EASA Part 66 - Module 6 - Materials and Hardware

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JAR 66 CATEGORY B1 MODULE 6 MATERIALS AND HARDWARE

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1 INTRODUCTION ... 1-1 2 PROPERTIES OF METALS ... 2-1 2.1 BRITTLENESS ... 2-1 2.2 CONDUCTIVITY ... 2-1 2.3 DUCTILITY ... 2-1 2.4 ELASTICITY ... 2-1 2.5 HARDNESS ... 2-1 2.6 MALLEABILITY... 2-1 2.7 PLASTICITY ... 2-1 2.8 TENACITY ... 2-1 2.9 TOUGHNESS ... 2-2 2.10 STRENGTH ... 2-2 2.10.1 Tensile Strength ... 2-2 2.10.2 Yield Strength ... 2-2 2.10.3 Shear Strength ... 2-2 2.10.4 Bearing Strength ... 2-2 3 TESTING OF MATERIALS ... 3-1 3.1 TENSILE TESTING ... 3-1 3.1.1 Tensile Strength ... 3-1 3.2 LOAD/EXTENSION DIAGRAMS ... 3-4 3.2.1 Ductility ... 3-7 3.2.2 Proof Stress ... 3-7 3.3 STIFFNESS ... 3-9 3.4 TENSILE TESTING OF PLