computer instructions and computer data associated with firmware or complex programmable electronics shall be considered as software.

This AQP follows a similar process as outlined in RTCA/DO-178C in that it is centered around the demonstration of specific software development objectives that depend on the level of criticality of the software. Unlike the DO-178C, the Army considers these “objectives” to be “requirements”. Another difference is the fact that there are quite a few more Army requirements than the DO-178C Annex A objectives. A complete list of the Army software requirements is summarized in Table 4.

Table 4 specifies the requirements for each software criticality level and the independence for each software criticality level. The process for determining the software criticality, herein referred to as the “SHCI Assignment Level”, is described in Section 3.

Section 2 provides more description of the aforementioned Table 4 requirements. It also explains in which phase of the software lifecycle the requirements are to be accomplished. These phases are shown graphically in Figure 1 below.

Sequencing through the software lifecycle in the correct order with the proper inputs is critical to developing and verifying airworthy software. All criteria are defined within each process section (Section 2). The Contractor shall ensure that the transition criteria are completed in the correct order.

RRMSASDALSARSPSSSA

RRMSTRDAL

RRMSVDDAL

RRMSTDDALSSA

RRMSDDSTDSSADALSAR

RRMSRSDALSSA

PSACSCMPSDP

SQAPDALSTPRRM

EMSAQS

SWPPSSDDFHADALIMSSSA

System Process Requirements

(2.1)

Software Planning Process

(2.2)

Software Requirements

Process(2.3)

Software Design Process

(2.4)

Software Coding Process

(2.5)

Software Integration

Process(2.6)

Software Test Verification

Process(2.7)

Software Test Approval Process

(2.8)

SRR PPR SWRR CDRPDR IR TRR TReR FQR

2.1.1 2.2.1 2.3.1 2.4.1 2.5.1 2.6.1 2.7.1 2.2.1

ProcessArtifacts

Figure 1. Aviation Propulsion System Software Lifecycle

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited. Other requests for this document shall be referred to FCDD-AE, 4488 Martin Road, Redstone Arsenal, AL 35898-5000

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StatementOf

Work (SOW)

AV-E-8593AQP

EngineModel

Specification(EMS)

InterfaceControl

Documents(ICD)

System/SubsystemDesign

Description(SSDD)

SoftwareDesign

Description(s)(SDD)

InterfaceDesign

Description(s)(IDD)

SoftwareRequirements

Specification(s)(SRS)

InterfaceRequirements

Specification(s)(IRS)

HardwareRequirementsSpecification

(HRS)

ProgrammableHardware

RequirementsDocument

(HRD)

ProgrammableHardware DesignRepresentation

Data(HDD)

Top-Level Requirements

System-Level Requirements

Hardware/Software Requirements

Detail Requirements

Requirements Flow Down Process

Figure 2. Requirements Flow Down Process

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G.2. Software Lifecycle Processes.

G.2.1 System Process. The Contractor shall develop a Software Safety Program Plan (SwSPP), that satisfies their System Safety Program Plan (SSPP) and government provided System Safety Management Plan (SSMP). The Contractor shall work with the PMO to deliver Airworthiness Qualification Specification inputs that satisfy this AQP. The Contractor shall analyze the system specification, and develop system requirements that meet the engine model specification. The Contractor shall architect the system and allocate the requirements to hardware and software elements in the System/Subsystem Design Description (SSDD). The Contractor shall implement a configuration management process to baseline, archive, and release all system documents. The Contractor shall generate a system level Functional Hazard Analysis (FHA) to identify the potential functional failures of the system and trace those hazards to the government provided Aircraft Level FHA hazards identified for the system. This shall be contained in system level FHA. System Safety requirements identified in the engine model specification and SSDD shall be traced to the system level FHA. The Contractor shall perform a safety analysis to allocate safety requirements from the system level FHA to subsystems, equipment, and software. For catastrophic or critical system hazards (level A/B), a system level Fault Tree Analysis (FTA) or Dependency Diagrams shall be used to demonstrate how the proposed architecture mitigates each hazard and determines safety requirements for lower level items. The allocation of safety requirements for marginal hazards can be accomplished by qualitative discussion or by using other modeling techniques depending upon system complexity. Negligible hazards only require discussion of separation from other more critical functions or hazards. The safety allocation analysis shall determine the failures that contribute the hazards in the FHA and be delivered as the System Safety Analysis (SSA). The Contractor shall propose the Software Hazard Criticality Index (SHCI) for all software in the system and be delivered as part of SSA. The SHCI shall be approved by the PMO and AED. The Contractor shall develop a Hazard Tracking System to document and track hazards and their controls (as identified in the system level FHA) thus providing an audit trail of hazard resolutions. The Contractor shall produce Software Hazard Analysis Tracking reports and include in the Hazard Tracking System (HTS). The HTS shall include:

• Description of each hazard to include the systems, software (and assigned SHCI), and equipment associated with the hazard and associated risk assessment.

• Status of each hazard and control.

• Traceability of resolution on each Hazard Log item from the time the hazard was identified to the time the risk associated with the hazard was reduced to a level acceptable to the managing activity.

• Identification of residual risk.

• Action person(s) and organizational element.

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• The recommended controls to reduce the hazard to a level of risk acceptable to the

managing activity.

• The signature of the managing activity accepting the risk and thus effecting closure of the Hazard Log item.

Note that the PMO also has a Hazard Tracking System for all the system hazards. Hazards can only be closed with concurrence of the hazard originator. The Contractor shall conduct a System Requirements Review (SRR).

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2.1.1.2System Safety Assessment (SSA)

2.1.1.3Engine Level FHA

2.4.1.5Safety Assessment Report (SAR)

2.1.1.2.1SSA (Sys Arch)

2.1.1.2.2SSA (SW Arch)

2.1.1.2.3SSA (SW FMEA)

2.1.1.2.4SSA (SW FTA)

2.1.1.2.5SSA Final

2.1Software Hazard Analysis Tracking

Report

*PMO Responsibility

Aircraft Level Functional Hazard Assessment (FHA)* System Safety Management Plan (SSMP)*

System Safety Program Plan (SSPP)

2.1.1.1Software Safety Program Plan (SwSPP)

NOTE: The SSPP and SwSPP can be combined

NOTE: The SAR and SSA can be combined

Figure 3. System Safety Process Flow

G.2.1.1 System Process Artifacts.

G.2.1.1.1 Software Safety Program Plan (SwSPP). The SwSPP shall describe in detail the tasks and activities of the system safety engineering and management program established by the Contractor. It shall also describe the safety, systems, and software engineering processes to be employed to identify, document, evaluate, and eliminate and/or control system hazards to the levels of acceptable risk for the program. If software safety is not addressed in the SSPP, then a Software Safety Program Plan (SwSPP) is required.

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G.2.1.1.2 System Safety Assessment (SSA). The SSA shall be prepared and updated/expanded throughout the development process. This document evolution includes the following stages: System Architecture (Sys Arch), Software Architecture (SW Arch), Software FMEA (SW FMEA), Software Fault Tree Analysis (SW FTA) and Final.

G.2.1.1.2.1 SSA (Sys Arch). The contents shall be tailored to an analysis of the system architecture that shows the systems, subsystems, and equipment that can contribute to each of the FHA failure conditions. ARP4761 Appendix B discusses a Preliminary System Safety Assessment that accomplishes the intended scope of this document.

a. The Sys Arch shall include functional block diagrams that relate the inter-relationship between equipment that implement a systems and/or aircraft function.

b. The Sys Arch shall, at a minimum, contain the FHA(s) used, the architecture

design documents including specifications, diagrams, drawings, lessons learned databases/sources, etc.

c. The Sys Arch contents shall include the selected analysis method, analysis

assumptions, and summary results. For the architectural top-down analysis it is usually acceptable to decompose down to level the equipment LRU level; however, some complex LRUs that incorporate significant amount of redundancy, independence, or internal fault detection, an architectural analysis of the equipment may be required. For catastrophic and critical hazards/failure conditions, FTA is the preferred technique; however, other techniques such as dependency diagrams, allocation matrix, Markov analysis, etc. may be used. Marginal failure conditions architecture safety analysis may include FTAs for complex, integrated functions or qualitative discussion for simple systems. Negligible failure conditions do not require an assessment other than how these systems are isolated from more critical functions. When FTAs are used, the tree diagram shall be supported by a discussion of the fault tree logic development decisions and an event list that defines event name, assigned/allocated failure rate and exposure time, probability calculation method, and reference to developed requirement implementing the equipment failure condition and probability requirements. Common modes that can affect redundant and/or independent items shall be considered and highlighted in results (include software as common modes, shared resources; single event upset response and recovery, etc.) For system modification projects, the system allocation analysis shall be included as either a separate system SSA or combined with the System FHA.

G.2.1.1.2.2 SSA (SW Arch). The SW Arch shall include an analysis of the software architecture that shows how the various software CSCIs can contribute to each of the FHA failure conditions. A combination of ARP 5580 functional FMEA, Interface FMEA, block diagramming, and/or FTA can be used to identify single failures or combination of failures at the lower levels which could cause the FHA failure conditions. ARP 4761 Appendix B discusses a Preliminary System Safety Assessment and ARP4761 Section 4.1.2 Appendix D12 discusses software or errors modeled in FTA that accomplishes the intended scope of this document. In addition, ARP 5580 defines the

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methodology for performing Software Functional Interface, and Detailed FMEA. These analyses can be combined with the SSA (Sys Arch).

G.2.1.1.2.3 SSA (SW FMEA). The SW FMEA shall include a Software functional and Interface FMEA. The functional FMEA is applied to each CSCI during top-level design. The Software Functional FMEA is a systematic analysis of the effects of software errors and possibly some hardware failures on system and software behavior. The Software functional FMEA shall identify failure modes appropriate to the individual functions, assigned to computer software components (CSCs) and to individual modules, and determines their effect on the system.. The Software Functional FMEA is used to identify software architecture vulnerabilities to single point failures and timing dependencies that could lead to unintended effects. All Software Functional FMEA failure modes identified as “Interface” shall be identical to corresponding failure effects in the Software Interface FMEA. A FHA may be used in lieu of a software functional FMEA. The Software Interface FMEA is similar to the Functional FMEA but focuses on the interface between different software and hardware elements. The Software Interface FMEA will include all software/software (CSCI to CSCI) and software/hardware interfaces as defined in the interface requirements specifications. The Software Interface FMEA shall identify failures which relate to data and control coupling in software interfaces between software items (between CSCIs) and software/hardware interfaces. Failure modes in the Software Interface FMEA may be classified as one of the following: Missing Data (Lost message, data loss due to hardware failure), Incorrect Data (inaccurate data, spurious data), Timing of Data (obsolete data, data arrives too soon for processing), Extra Data (data redundancy, data overflow). Safety actions for the Software Interface FMEA shall be produced through software safety requirements, which shall trace back to the functions identified in the FMEA. ARP5580 provides a methodology for interface FMEAs that may be adapted for this document as long as the hardware failure effects include software and software effects. In lieu of a detailed FMEA, an in depth code analysis shall be used

a. The SW FMEA shall, at a minimum, contain the design data used including requirement specifications, description documents, software functional FMEA, software requirements documents, engineering design drawings, etc. All data shall be marked with a revision level to identify specific configurations analyzed.

b. The Software interface FMEA may be combined with the hardware FMEA/FMECA

provided there is clear evidence that software failures are identified and the software response to hardware interface failures are clearly indentified. Detection methods in the FMEA shall also reflect software’s role in detection and missed detections shall be evaluated.

c. There shall be traceability between hardware/software interfaces and the test

cases.

G.2.1.1.2.4 SSA (SW FTA). The SW FTA shall include a qualitative FTA that determines code-level failures that contribute to defined system, equipment, or software CSCIs conditions. A comparison between the results of the Software (Functional and Interface) FMEAs and the FTA shall be performed to ensure all failure modes and significant effects are properly reflected in FTA as “Basic Events”. It is

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anticipated that detailed FTAs would be needed only for specific scenarios or CSCIs. The Contractor shall obtain Airworthiness Authority approval of software FTA top level events prior to initiating the analysis.

G.2.1.1.2.5 SSA (Final). The SSA Final shall include a summary of all safety analyses, supporting engineering analyses and test, SAS, etc. that provide a comprehensive evaluation of the implemented system functions to ensure relevant safety requirements established in the FHAs and architectural safety analysis are met. ARP 4761 Section 3.4 and Appendix C provide discussion and sample content that may be used; however, the analyses, as a minimum, shall include those required in earlier SSA submittals listed above.

a. Data. The data shall include documentation of the configuration analyzed with revision level. As a minimum this shall include engineering drawings, software requirements and design documentation, test reports, engineering analyses that support hazard mitigation or verification (stress analysis, functional testing, reliability prediction, reliability testing, etc.).

G.2.1.1.3 Functional Hazard Assessment (FHA). The FHA shall include analysis techniques for both the aircraft-level FHA and engine system-level FHA. The FHA shall be IAW ARP 4761 Section 3.2 and Appendix A. The FHA shall be used to identify the functions, operating conditions (environment and flight or operating phases), and abnormal/emergency configurations (including other systems used as compensation to limit failure effects) that are used in the FHA. Failure conditions shall address all functions not just those thought to be the most critical. The FHA shall serve as the rationale for the function criticality. Additionally, the FHA shall consider failure conditions for both the detected and undetected loss of function, as well as, detected and undetected malfunctions. The FHA shall provide evidence that all failures were considered and also aids in establishing the criticality of failure condition methods of detection. For changes to existing systems, the original safety analysis may be used to re-evaluate the existing failure conditions for impact to effects and severity, as well as, creation of any new failure conditions. If original safety analysis is not available, then an FHA shall be prepared using defined exchanged functions to establish interfaces to the unchanged portion of the aircraft or system. The FHA shall include the FHA worksheets and a results section that includes a summary listing of failure conditions by severity classification (operating/flight phase shall be part of the failure condition statement if condition has different severities for different operating conditions). The following shall also be included:

a. The resulting safety requirements for each failure condition that shall include but not limited to probability requirements (qualitative and/or quantitative), SHCI, and fault tolerance “no single failure”

b. Follow-on safety analyses (specific techniques) that will be accomplished to verify

compliance to failure condition requirements c. A listing of all crew actions that were used to mitigate (reduce) failure effects d. List of test requirements used as a means to rationalize the severity effect

(example, asymmetric braking classified as minor based on ground testing failure scenario.

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e. Requirements for additions to the aircraft FHA as a result of the system FHA (This

is required when the failure effects of a system failure use another aircraft system to lessen the severity. For example, if engine modulation is used to maintain directional control after a pitch failure, the aircraft FHA needs to add derived failure condition of loss of engine combined with pitch trim failure).

The FHA shall be maintained and updated throughout the development process. The FHA shall provide the initial framework for a listing of hazards and associated top-level safety requirements that require tracking and resolution during program design and development.

a. All Aircraft FHA failure conditions shall be allocated to one or more aircraft systems. This allocation or tracing of requirements may be accomplished in the FHA or as part of the SSA’s Architectural Safety Analysis submittal. Acceptable techniques are FTA, block diagramming (dependency diagrams), matrix cross-referencing failure conditions to all the various systems, and Markov. The technique shall be selected based on the type and complexity of the aircraft or scope of the modification. The Airworthiness Authority shall approve the selected techniques. Additionally, functions and failure conditions added due to engine system-level FHAs or architecture decisions shall be identified as a derived or exchanged function.

b. The Engine System FHA shall identify functions and failure conditions allocated

from the aircraft FHA. Interfaces to other aircraft systems shall be captured as exchanged functions.

G.2.1.1.4 System/Subsystem Design Description (SSDD). The SSDD shall detail the specific performance and operating characteristics of the engine control system, including all functionality, or derived functionality, reflected in the top-level requirements documents, such as engine performance specifications, engine model specifications, interface control documents, statement of work and any other miscellaneous requirements documents. The requirements shall be grouped into three sections: system, hardware, and software. If needed, the software category can be further split into various sub-categories such as application, operating system, partitions, etc. The system requirements shall be derived directly from the top-level requirements, stating the specific requirement document name, paragraph number and text. From this requirement, a detailed set of “derived” system-level requirements shall be defined using shall terminology. Measurable acceptability criteria shall be defined for all performance-related engine or control system requirements. All system requirements defined above, shall be further decomposed into hardware and high-level software requirements, and placed within separate sections of the SSDD. The specific system requirement paragraph number and text shall be repeated and the requirement then decomposed into hardware and/or software requirements, as appropriate. If necessary, the software requirements may further be broken down into separate subsections to accommodate partitioning. For any provided requirements breakdown, the next higher level authorizing requirement shall be identified by paragraph number, title and text, and include for reference.

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The SSDD shall be submitted to the Army Airworthiness Authority for review and approval. This document shall then become the sole source for all subsequent control system requirements definition and the criteria by which the requirements so derived are verified and validated.

G.2.1.1.5 Data Accession List (DAL). The DAL shall provide a medium for identifying contractor internal data, which has been generated by the contractor in compliance with the work effort described in the SOW. The DAL shall be prepared for each milestone review and include all Non-Deliverable Items that may be audited by the government.

G.2.1.1.6 Engine Specification. The engine specification shall include all engine software requirements in AV-E-8593 AQP.

G.2.1.1.7 Integrated Master Schedule (IMS). The IMS shall identify all major milestones including Contractor reviews. The IMS shall identify a Government approved incremental software product delivery schedule for submittals of documentation/code to meet program milestones. The reports shall define the capabilities and limitations of each build.

G.2.1.2 System Process Requirements. Software Safety Program Plan (SwSPP) shall be developed. The SwSPP shall be developed to formally document the processes, tasks, interfaces, and methods to ensure software is taken into consideration from a safety perspective. The SwSPP may be combined with either the SSPP or the PSAC provided a mapping to the SwSPP data requirements is provided.

a. Software Safety Program Plan shall be compliant with the System Safety Program Plan (SwSPP). The SwSPP shall be compliant to the PMO specified SSMP.

b. The engine specification shall be compliant with the AQP. c. System Requirements shall be developed. Each requirement shall be uniquely

identified. System Safety can be affected when underlying system assumptions are not documented. The requirements process shall explicitly capture all underlying assumptions used in the safety analysis until such time the assumption becomes a requirement or constraint.

d. System Requirements shall be allocated to hardware and software. The system

architecture shall be developed. The system requirements in the engine model specification shall be allocated to hardware, software, and firmware system elements. This shall be contained in the System/Subsystem Design Description (SSDD).

e. Configuration Management archive, retrieval, and release processes shall be

established. Configuration management procedures shall be established and documented. All system and safety engineering artifacts shall be put under configuration management.

f. Integrated Master Schedule (IMS) for the system/software development shall be

established. The IMS shall detail the program schedule, events, and criteria for monitoring progress.

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G.2.1.2.1 System Level Functional Hazard Assessment (FHA) Requirements. The FHA process between the aircraft and systems can be iterative. As new system functions are defined they shall be reviewed to ensure no new failure conditions are created at the aircraft level and evaluate potential impacts to existing conditions.

a. Aircraft Level FHA shall be developed. The aircraft level FHA shall identify potential functional failures and classifies the hazards associated with specific failure conditions for each aircraft function. The PMO is responsible for this activity.

1. For new functions added to existing aircraft, the aircraft FHA shall examine the

new high-level function, as well as, review existing aircraft functions to determine interactions between existing systems and the new function.

2. For modification to existing aircraft functions in which existing system-level safety analyses exists without an aircraft-level analysis, the aircraft FHA shall examine the modified high-level function, as well as, review of existing aircraft functions. The aircraft FHA shall determine interactions between existing systems and the modified function to determine if there are any aircraft-level hazards that require the modified system to have higher SHCI than the system-level FHA requires.

3. For modification to existing aircraft functions in which no existing safety analyses is available, the modified system shall be treated as a new function as discussed above.

4. For ground based mission planning systems, the top-level FHA shall identify how the generated data is used in the aviation platforms considering individual aircraft effects and multiple aircraft provided with the shared data. The FHA shall consider the effects of misleading, as well as, loss of the data.

5. For maintenance systems, the top-level FHA shall identify how the captured data is used in the aviation platforms in terms of the ability to detect, isolate, and eliminate faults.

b. System FHA shall be developed. The system-level FHA shall examine system functions to identify potential functional failures and classify the hazards associated with specific failure conditions. This may be included as either a separate system FHA or as part of the aircraft level FHA.

c. System safety requirements shall be traceable to the FHA. Traceability shall be

demonstrated by showing system safety requirements are established. Requirements shall include probability requirements, redundancy, assurance levels, design constraints and/or recommended flight crew or maintenance action. The traceability information and the system safety requirements are contained in the engine model specification.

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G.2.1.2.2 System Safety Assessment (SSA) – System Architecture (Sys Arch) Reqmt a. Aircraft Failure Conditions and Safety Requirements shall be developed and

allocated. An analysis of the aircraft architecture shall be accomplished to show the single system failures or combinations of systems failures that contribute to the hazards identified in the aircraft level FHA. This is a PMO activity.

b. System Failure Conditions and Safety Requirements shall be developed and

allocated. An analysis of the system architecture shall determine what single failures or combinations of failures can contribute to the hazards identified in the system FHA. The system allocation analysis may be included as either a separate system SSA or as part of the aircraft level SSA.

c. System Allocation Analysis shall be traceable to System FHA. System-Level

functions and safety requirements shall be allocated to lower levels. For catastrophic or critical hazards (level A/B), a system-level FTA or Dependency Diagrams may be used to demonstrate how the proposed architecture mitigates each hazard and determines safety requirements for lower level items. The allocation of safety requirements for marginal hazards can be accomplished by qualitative discussion or other modeling techniques, depending upon system complexity. Negligible hazards only require discussion of separation from other more critical functions or hazards. The system allocation analysis may be included as either a separate system SSA or as part of the aircraft level SSA.

d. Safety requirements shall be specified. Safety requirements derived from the

safety analyses shall be incorporated into the appropriate level of specification. This requirement shall be included in the engine model specification.

e. Safety Requirements shall be traceable to System FHA. All FHA failure conditions

shall be addressed in the SSA and have been allocated to system requirements. The trace information and the system safety requirements shall be included in engine model specification.

f. Software Hazard Criticality Index (SHCI) Assignment shall be determined and

justification provided. The system safety process shall determine the SHCI. Justification of all SHCI assignments shall be provided in the SSA. Refer to Section 4 for details of the SHCI levels. These requirements shall be included in the SSA.

G.2.2 Software Planning Process. The Contractor shall plan the software lifecycle activities and document them in the following software planning documents:

• Plan for Software Aspects of Certification (PSAC) • Software Development Plan (SDP) • Software Test Plan (STP) • Software Configuration Management Plan (SCMP) • Software Quality Assurance Plan (SQAP)

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The Contractor shall conduct a Software Planning Process Review (PPR).

G.2.2.1 Software Planning Process Artifacts.

G.2.2.1.1 Plan for Software Aspects of Certification (PSAC). The PSAC shall communicate the system overview, software overview, and airworthiness certification considerations. It shall identify the software life cycle, software life cycle data, schedule, and any other additional considerations that may affect the airworthiness approval process. The PSAC shall communicate the proposed development methods to the certification authority for agreement, and define the means of compliance with the lifecycle processes and requirements within this document. The PSAC shall be assessed for completeness and consistency of the proposed SHCI level(s), in comparison with the outputs of the software safety assessment process and other software life cycle data. The PSAC shall include compliance with the software aspects of the engine specification. Note: the PSAC contents may be included in the SDP instead of a separate PSAC document.

G.2.2.1.2 Software Development Plan (SDP). The SDP shall include the requirements, standards, and software life cycle(s) to be used in the software development processes. The SDP shall address the handling of any safety critical software to include the documentation, the flow down of all identified hazards mitigated by software to software safety requirements, safety requirements to the source code, and safety requirements to test cases. The SDP shall include the following Software Standards:

• Software Requirements Standards • Software Design Standards • Software Code Standards

G.2.2.1.3 Software Test Plan (STP). The STP shall describe plans for qualification testing of the software. It shall also describe the software test environment and descriptions and schedule to be used for the testing, identifies the test to be performed, and provide schedules for test activities. The STP shall address the method(s) used to perform software regression testing for the safety requirements identified in the SRS. The STP shall document the procedures for achieving structural coverage testing. An STP is required for each CSCI to be integrated into the platform, and each platform CSCI that was modified for the integration. The STP shall cover both normal operating requirements based testing as well as abnormal operating condition testing including fault injection testing.

G.2.2.1.4 Software Configuration Management Plan (SCMP). The SCMP shall identify the procedures, tools, methods, standards, and interfaces to be used in conjunction with configuration items. The SCM process and lifecycle objective evidence shall be recorded and made available to support contractual requirements and be available for review during all phases of the contractor’s performance. The SCMP shall include methods for identifying configurations, schedules for identifying configurations, baseline configuration establishment methods, baseline traceability, and methods for writing and closing associated problem reports. The SCMP shall identify methods for change control, methods for change review, methods for approval/disapproval and other status, and methods for recording, reporting’s, and retrieving configuration status. The SCMP shall identify the integrity controls, release methods, release authority, and data retention for configurations. The SCMP shall identify safeguards for software loading and safeguards for software loading tools. The SCMP

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shall include processes for control of the software development tools, software verification tools, and any equipment used to test the software; including hardware and lab facilities.

G.2.2.1.5 Software Quality Assurance Plan (SQAP). The SQAP shall identify the scope, organizational responsibilities, interfaces, standards, procedures, tools, and methods to be used in conjunction with the quality assurance process. The SQAP shall include a statement to identify the approval authority for all artifacts. The SQAP shall identify activities including reviews, audits, reporting, inspections, lifecycle process monitoring, problem reporting, problem tracking, and problem resolution actions. The SQAP shall include necessary transition criteria and timing for entering each software development process. The SQAP shall include a definition of the records produced from the software quality assurance process and a plan to ensure sub-tier suppliers’ processes follow this quality assurance process. The SQA process and lifecycle objective evidence shall be recorded and made available to support contractual requirements and be available for review during all phases of the developer’s performance. The SQA process and life cycle objective evidence shall be recorded and made available to the government to support contractual requirements. Provisions shall be made to permit Government representatives to review procedures and data during all phases of the developer’s performance. The developer shall provide software quality metrics in the quality assurance artifacts. The developer’s existing software quality data collection and reporting formats shall be made available to the Government for evaluation.

G.2.2.1.6 Report, Record of Meeting/Minutes (RRM). The RRM is an ongoing log that shall be maintained throughout the software lifecycle. The Contractor shall conduct software lifecycle process reviews with the Government during the program to establish a foundation for airworthiness substantiation and assure compliance with all airworthiness qualification requirements, and log the agendas and results of those reviews within the RRM. All meeting results (action items, acceptance, rejections, etc.) shall be included in the RRM.

G.2.2.1.7 Data Accession List (DAL) See Section 2.1.1.5.

G.2.2.2 Software Planning Process Requirements.

G.2.2.2.1 Plan for Software Aspects of Certification (PSAC) Requirements. a. Software development and integral process activities shall be defined and agreed

upon. The integral processes shall include verification, software quality assurance, and software configuration management. SHCI assignment level as defined by the SSA shall be specified, and associated requirements from Table 4 shall be identified. These requirements shall be included in the following:

• Plan for Software Aspects of Certification (PSAC) • Software Development Plan (SDP) • Software Test Plan (STP) • Software Configuration Management Plan (SCMP) • Software Quality Assurance Plan (SQAP)

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b. Transition criteria, inter-relationships, and sequencing among processes shall be defined. The entrance and exit criteria for transition between lifecycle phases shall be defined. The criteria for transitioning back to a previous phase shall also be defined. These requirements shall be captured in some combination of the following:

• Plan for Software Aspects of Certification (PSAC) • Software Development Plan (SDP) • Software Test Plan (STP) c. Software life cycle environment shall be defined. The methods and tools to be

used for the activities of each software life cycle process shall be defined. This shall be summarized in the PSAC. The development tools shall be detailed in the SDP. The test tools shall be detailed in the STP.

d. Required resources, schedule, facilities, organizations related to software

airworthiness qualification shall be determined and compliant with the engine specification. All defined plans within the PSAC shall comply with the stated engine specification. The PSAC requirements can be included in the SDP instead of a separate document.

G.2.2.2.2 Software Development Plan (SDP) Requirements. Standards shall be specified for all previously developed software, COTS software, and new software.

a. Software development standards shall be defined. The software development standards shall define the rules and constraints for requirements, design, and code. These requirements shall be delivered in the SDP.

G.2.2.2.3 Software Test Plan (STP) Requirements. a. Methods for verification independence shall be established. The plan for meeting

the independence requirements of this AQP shall be defined. See the definition of verification independence. These requirements shall be included in the STP.

b. Verification methods shall be defined. A description of review methods (checklists

and other aids), analysis methods (traceability and coverage analysis), test methods, and the test data to be produced shall be documented. This requirement shall be delivered in the STP.

c. Test case standards shall be defined. Test case standards shall be developed

that specify the test cases for each type of requirement. Topics that shall be addressed are: robustness testing, boundary condition testing, range testing, equivalence classes, tolerances, and how structural coverage will be achieved. This requirement shall be included in the STP.

d. Test environment(s) shall be defined. The test environment(s) shall be defined.

Include details of the tools to be used and account for any differences between the test environment and the final target hardware environment. This requirement shall be delivered in the STP.

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Refer to Section 6.3 for information on tool qualification.

G.2.2.2.4 Software Configuration Management Plan (SCMP) Requirements. The following requirements shall be specified in the SCMP:

a. Configuration items shall be identified. Documented software configuration management process shall ensure documents, source code, executable object code, and other lifecycle artifacts are configuration identified.

b. Baselines and baseline traceability shall be established. Documented software

configuration management process shall ensure software and document baselines are established. Change controls shall be followed between baselines. Traceability shall exist between a derived baseline and an established baseline.

c. Problem reporting, change control, change review, and configuration status

accounting shall be established. Documented software configuration management process shall detail the software problem reporting processes and procedures. The actual problem reports shall be included in the Report, Record of Meetings (RRM) and be used to report all of the following:

• process non-compliance with plans • output deficiency, formal review deficiency • software anomalous behavior The problem report summary shall be delivered periodically throughout the development

process. d. Archive, retrieval, and release shall be established. Processes shall be

established to archive configured items, retrieve configured items, and release configured items. e. Software life cycle environment control shall be established. The process for

controlling the development and test tools shall be established.

G.2.2.2.5 Software Quality Assurance Plan (SQAP) Requirements. The following requirements shall be specified in the SQAP unless otherwise stated.

a. Activities shall be defined to ensure software plans comply with this AQP and

established program requirements. SQA activities shall be defined to review the software plans to ensure the documented processes comply with this document and established program requirements. These shall be specified in the SQAP.

b. Activities shall be defined to ensure software plans are coordinated. SQA activities

shall be defined to review the software plans to ensure they are consistent with each other. These shall be specified in the SQAP.

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c. SQA Activities shall be defined. SQA activities shall be defined to ensure that the software product is compliant with requirements and the processes used to develop and verify the software are compliant with the documented plans and procedures. The method and timing of the SQA activities shall be explicitly specified. SQA shall assure the lifecycle requirements listed in Table 4 are completed correctly. These SQA activities shall be specified in the SQAP.

G.2.3 Software Requirements Process. The Contractor shall analyze the system requirements assigned to software and generate software high-level requirements that satisfy the allocated system requirements. This shall be delivered as Software Requirements Specification (SRS). The Contractor shall review the high-level software requirements and provide evidence that they are correct and complete. The Contractor shall generate a Software Failure Modes Effects Analysis (SwFMEA) to analyze the failure effects based on the software architecture. This shall be delivered as the SSA - SwFMEA. The Contractor shall analyze the high-level software requirements to ensure they meet the system safety requirements. This shall be included in final SSA. The Contractor shall conduct a Software Requirements Review (SWRR) and use MIL-STD-1521B Appendix C as a guide.

G.2.3.1 Software Requirements Process Artifacts.

G.2.3.1.1 Software Requirements Specification (SRS). The SRS shall include both the software requirements and interface requirements specification (IRS). The SRS shall reflect the software requirements associated with the re-used/legacy code not changed, re-used/legacy code modified, and newly developed code. The SRS shall distinguish all software safety requirements from the other software requirements IAW SAE-ARP-4754. The interface requirements specification portion of the document shall include all CSCI interfaces to be integrated into the platform and each platform CSCI that was modified for the integration. The SRS shall specify the data integrity requirements between CSCI’s. Each safety related requirement shall be individually flagged and potential failure conditions shall be documented. The SRS shall specify constraints for precision, accuracy, timing, and memory. The SRS shall include a traceability matrix that shows requirements traceability between the system–level requirements of the SSDD and high-level software requirements. The traceability matrix shall be bi-directional as described in 11.21 of RTCA/DO-178C.

G.2.3.1.2 Interface Requirements Specification (IRS). The IRS shall include all CSCI interfaces to be integrated into the platform and each platform CSCI that was modified for the integration. The IRS shall specify the requirements imposed on one or more system, subsystem, Hardware Configuration Items (HWCIs), Computer Software Configuration Items (CSCIs), manual operation, or other components to achieve one or more interfaces among these entities. The IRS shall be prepared for, and include in the SRS.

G.2.3.1.3 System Safety Assessment (SSA) – Software Architecture (SwArch). See Section 2.1.1.2.2.

G.2.3.1.4 Data Accession List (DAL). See Section 2.1.1.5.

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G.2.3.1.5 Report, Record of Meeting/Minute (RRM). See Section 2.2.1.6.

G.2.3.2 Software Requirement Process Requirements.

G.2.3.2.1 Software/Interface Requirements Specification (SRS/IRS) Requirements. a. High-level requirements shall be developed. The system requirements allocated

to software shall be analyzed. Software high-level requirements shall be developed and documented. This shall be included as the SRS.

b. Derived high-level requirements shall be defined. Derived high-level software

requirements shall be developed and documented. This shall be included in the SRS. c. Software high-level requirements shall comply with system requirements.

Compliance shall be demonstrated by showing that the system functions to be performed by the software are defined, that the functional, performance, and safety requirements of the system are satisfied by the high-level software requirements, and that derived high-level software requirements and the reason for their existence are correctly defined. The evidence shall be documented in the DAL.

d. High-level requirements shall be accurate and consistent. Accuracy and

consistency shall be demonstrated by showing that each high-level software requirement is accurate, unambiguous, and sufficiently detailed and that the requirements do not conflict with each other. The evidence shall be documented in the DAL.

e. High-level requirements shall be compatible with target computer. Compatibility

shall be demonstrated by showing that no conflicts exist between the high-level software requirements and the hardware/software features of the target computer. The evidence shall be documented in the DAL.

f. High-level requirements shall be verifiable. Verifiability shall be demonstrated by

showing that all high-level software requirements can be proven correct through review, analysis, and/or testing. The evidence shall be documented in the DAL.

g. High-level requirements shall conform to standards. Conformance shall be

demonstrated by showing that the Software Requirements Standards are followed during the software requirements process and that deviations from the standard are justified. The evidence shall be documented in the DAL.

h. High-level requirements shall be traceable to system requirements. Traceability

shall be demonstrated by showing that functional and performance requirements of the system that are allocated to the software are contained in the software high-level requirements. The trace information shall be documented in the SRS. The trace data shall be bi-directional (i.e. system

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requirements to high-level software requirements and high-level software requirements to system requirements).

i. Algorithms shall be accurate. Algorithm accuracy shall be demonstrated by

showing that the accuracy, precision, and behavior of the proposed algorithms in the high-level software requirements are correct. The evidence shall be documented in the DAL.

j. High-level safety-critical requirements shall be identified. All high-level software

requirements identified in the software safety analyses as contributors or mitigation to hazards shall be labeled as such. This shall be documented in SRS.

k. High-level safety-critical requirements shall be traceable to FHA. Traceability shall

be demonstrated by showing that safety-critical high-level software requirements trace to a system function (allocated to software as specified in the SSDD, which has associated hazards. The trace information shall be documented in the SRS.

l. Safety-critical constraints shall be developed. Safety-critical constraints shall

include timing, throughput, and sizing constraints for safety-critical software elements. These constraints shall be developed into high-level software requirements. The resulting high-level software requirements shall be documented in the SRS.

G.2.3.2.2 System Safety Assessment (SSA) – Software Architecture (SwArch) Reqmt a. Software Safety Architectural Analysis (Dependency Diagram or FTA & Software

Interface FMEA) shall be performed. Analysis shall be performed that evaluates the software architecture against requirements allocated to software from the system-level. The software architecture analyses shall identify system functions allocated to software, software failure modes of those functions (more detailed than the system level allocation), and techniques to mitigate/prevent the failures. Special attention shall be provided for any segregation/partitioning techniques planned. The requirements of this analysis is to understand the elements of the system containing software, common software used in multiple subsystems and equipment, and what safety requirements for interfaces that shall be established. This shall be included in the SSA.

b. Software Safety Architectural Analysis shall be traceable to System FHA.

Traceability shall be demonstrated by showing that all software functions are included in architectural analysis and all software analyzed for interface functional failures, all failure modes have been addressed (for example, loss of interface and malfunctions such as, interface is out-of-tolerance high, out-of-tolerance low, etc.), and all software interface requirements are identified. Any differences shall be resolved. This shall be included in the SSA.

c. Safety requirements shall be validated. The safety requirements derived from

safety analyses to each hazard shall be validated. Some examples are: assumptions used in the safety analysis are captured as requirements (captured as requirements to ensure review for

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validity of the assumption and/or safety analysis modified to correct condition), validating independence between software or CSCI partitions, incorporation of software design features (BITS, parameter state change monitoring, etc.), and analyzing common mode failures (power, operating environment, memory, processor, etc.). This shall be included in the SSA.

G.2.4 Software Design Process. If required based on the Software Hazard Criticality Index (SHCI), the Contractor shall decompose the software high-level requirements into software architecture and low-level software requirements that satisfy the high-level software requirements. This shall be included in the SDD. The Contractor shall review the software architecture and low-level software requirements and provide evidence that they are correct and complete. The Contractor shall generate software test cases that verify the high-level software requirements and the low-level software requirements. The test cases shall include both normal range and robustness test cases for each high-level and low-level software requirement. This shall be included in the STD. The Contractor shall generate a Software Interface FMEA, which relates to data and control coupling in software interfaces between software items (between CSCIs) and software/hardware interfaces. This FMEA shall include all software/software (CSCI to CSCI) and software/hardware interfaces as defined in the interface requirements specifications to analyze the failure modes and assess the effects of the software and the hardware failure effects to software interfaces. SAE ARP5580 provides guidance on FMEAs. This shall be included in the SSA. The Contractor shall analyze the system and software test cases used to verify the software safety requirements to ensure they are complete and correct. This shall be included in the SSA. The Contractor shall generate a preliminary assessment of the hazard mitigation, control of failure effects, and residual risk. This shall be included in the SAR. The Contractor shall conduct a Software Preliminary Design Review (PDR) using MIL-STD-1521B Appendix D as a guide. The Developer shall conduct a Software Critical Design Review (CDR) using MIL-STD-1521B Appendix E as a guide.

G.2.4.1 Software Design Process Artifacts.

G.2.4.1.1 Software Design Description (SDD). The SDD shall include both the software design and interface design description (IDD). The SDD shall describe low-level requirements and software architecture implemented to satisfy the high-level requirements. The SDD shall include a data and control flow of the design. The SDD shall include timing and memory resource limitations, and all details necessary to describe the partitioned behavior, including scheduling procedures, inter-partition communication methods, partitioning methods, and partition protection methods. The SDD shall include an analysis to ensure no deactivated code can be executed on the target environment. The SDD shall specify the interface design details for all SRS interface requirements; which shall cover interfaces between all CSCI’s. An SDD is required for each CSCI to be integrated into the platform, and each platform CSCI that was modified for the integration.

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The SDD shall include a traceability matrix that shows requirements traceability between the high–level requirements of the SRS and low-level software requirements. The traceability matrix shall be bi-directional as described in 11.21 of RTCA/DO-178C.

G.2.4.1.2 Interface Design Description (IDD). The IDD shall describe the interface characteristics of all systems, subsystems, Hardware Configuration Items (HWCIs), Computer Software Configuration Items (CSCIs), manual operations, or other system components. An IDD shall be prepared for, and included in each SDD.

G.2.4.1.3 Software Test Description (STD). A complete STD shall be required to cover all CSCIs to be integrated into the target platform, and every platform CSCI that was modified for the integration. The STD shall include all plans necessary to complete a successful FQT, which includes the analysis, test procedures, test cases and requirements traceability, and regression test suite. Tests shall adequately demonstrate the functionality of the software requirements (high level requirements) and design requirements (low level requirements). The STD shall describe the FQT test environment, which shall be completed on the system representative/target/flight configuration. The test procedures shall be detailed down to step-by-step instructions for how each test case is to be setup and executed, how the test results are evaluated and the test environment to be used. The STD shall include a traceability matrix that shows requirements traceability between all software requirements (high and low) and the test cases. The STD shall also include a traceability matrix that shows requirements traceability between all test cases and test procedures (test scripts). The traceability matrices shall be bi-directional as described in 11.21 of RTCA/DO-178C. If test scripts are used and referenced in the STD, they shall be made available as part of the STD review.

G.2.4.1.4 System Safety Assessment (SSA) – SW FMEA. See Section 2.1.1.2.3

G.2.4.1.5 Safety Assessment Report (SAR). The SAR and SSA (final) can be integrated into a single document or the contents coordinated between the documents to avoid duplication. When this is implemented, a mapping to the requirements shall be provided. If the SSA contains a list of failure conditions traceable to the FHAs, then only the residual risk assessment has to be provided in the SAR.

G.2.4.1.6 Report, Record of Meeting/Minutes (RRM). See Section 2.2.1.6.

G.2.4.1.7 Data Accession List (DAL). See Section 2.2.1.7.

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G.2.4.2 Software Design Process Requirement.

G.2.4.2.1 Software/Interface Design Description (SDD/IDD) Requirements. a. Low-level requirements shall be developed. The high-level software requirements

shall be analyzed. Software low-level requirements shall be developed and documented. This shall be included in the SDD, if required for the SHCI level assignment.

b. Derived low-level requirements shall be defined. Derived low-level software

requirements shall be developed and documented. This shall be included in the SDD, if required for the SHCI level assignment.

Software architecture shall be developed. The software architecture shall be developed

and documented based on the high-level software requirements, low-level software requirements, and target hardware. This shall be included in the SDD, if required for the SHCI level assignment.

c. Low-level requirements shall comply with high-level requirements. Compliance

shall be demonstrated by showing that the software low-level requirements satisfy the software high-level requirements (e.g. bus loading, system response times, and input/output hardware). Derived low-level software requirements and the reason for their existence shall also be correctly defined. The evidence shall be documented in the DAL.

d. Low-level requirements shall be accurate and consistent. Accuracy and

consistency shall be demonstrated by showing that the low-level software requirements are unambiguous and do not conflict with each other. The evidence shall be documented in the DAL.

e. Low-level requirements shall be compatible with target computer. Compatibility

shall be demonstrated by showing that no conflicts between the low-level software requirements and the hardware/software features of the target computer exist. The evidence shall be documented in the DAL.

f. Low-level requirements shall be verifiable. Verifiability shall be demonstrated by

showing that all low-level software requirements can be proven correct through review, analysis, and/or testing. The evidence shall be documented in the DAL.

g. Low-level requirements shall conform to standards. Conformance shall be

demonstrated by showing that the low-level software requirements do not conflict with the Software Design Standards. (E.g. complexity restrictions, restricted design constructs). Any deviations shall be justified. The evidence shall be documented in the DAL.

h. Low-level requirements shall be traceable to high-level requirements. Traceability

shall be demonstrated by showing the high-level software requirements and derived high-level software requirements were developed into low-level software requirements. The trace information shall be included in the SDD. The trace data shall be bi-directional (i.e. high-level software requirements to low-level software requirements and low-level software requirements to high-level software requirements).

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i. Algorithms shall be accurate. Algorithm accuracy shall be demonstrated by

showing the accuracy and behavior of the low-level algorithms is correct. The evidence shall be documented in the DAL.

j. Software architecture shall be compatible with high-level requirements.

Compatibility shall be demonstrated by showing that the software architecture does not conflict with the high-level software requirements. The evidence shall be documented in the DAL.

k. Software architecture shall be consistent. Consistency shall be demonstrated by

showing that the data flow and control flow between the components of the software architecture is correct. The evidence shall be documented in the DAL.

l. Software architecture shall be compatible with target computer. Compatibility shall

be demonstrated by showing that no conflicts exist between the software architecture and the hardware/software features of the target computer (e.g. initialization, asynchronous operation, synchronization, interrupts). The evidence shall be documented in the DAL.

m. Software architecture shall be verifiable. Verifiability shall be demonstrated by

showing that the software architecture can be proven correct through review, analysis, and/or testing (e.g., no unbounded recursive algorithms, no unreachable states). The evidence shall be documented in the DAL.

n. Software architecture shall conform to standards. Conformance shall be

demonstrated by showing that the software architecture does not conflict with the Software Design Standards. (e.g., complexity restrictions, restricted design constructs, etc...). The evidence shall be documented in the DAL.

o. Software partitioning integrity shall be confirmed. Software partitioning integrity

shall be demonstrated by showing that partitioning breaches are prevented or isolated. If different SHCI are used in the various partitions, the SHCIs shall be justified in cooperation with the system safety team. The evidence shall be documented in the DAL.

p. Low-level safety-critical requirements shall be developed. Low-level software

requirements are safety-critical if they are derived from high-level safety-critical requirements or from detailed software safety analyses. The resulting low-level requirements shall be documented in the SDD.

q. Low-level safety-critical requirements shall be traceable to SSA. The safety-critical

requirements shall trace to the SSA through the high-level software requirements. During the design phase the low-level requirements shall be traced to the high-level software requirements. This process shall result in low-level safety-critical requirements that trace to the SSA. The trace information shall be documented in the SDD.

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G.2.4.2.2 Software Test Description (STD) Requirement. a. Test cases shall be developed. Test cases shall be developed to verify all of the

software high-level requirements and software low-level requirements. This shall be included in the STD.

b. Test cases shall be traceable to requirements. Traceability shall be demonstrated

by showing that test cases fully cover all high-level and low-level software requirements and that every test case traces to one or more requirement. The trace information shall be included in STD. The trace data shall be bi-directional (i.e. high-level and low-level requirements to test cases and test cases to high-level and low-level software requirements).

c. Test cases shall comply with requirements. Compliance shall be demonstrated by

showing that test cases accurately and fully test the requirements they trace to. The evidence shall be included in the DAL.

d. Test cases shall be complete. Completeness shall be demonstrated by showing

that all test cases include a set of inputs, conditions, expected results, and pass/fail criteria. The evidence shall be documented in the DAL.

e. Test procedures shall be traceable to test cases. Traceability shall be

demonstrated by showing bi-directional association between test cases and test procedures. The trace information shall be included in STD.

G.2.4.2.3 System Safety Assessment (SSA) – SW FMEA Requirements. a. Software Interface FMEA is developed. The Software Interface FMEA shall

identify failures, which relate to data and control coupling in software interfaces between software items (between CSCIs) and software/hardware interfaces. This FMEA shall include all software/software (CSCI to CSCI) and software/hardware interfaces as defined in the interface requirements specifications.

Note: Best practices to ensure the software interfaces are consistent and safe are: 1)

Have strong data typing standards for all software interfaces, 2) Use message passing across interfaces (i.e. don’t use common data structures or global data), 3) Have rigorous design and coding standards, and 4) Use automated tools to enforce the standards.

b. Safety Test Cases and Software Test procedures are consistent with Safety

Analyses. Software and system test cases are used to verify the software safety requirements. Software Test Procedures and safety test cases shall be evaluated to ensure they are consistent with and compliant with the failure modes, fault conditions, mitigations, and derived safety requirements of the system safety process. Tests shall trace to safety requirements derived from the safety analyses.

Note: The SSA and SAR can be separate or combined submittals. When combined, a

mapping to each data requirements shall be provided.

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G.2.4.2.4 Safety Assessment Report (SAR) Requirements. a. Software-related hazards shall be controlled and mitigated. Preliminary

assessment, that the software safety requirements that trace to hazards shall eliminate and/or control these hazards to sufficient levels. This analysis shall be included in the SAR.

b. Control of failure effects shall be satisfactory. Preliminary assessment, that the

failure effects’ probability of occurrence and severity level shall be reduced to satisfactory levels. This analysis shall be included in the SAR.

c. Residual risk shall be acceptable. Preliminary assessment, that the residual risk

that remains from controlled and uncontrolled hazards shall be satisfactory. This analysis shall be included in the SAR.

G.2.5 Software Coding Process. The Contractor shall generate source code that implements the software high-level requirements, the software architecture, and the software low-level requirements (if applicable for the SHCI). The Contractor shall review the source code and provide evidence that it is correct and complete. The Contractor shall revise the software test cases based on any changes during the software coding process. The test cases shall include both normal range and robustness test cases for each high-level and low-level software requirement. This shall be included in the STD. Test scripts, if any are developed, shall be included as part of the STD. The Contractor shall analyze the test cases to ensure the structural coverage required for the SHCI is achieved. The Contractor shall generate a Software Fault Tree Analysis, if identified in the SwSPP or safety analysis process that examines the source code to determine all possible ways in which the source code can cause specific hazards identified as requiring the detailed analysis. The analysis shall identify and mitigate any areas of the source code that can directly cause the system hazards. Since the any detailed FTAs will be developed during the coding process, the FTA shall be part of peer review process and the materials available during Government audit of peer reviews. The final detailed FTAs, when specified, shall be included in the SSA.

G.2.5.1 Software Coding Process Artifacts

G.2.5.1.1 System Safety Assessment – Software Fault Tree Analysis (SW FTA). See Section 2.1.1.2.4.

G.2.5.1.2 Report, Record of Meeting/Minutes (RRM). See Section 2.2.1.6.

G.2.5.1.3 Data Accession List (DAL). See Section 2.2.1.7.

G.2.5.1.4 Software Test Description (STD). See Section 2.4.1.3.

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G.2.5.2 Software Coding Process Requirements.

G.2.5.2.1 Software Code Requirements. a. Source code shall be developed. The source code shall be generated using the

low-level requirements and the software architecture. b. Source code shall comply with low-level requirements. Compliance shall be

demonstrated by showing that source code implements the low-level software requirements and does not implement undocumented functions. The evidence shall be documented in the DAL.

c. Source code shall comply with software architecture. Compliance shall be

demonstrated by showing that source code matches the data flow and control flow defined in the software architecture. The evidence shall be documented in the DAL.

d. Source code shall be verifiable. Verifiability shall be demonstrated by showing that

source code does not contain statements/structures that cannot be verified and that the code does not have to be altered to test it. The evidence shall be documented in the DAL.

e. Source code shall conform to standards. Conformance shall be demonstrated by

showing that source code meets the Software Coding Standards specified (e.g. complexity restrictions, restricted code constructs). The evidence shall be documented in the DAL.

f. Source code shall be traceable to low-level requirements. Traceability shall be

demonstrated by showing that source code includes tags that trace back to low-level software requirements. The evidence shall be documented in shall be DAL. The trace data must be bi-directional (i.e. low-level software requirements to source code and source code to low-level software requirements).

g. Source code shall be accurate and consistent. Accuracy and consistency shall be

demonstrated by showing that source code is correct and consistent for potential issues such as stack usage, fixed point arithmetic overflow and resolution, worst-case execution timing, exception handling, use of uninitialized variables or constants, unused variables or constants, and data corruption due to task or interrupt conflicts. The evidence shall be documented in the DAL.

h. Safety-critical algorithms shall be correct. Algorithm accuracy shall be

demonstrated by showing correct results for analyses/reviews of algorithm accuracy, behavior, and boundary value conditions. The evidence shall be documented in the DAL.

i. Safety-critical source code shall be traceable to hazard analyses. Traceability

shall be demonstrated by showing that source code includes tags that trace back to safety-critical high-level and low-level requirements, which trace back to system safety requirements, which trace back to safety analyses. The evidence shall be documented in the DAL.

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G.2.5.2.2 Software Test Description (STD) Requirements. The STD shall be updated to include verification of the low-level software requirements. The following requirements shall be added:

a. Test procedures shall be correct. The test cases shall be reviewed to ensure they fully cover all aspects of each high-level and low-level software requirement, that each step is properly detailed, and that the inputs and expected outputs are determined from the requirements. The revised test procedures shall be included in the STD.

b. Test coverage of high-level requirements shall be achieved. Coverage shall be

demonstrated by showing test cases and procedures exist for each high-level software requirement and satisfy the criteria for normal and robustness testing of each high-level software requirement. The evidence shall be documented in shall be DAL. The trace data shall be bi-directional (i.e. high-level software requirements to test cases and test cases to high-level software requirements).

c. Test coverage of low-level requirements shall be achieved. Coverage shall be

demonstrated by showing test cases and procedures exist for each low-level software requirement and satisfy the criteria for normal and robustness testing of each low-level software requirement. The evidence shall be documented in the DAL. The trace data must be bi-directional (i.e. low-level software requirements to test cases and test cases to low-level software requirements).

d. Test coverage of software structure (modified condition/decision) is achieved. If

required based on the SHCI assignment level (refer to Table 4), coverage shall be demonstrated by showing the test cases and procedures provide modified condition/decision coverage of all source code. The evidence shall be documented in DAL.

e. Test coverage of software structure (decision coverage) shall be achieved. If

required based on the SHCI assignment level (refer to Table 4), coverage shall be demonstrated by showing the test cases and procedures provide decision coverage of all source code. The evidence shall be documented in the DAL.

f. Test coverage of software structure (statement coverage) shall be achieved. If

required based on the SHCI assignment level (refer to Table 4), coverage shall be demonstrated by showing the test cases and procedures provide statement coverage of all source code. The evidence shall be documented in the DAL.

g. Test coverage of software structure (data coupling and control coupling) shall be

achieved. If required based on the SHCI assignment level (refer to Table 4), coverage shall be demonstrated by showing the test cases and procedures provide coverage of all data and control coupling in the software architecture and low-level requirements. The evidence shall be documented in the DAL.

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h. Regression test suite shall be developed. A regression test suite shall be specified that includes a minimum set of test cases and procedures for regression testing, which shall cover all safety critical software functionality. The regression test suite shall be specified in the STD.

G.2.5.2.3 System Safety Assessment (SSA) – Fault Tree Analysis (SwFTA) Reqmts. a. Software Detailed Causal Analysis (SwFTA or detailed SwFMEA) shall be

developed. The Software Detailed FTA shall examine the source code to determine all possible ways in which the source code can cause specific hazards identified as requiring the detailed analysis. The analysis shall identify and mitigate any areas of the source code that can directly cause the system hazards. Since the detailed FTAs will be developed during the coding process, the FTA shall be part of peer review process and the materials available during Government audit of peer reviews. The final detailed FTAs, when specified, shall be included in the SSA.

b. Software Detailed FTA shall be compliant with the FHA. Compliance shall be

demonstrated by showing how the hazards identified faults in the FHA have been analyzed in the SwFTA. This shall be included in the SSA.

G.2.6 Software Integration Process. The software integration process shall be the activities that translate the source code into executable object code. This includes the definition of memory maps, linker commands, build scripts, module lists and the compile and link process. The outputs of the Software Integration Process shall be the executable object code, linker map, memory map, and build and load procedures. This process shall also include the steps required to run the executable object code on the target computer. This process shall also include the integration of the individual source modules into a single executable. The Contractor shall integrate the source code, generate executable code, execute the code on the target system, and verify that the integration is complete and correct. The Contractor shall provide evidence that the integration is correct and complete. The Contractor shall generate software build and load instructions and capture the configuration of the executable software, source code, software build tools, and all the lifecycle data for developing and verifying the software. This shall be included in the Software Version Description (SVD).

G.2.6.1 Software Integration Process Artifacts

G.2.6.1.1 Software Version Description (SVD). The SVD shall be developed and delivered for any version of the software released for formal testing/production/fielding. The SVD shall identify the software lifecycle environment hardware and operating system software, the software development tools, the data integrity, the software test environment, and any qualified tools and their associated qualification data. The SVD shall capture the configuration of the executable software, source code, software build tools, software build and load instructions, and all the lifecycle data for developing and verifying the control system embedded software for each CSCI. The SVD shall contain a list of all changes incorporated into the software version since the previous version. The SVD shall identify the problem reports, change proposal and change notices associated with each change and the effects of each change on the system operation and on interfaces with other hardware and software.

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G.2.6.1.2 Data Accession List (DAL). See Section 2.1.1.5.

G.2.6.1.3 Report, Record of Meeting/Minute (RRM). See Section 2.2.1.6.

G.2.6.2 Software Integration Process Requirements.

G.2.6.2.1 Executable Code Requirements. a. Executable object code shall be produced and integrated in the target computer.

Scripts, commands, memory maps, and tools shall be developed to integrate the source modules into executable object code that runs on the target computer. These results shall be documented in the SVD.

b. Output of software integration process shall be complete and correct.

Completeness and correctness shall be demonstrated by an analysis of the link data and memory map to detect incorrect hardware addresses, memory overlaps, and missing software components. The evidence shall be documented in the DAL.

G.2.6.2.2 Software Version Description (SVD) Requirements. a. Software version description shall be complete. The Software Version Description

document shall capture the configuration of the executable software, source code, software build tools, software build and load instructions, and all the lifecycle data for developing and verifying the software. This shall be included in the SVD.

b. Software load control shall be established. Procedures shall be developed and

documented to ensure that only properly configured software versions are loaded on the target and used for testing. The procedures shall include methods such as a 32-bit CRC to minimize the likelihood of an undetected bad software load.

G.2.7 Software Test Verification Process.

G.2.7.1 Software Test Verification Process Artifacts.

G.2.7.1.1 Software Test Report (STR). The STR shall include the results of each test, the procedures and test cases associated with the results, and the personnel that witness the test. A STR shall be required for all CSCIs to be integrated in to the platform. The STR shall include Requirements Verification Matrix. The Requirement Verification Matrix shall contain demonstrate verification of each software requirement indication the test case identification number and summary of the test results that verify each software requirement. The STR shall include the configuration item of software version reviewed. The STR shall include all structural coverage, for Modified Condition Decision Coverage, results and traceability analyses with test result.

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Structural coverage testing shall be performed and included with the STR for any source code assigned an SHCI level of A through C. The Contractor may choose to include the results of high-level test cases and low-level test cases in separate documents.

G.2.7.1.2 Data Accession List (DAL). See Section 2.1.1.5.

G.2.7.1.3 Report, Record of Meeting/Minute (RRM). See Section 2.2.1.6.

G.2.7.2 Software Test Verification Process Requirements.

G.2.7.2.1 Software Test Report (STR) Requirements. a. Test results shall be correct and discrepancies explained. Correctness shall be

demonstrated by showing that expected and actual test results match and any discrepancies are explained. Test discrepancies shall be documented as problem reports. The test results shall be included in the STR.

b. High-level requirements test results shall be correct and complete. All high-level

software requirements based test cases (both normal range and robustness) shall be demonstrated as complete and passed. This shall be contained in the STR. Bi-directional trace data shall be part of the evidence to demonstrate this (i.e. high-level test cases to test results and test results to high-level test cases).

c. Low-level requirements test results shall be correct and complete. All low-level

software requirements based test cases (both normal range and robustness) shall be demonstrated as complete and passed. This shall be contained in the STR. Bi-directional trace data shall be part of the evidence to demonstrate this (i.e. low-level test cases to test results and test results to low-level test cases).

d. Coverage results of software structure shall be complete. Structural coverage

results for required coverage for the SHCI assignment level (modified condition/decision, decision, statement, data coupling, control coupling) shall be complete and acceptable. This shall be contained in the STR.

e. Regression test results shall be correct and complete. The results of all tests (that

verify the test software changes were correctly implemented and do not cause any inadvertent changes) shall be completed and passed. This shall be contained in STR.

f. Executable object code shall comply with high-level requirements. Compliance

shall be demonstrated by executing the high-level software requirements based test cases on the target computer. The results shall be contained in the STR.

g. Executable object code shall be robust with high-level requirements. Compliance

shall be demonstrated by executing the high-level software requirements based robustness test cases on the target computer. The results shall be contained in the STR.

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h. Executable object code shall comply with low-level requirements. Compliance

shall be demonstrated executing the low-level software requirements based test cases on the target computer. This shall be contained in the STR.

i. Executable object code shall be robust with low-level requirements. Compliance

shall be demonstrated by executing the low-level software requirements based robustness test cases on the target computer. The results shall be contained in the STR.

j. Executable object code shall be compatible with target computer. Executing the

high-level and low-level software requirements-based test cases that verify the software works on the target computer shall demonstrate software compatibility. The results shall be contained in the STR.

G.2.8 Software Approval Process

G.2.8.1 Software Approval Process Artifacts.

G.2.8.1.1 Software Accomplishment Summary (SAS). The SAS shall show the compliance of the software to the airworthiness requirements. The summary shall show compliance to the proposed system overview and software overview. It shall restate the certification considerations from the PSAC and describe any differences. It shall show compliance to the stated PSAC software life cycle and software lifecycle data. It shall identify the software configuration, include a software change history, and summarize any unresolved problem reports. The SAS shall address all software aspects of the engine specificationto demonstrate how the software data and processes show compliance to the engine specification.

G.2.8.1.2 Software Product Specification (SPS). The SPS shall document the source code; build scripts, linker command, compiler options and all other data necessary to build the software from the source code. The information shall provide all the information needed to maintain the software after initial delivery.

G.2.8.1.3 Safety Assessment Report (SAR). See Section 2.4.1.5

G.2.8.1.4 System Safety Assessment (SSA) – Final. See Section 2.1.1.2.5.

G.2.8.1.5 Data Accession List (DAL). See Section 2.1.1.5.

G.2.8.1.6 Report, Record of Meeting/Minute (RRM). See Section 2.2.1.6.

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G.2.8.2 Software Approval Process Requirements.

G.2.8.2.1 Software Accomplishment Summary (SAS) Requirements. a. Software development processes shall comply with approved software plans and

standards. Compliance shall be demonstrated by showing software quality assurance records contain completed checklists for process compliance reviews in each process. This shall be included in the SAS.

b. Transition criteria for all processes shall be complete. Completeness shall be

demonstrated by showing software quality assurance records contain completed checklists for transitioning between processes in the lifecycle per the software plans. This shall be summarized in the SAS.

c. Lifecycle processes and data shall be complete. Completeness shall be

demonstrated by showing software quality assurance records contain completed checklists for all developed artifacts in each process. This shall be summarized in SAS. The final source code and build instructions shall be included in the SPS.

G.2.8.2.2 Safety Assessment Report (SAR) Requirements. a. Airworthiness Qualification Substantiating Report (AQSR) shall summarize the

compliance with the engine specification. It shall include or references all the test results, documents and activities that demonstrate compliance with each engine specification requirement. The AQSR shall be completed by the Airworthiness Authority. The Contractor shall provide all artifacts requested by Airworthiness Authority to support generation of the AQSR. This shall be included in the SAR.

b. Software-related hazards shall be controlled and mitigated. The software safety

requirements that trace to hazards shall be eliminated and/or controlled to sufficient levels. Evidence of this shall be included in the SAR.

c. Control of failure effects shall be satisfactory. Failure conditions/hazards shall be

reduced to satisfactory risk levels. Evidence of this shall be included in the SAR. d. Residual risk shall be acceptable. The SAS shall document completion of the

software lifecycle requirements by the PSAC. Open problem reports in the SAS shall be reviewed to determine impacts to safety analysis. Software that does not accomplish all requirements shall be identified as shortfalls in the SAR. The assessment of residual risk shall be included in the SAR.

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G.3. Software Hazard Criticality Index (SHCI). The Software Hazard Criticality Index provides a method for determining the criticality level of the CSCI and associated system safety activities that shall be used during system and software development. Note: The terminology SHCI is used to prevent any conflicts with the SAE ARP4754A and RTCA DO-178 term Development Assurance Level.

G.3.1 Software Hazard Criticality Index (SHCI) Definitions. The Software Hazard Criticality Indexes shall be defined based on the type of system in which the software will be implemented (Table 1). Matrices (Table 2 and 3) provide guidelines for determining the SHCI categorization based on the effects created by hazards.

Table 1: SYSTEM CATEGORIES AND DEFINITION System Type Category System Type Definition Flight Aid System (A) Table 2

Any system or equipment that does not directly control the aircraft while in flight but provides data/information to aircraft systems or the crew that may be used to control the aircraft, interface with aircraft systems, or manage aircraft systems is classified under this category. Examples are Engine Monitoring Systems (EMS), Ground-Based Pre-Flight Mission Planning Systems, Electronic Flight Bag Systems , and Ground-Based Post-Flight Condition Maintenance Systems.

Manned Aircraft (M) Table 3

Manned aircraft are defined as aircraft with human crew onboard. This includes aircraft that could be used to transport causalities that are remotely piloted.

G.3.2 SHCI Assignment Method. The SHCI shall be determined based on the failure condition (hazard) severity classifications starting at the aircraft-level and allocating down to system-level, equipment-level, and ultimately the software-level. Each failure condition defined by the Aircraft FHA shall be assigned a SHCI level based on the failure effects defined in Table 2 and 3. The SHCI level shall be allocated to the systems using the criteria defined in ARP4754 Sections 4.2 and 5.2. The system-level FHA failure conditions shall be assigned in a similar manner considering that the system SHCI requirement may be higher due to the aircraft SHCI allocation than the effects categorization defined by Table 2 and 3.

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G.3.3 Guidelines for Determining SHCI Assignment Levels.

Table 2: FLIGHT AID SYSTEM SHCI TABLE Risk Type Level A Level B Level C Level D Personnel Risk

Failure could indirectly result in a loss of life or permanent total disability due to improper configuration of aircraft or failure to configure aircraft.

Failure could indirectly result in permanent partial disability, injuries or occupational illness that may result in hospitalization of at least three personnel injury due to improper configuration of aircraft or failure to configure aircraft.

Failure could indirectly cause injury or occupational illness resulting in one or more lost work days(s) or physical distress to passengers

Failure could directly cause minor physical discomfort to passengers or injury or illness not resulting in a lost workday.

Aircraft Control Risk

Failure could indirectly result in immediate or near-immediate loss of aircraft control; failure to detect and replace faulty safety-critical parts, or misleading pilot information to establish safe route in integrated airspace. Failure could prevent continued safe flight and landing.

Failure could indirectly result in immediate emergency landing or cause pilot to take emergency action. Indirectly result in failure to provide pilot information to establish safe route in integrated airspace. It is remotely possible that the failure could result in loss of life. Indirectly causes physical distress or excessive workload such that the flight crew cannot be relied upon to perform their tasks accurately or completely

Failure could indirectly result in reduced control safety through lack of redundant systems; failure to detect and replace faulty aircraft parts; failure to provide proper pre-flight safety check. Failure would result in discomfort and/or physical distress to the crew. It is improbable that the failure could result in loss of life.

Failure could indirectly result in a slight reduction in safety margins or functional capabilities; Slight increase in crew workload, such as routine flight plan changes, or some physical discomfort to passengers.

Environmental Risk

Failure could directly cause irreversible severe environmental damage that violates law or regulation.

Failure could directly cause irreversible severe environmental damage that violates law or regulation

Failure could directly cause mitigable environmental damage without violation of law or regulation

Failure could cause minimal environmental damage not violating law or regulation

NOTES: 1. If the top-down SHCI assignment approach is used, the Flight Aid system would be included in the SHCI

allocation from the aircraft and system-level analyses. 2. If the aircraft integration is unknown or if the flight aid is used across aircraft types, different operations, and

different organizations, an FHA of the Flight Aid system can be used to identify Flight Aid functions within the various operating scenarios. Failure conditions can be assessed for each operating condition with the worst-case resulting severity establishing the SHCI requirements.

Table 3: MANNED AIRCRAFT SHCI TABLE Risk Type Level A Level B Level C Level D Personnel Risk

Failure could indirectly result in a loss of life or permanent total disability due to improper configuration of aircraft or failure to configure aircraft.

Failure could indirectly result in permanent partial disability, injuries or occupational illness that may result in hospitalization of at least three personnel injury due to improper configuration of aircraft or failure to configure aircraft.

Failure could indirectly cause injury or occupational illness resulting in one or more lost work days(s) or physical distress to passengers

Failure could directly cause minor physical discomfort to passengers or injury or illness not resulting in a lost workday.

Aircraft Control Risk

Failure could indirectly result in immediate or near-immediate loss of aircraft control; failure to detect and replace faulty safety-critical parts, or misleading pilot information to establish safe route in integrated airspace. Failure could prevent continued safe flight and landing.

Failure could indirectly result in immediate emergency landing or cause pilot to take emergency action. Indirectly result in failure to provide pilot information to establish safe route in integrated airspace. It is remotely possible that the failure could result in loss of life. Indirectly causes physical distress or excessive workload such that the flight crew cannot be relied upon to perform their tasks accurately or completely

Failure could indirectly result in reduced control safety through lack of redundant systems; failure to detect and replace faulty aircraft parts; failure to provide proper pre-flight safety check. Failure would result in discomfort and/or physical distress to the crew. It is improbable that the failure could result in loss of life.

Failure could indirectly result in a slight reduction in safety margins or functional capabilities; Slight increase in crew workload, such as routine flight plan changes, or some physical discomfort to passengers.

Environmental Risk

Failure could directly cause irreversible severe environmental damage that violates law or regulation.

Failure could directly cause irreversible severe environmental damage that violates law or regulation

Failure could directly cause mitigable environmental damage without violation of law or regulation

Failure could cause minimal environmental damage not violating law or regulation

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G.4. Software Lifecycle Process Requirements. The Software Safety Lifecycle Project Requirements table below summarizes the lifecycle requirements in Section 2 that are required for each SHCI in Section 3.3.

Table 4: SOFTWARE SAFETY LIFECYCLE PROCESS REQUIREMENTS

SHCI Levels

Requirements A B C D System Process (2.1) System Process Requirements (2.1.2) 1. Software Safety Program Plan shall be developed. (SwSPP) X X X X 2. Software Safety Program Plan shall be compliant with the System Safety Program Plan. (SwSPP)

X X X X

3. Engine Specification shall be compliant with the AQP. X X X X 4. System requirements shall be developed. (engine model specification) X X X X 5. System requirements shall be allocated to hardware and software. (SSDD) X X X X 6. Configuration Management archive, retrieval, and release processes shall be established. (DAL) X X X X

7. Integrated master schedule for the system/software development shall be established. (IMS)

X X X X

Functional Hazard Assessment Requirements (2.1.2.1) 1. Aircraft Level FHA shall be developed. {FHA} X X X X 2. System FHA shall be developed. {FHA} X X X X 3. System safety requirements shall be traceable to the FHA. X X X System Safety Assessment Requirements (2.1.2.2)

1. Aircraft Failure Conditions and Safety Requirements shall be developed and allocated. {SSA} X X X X 2. System Failure Conditions and Safety Requirements shall be developed and allocated. {SSA} X X X X 3. System Allocation Analysis shall be traceable to System FHA. {SSA} X X X 4. Safety requirements shall be specified. {engine model specification} X X X X 5. Safety requirements shall be traceable to System FHA. {engine model specification} X X X 6. Software Hazard Criticality Index (SHCI) assignment shall be determined and justification provided. {SSA}

X X X X

Software Planning Process Requirements (2.2.2) Plan for Software Aspects of Certification Requirements (2.2.2.1) 1. Software development and integral process activities shall be defined and agreed upon. (PSAC)

X X X X

2. Transition criteria, inter-relationships, and sequencing among processes shall be defined. (PSAC)

X X X

3. Software lifecycle environment shall be defined. (PSAC) X X X

4. Required resources, schedule, facilities, organizations related to software airworthiness qualification shall be determined and compliant with the engine specification. (PSAC)

X X X X

Software Development Plan Requirements (2.2.2.2) 1. Software development standards shall be defined. (SDP) X X X Software Test Plan Requirements (2.2.2.3) 1. Methods for verification independence shall be established. {STP} X X X X

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SHCI Levels

Requirements A B C D 2. Verification methods shall be defined. {STP} X X X X 3. Test case standards shall be defined. {STP} X X X 4. Test environment(s) shall be defined. {STP} X X X Software Configuration Management Plan Requirements (2.2.2.4) 1. Configuration items shall be identified. (SCMP) X X X X 2. Baselines and traceability shall be established. (SCMP) X X X X 3. Problem reporting, change control, change review, and configuration status accounting shall be established. (SCMP)

X X X X

4. Archive, retrieval, and release shall be established. (SCMP) X X X X 5. Software life cycle environment control shall be established. (SCMP) X X X X Software Quality Assurance Plan Requirements (2.2.2.5) 1. Activities shall be defined to ensure software plans comply with this AQP and established program requirements. (SQAP)

X X X

2. Activities shall be defined to ensure software plans are coordinated. (SQAP) X X X

3. SQA activities shall be defined. (SQAP) X X X X Software Requirements Process Requirements (2.3.2) Software/Interface Requirements Specification Requirements (2.3.2.1) 1. High-level requirements shall be developed. (SRS) X X X X 2. Derived high-level requirements shall be defined. (SRS) X X X X 3. Software high-level requirements shall comply with system requirements. (DAL) O O X X 4. High-level requirements shall be accurate and consistent. (DAL) O O X X 5. High-level requirements shall compatible with target computer. (DAL) X X

6. High-level requirements shall be verifiable. (DAL) X X X

7. High-level requirements shall conform to standards. (DAL) X X X

8. High-level requirements shall be traceable to system requirements. (SRS) X X X X 9. Algorithms shall be accurate. (DAL) O O X

10. High-level safety-critical requirements shall be identified. (SRS) X X X X 11. High-level safety-critical requirements shall be traceable to FHA. (SRS) X X X

12. Safety-critical constraints shall be developed. (SRS) X X X X System Safety Assessment Requirements (2.3.2.2) 1. Software Safety Architectural Analysis shall be performed. (SSA) X X X 2. Software Safety Architectural Analysis shall be traceable to System FHA. (SSA) X X X 3. Safety requirements shall be validated. (SSA) X X X Software/Interface Design Process Requirements (2.4.2) Software/Interface Design Description Requirements (2.4.2.1) 1. Low-level requirements shall be developed. (SDD) X X X 2. Derived low-level requirements shall be defined. (SDD) X X X 3. Software architecture shall be developed. (SDD) X X X 4. Low-level requirements shall comply with high-level requirements. (DAL) O O X 5. Low-level requirements shall be accurate and consistent. (DAL) O O X 6. Low-level requirements shall be compatible with target computer. (DAL) X X

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SHCI Levels

Requirements A B C D 7. Low-level requirements shall be verifiable. (DAL) X X 8. Low-level requirements shall conform to standards. (DAL) X X X 9. Low-level requirements shall be traceable to high-level requirements. (SDD) X X X

10. Algorithms shall be accurate. (DAL) O O X 11. Software architecture shall be compatible with high-level requirements. (DAL) O X X 12. Software architecture shall be consistent. (DAL) O X X 13. Software architecture shall be compatible with target computer. (DAL) X X 14. Software architecture shall be verifiable. (DAL) X X 15. Software architecture shall conform to standards. (DAL) X X X 16. Software partitioning integrity shall be confirmed. (-DAL) O X X X 17. Low-level safety-critical requirements shall be developed. (SDD) X X X 18. Low-level safety-critical requirements shall be traceable to SSA. (SDD) X X X Software Test Description Requirements (2.4.2.2) 1. Test cases shall be developed. (STD) X X X X 2. Test cases shall be traceable to requirements. (STD) X X X X 3. Test cases shall comply with requirements. (DAL) O O X X 4. Test cases shall be complete. (DAL) O O X X

5. Test procedures shall be traceable to the test cases (STD) O O X X System Safety Assessment Requirements (2.4.2.3) 1. Software Interface FMEA is developed. (CDRL SS-SSA) X X X 2. Safety Test Cases and Software Test procedures are consistent with Safety Analyses. (SSA) X X X X Safety Assessment Report Requirements (2.4.2.4) 1. Software-related hazards are controlled and mitigated (preliminary). (SAR) X X X X 2. Control of failure effects is satisfactory (preliminary). (SAR) X X X X 3. Residual risk shall be acceptable (preliminary). (SAR)

X X X X

Software Coding Process Requirements (2.5.2) Source Code Requirements (2.5.2.1) 1. Source code shall be developed. X X X 2. Source code shall comply with low-level requirements. (DAL) O O X 3. Source code shall comply with software architecture. (DAL) O X X 4. Source code shall be verifiable. (DAL) X X 5. Source code shall conform to standards. (DAL) X X X 6. Source code shall be traceable to low-level requirements. (DAL) X X X 7. Source code shall be accurate and consistent. (DAL) O X X 8. Safety-critical algorithms shall be correct. (DAL) O O X 9. Safety-critical source code shall be traceable to hazard analyses. (DAL) X X X Software Test Description Requirements (2.5.2.2) 1. Test procedures shall be correct. (STD) O X X X 2. Test coverage of high-level requirements shall be achieved. (DAL) O X X X 3. Test coverage of low-level requirements shall be achieved. (DAL) O X X

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SHCI Levels

Requirements A B C D 4. Test coverage of software structure (modified condition/decision) shall be achieved. (DAL)

O

5. Test coverage of software structure (decision coverage) shall be achieved. (DAL) O O

6. Test coverage of software structure (statement coverage) shall be achieved. (DAL) O O X

7. Test coverage of software structure (data coupling and control coupling) shall be achieved. (DAL)

O O X

8. Regression test suite shall be developed. (STD) X X X X System Safety Assessment Requirements (2.5.2.3) 1. Software Detailed Causal Analysis shall be developed. (SSA) X X 2. Software Detailed Causal Analysis shall be compliant with the FHA. (SSA) X X Software Integration Process Requirements (2.6.2) Executable Code Requirements (2.6.2.1) 1. Executable object code shall be produced and integrated in the target computer. (SVD)

X X X X

2. Output of software integration process shall be complete and correct. (DAL) X X X Software Version Description Requirements (2.6.2.2) 1. Software version description shall be complete. (SVD) X X X X 2. Software load control shall be established. X X X X Software Test Verification Process Requirements (2.7.2) Software Test Report (2.7.2.1) 1. Test results shall be correct and discrepancies explained. (STR) O X X X 2. High-level requirements test results shall be correct and complete. (STR) O X X X 3. Low-level requirements test results shall be correct and complete. (STR) O X X 4. Coverage results of software structure shall be complete. (STR) O O X 5. Regression test results shall be correct and complete. (STR) O X X X 6. Executable object code shall comply with high-level requirements. (STR) X X X X 7. Executable object code shall be robust with high-level requirements. (STR) X X X X 8. Executable object code shall comply with low-level requirements. (STR) O O X 9. Executable object code shall be robust with low-level requirements. (STR) O X X 10. Executable object code shall be compatible with target computer. (STR) X X X X Software Approval Process (2.8.2) Software Accomplishment Summary Requirements (2.8.2.1) 1. Software development processes shall comply with approved software plans and standards. (SAS)

O O O O

2. Transition criteria for all processes shall be complete. (SAS) O O O

3. Lifecycle processes and data shall be complete. (SAS) { SPS} O O O O Safety Assessment Report Requirements (2.8.2.2) 1. AQSR complies with engine specification. (SAR) O O O X 2. Software-related hazards shall be controlled and mitigated. (SAR) X X X X 3. Control of failure effects shall be satisfactory. (SAR) X X X X 4. Residual risk shall be acceptable. (SAR) X X X X

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O – Must be satisfied with independence X – Must be satisfied. Blank – Does not need to be satisfied.

G.5. Software Change Procedures. Changes to the software after initial airworthiness approval, shall be thoroughly evaluated.

G.5.1 Post Baseline. Upon receiving an initial AWR, no further changes shall be made unless specifically approved by the Airworthiness Authority. From that point forward, each proposed change to requirements and/or software shall be provided to the Airworthiness Authority for review and approval. Editorial changes to requirements or changes to test procedures, shall not require approval, but shall be provided to the Airworthiness Authority for information purposes.

G.5.2 Developmental Flight Test. During flight test, there will be a need to make timely changes to software, particularly in the area of engine/airframe integration dynamics. In some cases, an allowable range of changes may be pre-defined so that alterable constants can be adjusted on the flight line with minimal supporting justification, while other changes may require various levels of verification and validation prior to uploading revised software to the control system. In all cases, a formal change procedure document shall be prepared and approved by the Using Service, the Airworthiness Authority, the Air-framer, and the engine supplier. Note that any changes outside the scope of the approved procedure shall require a supplementary airworthiness release; therefore the procedure shall attempt to address the complete range of foreseeable changes.

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G.6. Programmable Hardware Lifecycle Processes The programmable hardware lifecycle processes and activities below shall be followed to provide design assurance guidance for the development of airborne electronic hardware such that it safely performs its intended function, in its specified environments. All programmable hardware process reviews shall be combined with the appropriate software reviews when possible.

G.6.1 Programmable Hardware Planning Process The Programmable Hardware Planning Process requirements and activities shall be IAW RTCA/DO-254 Section 4.0, and documented in the following hardware planning process documents:

• Plan for Hardware Aspects of Certification (PHAC) • Hardware Design Plan • Hardware Validation Plan • Hardware Verification Plan • Hardware Configuration management Plan (HCMP) • Hardware Process Assurance Plan (HPAP) • Hardware Development Standards

The following process reviews shall be held: • System Requirements Review (SRR) • Programmable Hardware Planning Process Review (PPR)

G.6.1.1 Programmable Hardware Planning Process Artifacts

G.6.1.1.1 Plan for Hardware Aspects of Certification (PHAC) The PHAC shall define the processes, procedures, methods and standards to be used to achieve the requirements of this AQP and obtain airworthiness qualification. The PHAC shall be prepared IAW RTCA/DO-254 Section 10.1.1.

G.6.1.1.2 Hardware Design Plan (HDP) The Hardware design plan shall describe the procedures, methods and standards to be applied, and the processes and activities to be conducted for the design of the hardware item. The hardware design plan shall be prepared IAW DO-254 Section 10.1.2.

G.6.1.1.3 Hardware Validation Plan The Hardware Validation Plan shall describe the procedures, methods and standards to be applied and the processes and activities to be conducted for the validation of the hardware item derived requirements to achieve the validation requirements of this AQP. The Hardware Validation Plan shall be prepared IAW DO-254 Section 10.1.3.

G.6.1.1.4 Hardware Verification Plan The Hardware Verification Plan shall describe the procedures, methods, and standards to be applied and the processes and activities to be conducted for the verification of the hardware items to achieve the verification requirements of this AQP. The Hardware Verification Plan shall be prepared IAW DO-254 Section 10.1.4.

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G.6.1.1.5 Hardware Configuration Management Plan (HCMP) The HCMP shall describe the policies, procedures, standards and methods to be used to satisfy the configuration management requirements of this AQP. The HCMP shall be prepare IAW DO-254 Section 10.1.5.

G.6.1.1.6 Hardware Process Assurance Plan (HPAP) The Hardware Process Assurance Plan shall describe the procedures, methods and standards to be applied and the processes and activities to be conducted for achieving the process assurance requirements of this AQP. The Hardware Process Assurance Plan shall be prepared IAW DO-254 Section 10.1.6

G.6.1.1.7 Hardware Design Standards The Hardware Development Standards document shall identify all requirements standards, hardware design standards, validation and verification standards, and hardware archive standards. These standards documents may be combined into a single document or provided as individual documents.

G.6.1.1.7.1 Requirements Standards Requirements Standards shall be used during the requirements capture process to define the rules, procedures, methods, guidance and criteria for developing the requirements. The Requirements Standards shall be prepared IAW DO-254 Section 10.2.1

G.6.1.1.7.2 Hardware Design Standards Hardware Design Standards shall be used during the conceptual design process and detailed design process to define the rules, procedures, methods, guidance and criteria for developing and specifying the hardware design. The Hardware Design Standards shall be prepared IAW DO-254 Section 10.2.2.

G.6.1.1.7.3 Validation and Verification Standards Hardware Validation and Verification standards shall be used during the validation and verification processes to define the rules, procedures, methods, guidance and criteria for validating and verifying the hardware design and implementation. The Hardware Validation and Verification Standards shall be prepared IAW DO-254 Section 10.2.3.

G.6.1.1.7.4 Hardware Archive Standards Hardware Archive Standards shall be used to define the procedures, methods and criteria used to retain and archive product data and develop and maintain program and project archives. Hardware archive standards shall be prepared IAW DO-254 Section 10.2.4.

G.6.2 Programmable Hardware Design Processes The Programmable Hardware Design Process requirements and activities shall be IAW DO-254 Section 5.0, and documented in the following hardware design process artifacts:

• Programmable Hardware Requirements Document (HRD) • Programmable Design Representation Data (HDD)

The following design reviews shall be held: • Preliminary Design Review (PDR) • Critical Design Review (CDR)

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G.6.2.1 Programmable Hardware Design Process Artifacts

G.6.2.1.1 Hardware Requirements Document (HRD) The HRD shall specify the functional, performance, quality, and safety, maintainability, and reliability requirements for the programmable hardware items. The HRD shall include a traceability matrix that shows requirements traceability between the system level requirements of the SSDD and the high-level programmable hardware requirements. The HRD shall be prepared IAW DO-254 Section 10.3.1.

G.6.2.1.2 Hardware Design Representation Data (HDD) The Hardware Design Representation Data shall provide a definition of the hardware item(s) and is comprised of the set of drawings, documents and specifications used to build the hardware item(s). The Hardware Design Representation Data shall contain Conceptual Design Data and Detail Design Data.

G.6.2.1.2.1 HDD - Conceptual Design Data The Conceptual Design Data shall describe the hardware item’s architecture and functional design. The Conceptual Design Data shall be IAW DO-254 Section 10.3.2.1.

G.6.2.1.2.2 HDD – Detail Design Data The Detailed Design data shall describe all data necessary to implement the hardware item consistently with its requirements. The Detail Design Data shall be IAW DO-254 Section 10.3.2.2, and shall include at a minimum the following:

• Top-level drawings • Assembly drawings • Installation control drawings • Hardware/software interface data

G.6.2.1.2.2.1 Top-Level Drawings Top Level drawings shall uniquely identify the hardware item and identify all assemblies, subassemblies, components and relevant documentation that define the hardware item.

G.6.2.1.2.2.2 Assembly Drawings Assembly drawings shall include additional detailed information needed to assemble the hardware item, assembly, or subassembly. Assembly drawings shall be IAW DO-254 Section 10.3.2.2.2.

G.6.2.1.2.2.3 Installation Control Drawings Installation Control Drawings shall ensure correct installation of the hardware item into a system or correct installation of a hardware item into another hardware item. Installation Control Drawings shall be IAW DO-254 Section 10.3.2.2.3.

G.6.2.1.2.2.4 Hardware/Software Interface Data Hardware/Software Interface Data shall be IAW DO-254 Section 10.3.2.2.4.

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G.6.3 Programmable Hardware Supporting Processes The Programmable Hardware Supporting Processes activities and requirements shall be IAW DO-254 Section 6 thru Section 8, and documented in the following hardware supporting process artifacts:

• Validation and Verification Data • Hardware Acceptance Test Criteria • Problem Reports • Hardware Configuration Management Records • Hardware Process Assurance Records • Hardware Accomplishment Summary (HAS)

The following reviews shall be held: • Test Readiness Review (TRR) • Test Results Review (TReR) • Final Qualification Review (FQR)

G.6.3.1 Programmable Hardware Supporting Processes Artifacts

G.6.3.1.1 Hardware Validation and Verification Data The Hardware Validation and Verification Data shall provide evidence of the completeness and correctness of the hardware design results and of the hardware item itself. The Hardware Validation and Verification Data shall be prepared IAW DO-254 Section 10.4 and shall include:

• Hardware Traceability Data • Hardware Review and Analysis Procedures • Hardware Review and Analysis Results • Hardware Test Procedures • Hardware Test Results

G.6.3.1.1.1 Hardware Traceability Data The Hardware Traceability Data shall establish a correlation between the requirements, detailed design, implementation and verification data that facilitates configuration control, modification and verification of the hardware item. The Hardware Traceability Data shall be prepared IAW DO-254 Section 10.4.1.

G.6.3.1.1.2 Hardware Review and Analysis Procedures The Hardware Review and Analysis Procedures shall define the process and criteria for conducting reviews and analysis. The Hardware Review and Analysis Procedures shall be prepared IAW DO-254 Section 10.4.2.

G.6.3.1.1.3 Hardware Review and Analysis Results The Hardware Review and Analysis Results as shall provide evidence that the reviews and analyses have been completed to approve procedures and criteria. The Hardware Review and Analysis Results shall be prepared IAW DO-254 Section 10.4.3.

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G.6.3.1.1.4 Hardware Test Procedures The Hardware Test Procedures shall define the methods, environment and instructions for conducting both functional and environmental qualification testing used for the verification of the hardware item. The Hardware Test Procedures shall be IAW DO-254 Section 10.4.4.

G.6.3.1.1.5 Hardware Test Results The Hardware Test Results shall provide evidence that the tests have been completed to approved procedures in support of the verification of the hardware items. The Hardware Test Results shall be IAW DO-254 Section 10.4.5.

G.6.3.1.2 Hardware Acceptance Test Criteria The Hardware Acceptance Test Criteria shall provide the criteria and assessment data that the test and associated test results are capable of ensuring that an item is manufactured or repaired correctly. Hardware Acceptance Test Criteria shall be IAW DO-254 Section 10.5.

G.6.3.1.3 Problem Reports Problem Reports shall identify and record the resolution to hardware design problems, process non-compliance with hardware plans and standards, and deficiencies in hardware life cycle data. Problem Reports shall be IAW DO-254 Section 10.6.

G.6.3.1.4 Hardware Configuration Management Records The Hardware Configuration Management Records shall document the results of the configuration management process activities. These activities shall include configuration identification lists, baseline or electronic records, change history reports, problem report summaries, tool identification data, archive records and release records.

G.6.3.1.5 Hardware Process Assurance Records The Hardware Process Assurance Records shall document the results of the process assurance process activities. These activities shall include review or audit reports, meeting minutes, records of authorized process deviations, or conformity review records.

G.6.3.1.6 Hardware Accomplishment Summary (HAS) The HAS shall show compliance with the PHAC and demonstrate to the Airworthiness Authority that the objectives of DO-254 have been achieved for the hardware items. The HAS shall be IAW DO-254 Section 10.9.

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G.7. Testing Requirements and Terminology. Contractors have numerous testing classification types they may employ during the software lifecycle. These include unit tests, integration tests, regression tests, and acceptance tests. These tests do not have clear definitions, as one developer’s notion of a “regression test” may be very different from another developer. Therefore, the distinction of test types shall not be an important issue for the independent test reviewers. Instead, the test reviewers should be concerned only with the submitted test cases and accompanying test procedures. All test cases shall be developed from, trace to, and fully test the requirements. Whether the developer classifies a test as a unit test, module test, closed-loop test, open-loop test, integration test, regression test, etc., it shall trace back and fully test the requirements. The developer shall select a set of test cases that will be run with every software release. This set of tests shall verify the basic system functionality and verify that all safety features execute as expected. The intent of this set of test cases is to ensure that the software change did not inadvertently create a software error. While only a subset of test cases and procedures may be witnessed during Formal Qualification Testing (FQT), all test cases, procedures, and results that ensure correct high-level and low-level requirements implementation and applicable structural coverage shall be submitted with the test artifacts. Any failure that occurs in testing shall not be merely reran once and passed without a thorough examination and justification as to why the failure occurred before just accepting a one-time re-execution. It is encouraged to rerun the failed event multiple times to establish a statistical confidence that the failure was an anomaly of test procedure or environment and not an anomaly of the code. Structural coverage is the task of measuring which parts of the executable software are run during the requirements based testing. This is only effective if the test cases are based on the requirements. After executing the requirements based test cases and determining which portions of the code have not been executed an analysis is performed to determine why portions of the code have not been executed.

G.7.1 Robustness Test Case. The requirements of robustness test cases shall be to demonstrate the ability of the software to respond to abnormal inputs and conditions. Specific requirements includes:

a. Test cases shall be developed to verify out of range conditions for the specific variable and the underlying machine representation. .

b. Test cases shall be developed to verify out of range enumeration values. c. Test cases shall include abnormal system initialization conditions. d. Test cases shall include failed input conditions. This includes out of range analog

inputs and incorrect bus communications. Incorrect bus communications include invalid data, incorrect data timing, and incorrect data sequences.

e. Test cases shall verify scheduler responses to incorrect frame times including

frame overruns.

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f. Test cases shall attempt to provoke overflow of arithmetic operations including

divide-by-zero. g. Test cases shall attempt to provoke invalid state transitions in state transition

diagrams. h. Test cases shall include operations under heavy processing load and under heavy

bus utilization load (> 75%) i. Test cases shall include forcing faults and failovers to verify error handling. Some

of these tests may only be feasible in a lab or simulation environment.

G.7.2 Software Test Environment Validation. The ability of software testing to verify that the software will function free from errors depends on the quality of the software test environment. For most systems, more than one test environment is required to show that each software module meets its requirements and that the integrated system performs without error. The following requirements apply to all software test environments, from host testing environments to system/software integration laboratories (SILs).

a. The differences between the software test environment and the target environment shall be documented and evaluated. The differences shall be summarized in the STR, to allow the reader to draw the proper conclusions about the test results. Factors such as timing, interfaces, memory size, hardware fidelity, and simulation fidelity shall be considered. Exceptions may be made if the condition being tested is not safe to fly.

b. The test environment shall be under configuration control. The exact configuration

of the test environment shall be recorded for each test. The test environment shall be able to be re-created for any required re-testing. Test environment configuration control shall include any software simulations, hardware (including sensors, actuators, electrical harnesses, displays, & data busses), computers, operating systems, configuration tables, and COTS software.

G.7.3 Tool Qualification. DO-178C specifies that the output of tools cannot be assumed to be correct. Tools shall be qualified if they eliminate, reduce, or automate any requirements in Section 2 of this document. If a tool needs to be qualified, the potential impacts of the tool on the lifecycle shall be evaluated against the criteria in DO-178C Section 12.

G.7.4 Model Based Development. The use of models is becoming more common for software development. RTCA DO-331 provides guidance on using the tools and techniques. In general, these techniques and tools may be acceptable for developing airworthy software. All the requirements in this AQP shall still be met. The planning documents shall specify how the tools and lifecycle process map to the requirements.

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G.7.5 Dead and Deactivated Code. Executable object code that is not traceable to any system or software requirements is defined as dead code and shall be removed. Executable object code that is present for different configuration of software is defined as deactivated code and shall be demonstrated that:

a. There are requirements for all of the executable object code b. The deactivation mechanism shall be correct and robust. Code errors, hardware

faults, maintenance, and inadvertent operator selection shall be considered. c. Inadvertent activation of deactivated code shall not cause a system hazard d. All configurations shall be fully verified. All executable object code shall meet the

verification requirements for the SHCI assignment level.

G.7.6 Software Load Integrity. Each version of software shall be uniquely identified and contain correct information to ensure the software is correctly loaded into the LRU. Requirements includes:

a. The load integrity shall use a Cyclic Redundancy Check (CRC) or similar algorithm. A 32-bit CRC is preferred, but the number of bits in the CRC shall meet the safety requirements. A checksum is not reliable enough to provide load integrity for any safety critical software.

b. The software shall not have a date or timestamp embedded in the software. c. The software shall have an embedded software part and version number. d. The software part number, version number, and CRC shall be visible externally.

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FIGURE 1. EXTERNALLY APPLIED FORCES (REF. 3.1.2.5).

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited. Other requests for this document shall be referred to FCDD-AE, 4488 Martin Road, Redstone Arsenal, AL 35898-5000

-11-10

-9-8-7-6-5-4-3-2-1012345678

-6 -4 -2 0 2 4 6 8 10 12

S.L. = 4.0

ψ̈

θS.L. = 1.5.

¨

FLIGHT

FORE

= 6θ

UP

= 0

DOWN

= 0

= 0.ψ

= 2θ.

AFT

+

+

1. LOAD FACTORS AND ANGULAR VELOCITIES ANDACCELERATIONS SHOULD BE TAKEN AT OR ABOUTTHE C.G. OF THE ENGINE.

2. SIDE LOAD FACTORS (SL) ACT TO EITHER SIDE.

3. AND ARE PITCHING VELOCITY (RAD/SEC) ANDACCELERATION (RAD/SEC2).

4. AND ARE YAWING VELOCITY (RAD/SEC) AND ACCELERATION (RAD/SEC2).

5. DOWN LOADS OCCUR DURING PULLOUT.

θ.

θ̈

ψ.

ψ̈

-11-10

-9-8-7-6-5-4-3-2-10 1 2 3 4 5

-6 -4 -2 0 2 4 6 8 10 12

S.L. = 2.0

¨ ψ

θ ψ

AFT

. = 0

. = 0

¨

LANDING

FORE

= ý ±14 θ DOWN

UP

= ý ±6

Applicable to complete rectangle from 7g up to 10g down.

Applicable to complete cross hatched area.

●

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FIGURE 2. ENGINE SURFACE TEMPERATURE VERSUS ENGINE LENGTH.

ENGINE SURFACE TEMPERATURE - TS (deg F)

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FIGURE 3. ENGINE SURFACE EMISSIVITY VERSUS LENGTH (Ref 3.1.2.8).

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FIGURE 4. Not Used.

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IDLE NSD = ______ RPM

-80 -60 -40 -20 0 20 40 60

GASGENERATORSPEED,RPM

RATING POINTS

INLET TOTAL TEMPERATURE, DEG F14012010080

INLET TOTAL TEMPERATURE, DEG C

6040200-20-40-60-80-100

FIGURE 5-1. IDLE GAS GENERATOR SPEED (REF 3.2.1.5.4.1).

Gas

Gen

erat

or S

peed

(RPM

)

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IDLE NSD = ______ RPM

FIGURE 5-2. IDLE OUTPUT SHAFT TORQUE (REF 3.2.1.5.4.1)

-30

-20

-10

0

10

20

30

-80 -60 -40 -20 0 20 40 60

OU

TPU

T SH

AFT

TOR

QU

E, F

T-LB

SEA LEVEL

INLET TOTAL TEMPERATURE, DEG C

-20

INLET TOTAL TEMPERATURE, DEG F

-10

10

0

14012010080

OU

TPU

T SH

AFT

TOR

QU

E, N

-M

6040200-20-40-60-80-100

20

6.1 KM (20014 FT)

4 KM (13122 FT)

2 KM (6362 FT)

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NO LOAD NSD = ______ RPM

-60 -40 -20 0 20 40 60

-60 -40 140INLET TOTAL TEMPERATURE, DEG F

GASGENERATORSPEED,RPM

120100806040200-20

INLET TOTAL TEMPERATURE, DEG C

6.1 KM (20014 FT)SEA LEVEL

3 KM (9842 FT)

FIGURE 6. NO LOAD GAS GENERATOR SPEED (REF 3.2.1.5.4.2)

Gas

Gen

erat

or S

peed

(RPM

)

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FIGURE 7. Engine Altitude and Operating Envelope.

(Ref. 3.2.1.4.1, 3.2.1.4.4, 3.2.1.4.2, 3.2.1.4.3, 3.2.1.4.5, 3.2.5.1, 4.5.1.3, 4.6.1.3).

ALT

ITU

DE

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FIGURE 8. OPERATING LIMITS (REF. 3.2.1.4.1, 3.2.1.4.4).

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-60

-40

-20

0

20

40

60

80

100

-50 -40 -30 -20 -10 0 10 20 30 40 50

LEFTINCLINATION

(ROLL)

RIGHTINCLINATION

(ROLL)

ENGINE ATTITUDE

* ABOVEHORIZONTAL

* BELOWHORIZONTAL

NOTES (1) For the purpose of defining direction of the acceleration vector from the engine CG,

the figure assumes no acceleration other than gravity; the engine shall be capable of operating at all possible acceleration conditions.

(2) * referenced to ground (3) Engine centerline perpendicular to plane of paper. (4) Continuos operation in clear area. (5) 30 second operation in shaded area. (6) Symbol “O” indicates points for test.

FIGURE 9. ENGINE ATTITUDE LIMITS (REF. 3.2.1.5.1, 4.5.4.4, 4.6.4.12)

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Figure 10-1. ESTIMATED ACCELERATION TIME VERSUS ALTITUDE (REF. 3.2.1.5.6.2).

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FIGURE 10-2. ESTIMATED ACCELERATION TIME VERSUS POWER (REF. 3.2.1.5.6.2).

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351

(CONTRACTOR TO PROVIDE.)

FIGURE 10-3A THRU 10-3G POWER ABSORBER LOADING FOR ACCELERATION (REF. 3.2.1.5.6.1, 3.2.1.5.6.1a)

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FIGURE 11. CONTINUOUS MAXIMUM ICING CONDITIONS (REF. 3.2.5.2).

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FIGURE 12. INTERMITTENT MAXIMUM ICING CONDITIONS (REF. 3.2.5.2).

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Diameter, d is the diameter of the vitiated airflow exhaust nozzle at the engine exhaust exit plane m (ft). The exhaust exit plane is the first downstream plane, normal to the exhaust system that does not contain a solid surface around the stream. Smoke may not be visible when viewing the exhaust stream perpendicular to the gas flow.

FIGURE 13. EXHAUST INVISIBILITY LIMIT (REF. 3.2.5.8.1).

Maximum for any gas turbine engine

Visible region of unacceptable smoke emission

Invisible

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FIGURE 14-1. OXIDES OF NITROGEN (REF. 3.2.5.8.2).

PRESENT OR NOX LEVEL

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FIGURE 14-2. GASEOUS EMISSIONS (REF. 3.2.5.8.2).

CO

CxHy

NOx

ENGINE POWER SETTING (% NET POWER)

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FIGURE 15. CONTROL LEVER ANGLES VERSUS CONTROL LEVER TORQUES (REF. 3.7.2.1.1).

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FIGURE 16. CONTROL LIMITER REGIMES (REF. 3.7.2.2.2).

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FIGURE 17. STARTING TORQUE AND SPEED (REF. 3.7.9.1.1, 4.3.3.4, 4.6.4.1, 4.7.1.4).

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FIGURE 18. ENGINE GROUND START TIME VERSUS AMBIENT AIR TEMPERATURE – STATIC, NO RAM, SEA LEVEL TO 5 KM (REF. 3.7.9.2, 4.6.4.1).

Ambi

ent A

ir Te

mpe

ratu

re (°

C)

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FIGURE 19. ENGINE ALTITUDE AND STARTING TEST POINTS (REF. 4.5.3, 4.5.3.2, 4.6.3, 4.6.3.2).

ALT

ITU

DE

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FIGURE 20. Not Applicable.

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FIGURE 21. Not Applicable.

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FIGURE 22. Not Applicable.

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FIGURE 23. ENGINE CORROSION OPERATING CYCLE (REF. 4.6.4.3).

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FIGURE 24. AMBIENT TEMPERATURE EXTREMES VS. ALTITUDE (REF. 3.2.1, 3.2.5.1).

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FIGURE 25. JET WAKE (REF. 3.7.10.3).

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368

(CONTRACTOR TO PROVIDE.)

FIGURE 26. REFERENCE EXHAUST NOZZLE (REF. 3.2.5.8.1, 3.7.10.2).

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FIGURE 27. Not Used.

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370

(CONTRACTOR TO PROVIDE.)

FIGURE 28. COMPONENT SURFACE LIMIT TEMPERATURES.

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371

(CONTRACTOR TO PROVIDE.)

FIGURE 29. ENGINE VIBRATION SPECTRUM (REF. 3.2.1.4.10, 4.7.5.3.1).

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FIGURE 30-1. NEAR FIELD OCTAVE BAND SOUND PRESSURE LEVEL CONTOURS (DB - REF. 0.0002 µBARS) CENTER FREQUENCY - 250 Hz

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FIGURE 30-2. FAR FIELD OVERALL SOUND PRESSURE LEVEL CONTOURS (DB - REF. 0.0002 µBARS).

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FIGURE 30-3. ESTIMATED OVERALL SOUND PRESSURE LEVEL CONTOURS AT IDLE.

EXHAUST

ANGLE FROM EXHAUST ENGINE AXIS

ENGINE FACE

OASPL in dB re 0.0002 dyne/cm2

SEA LEVEL STATIC STANDARD DAY CONDITIONS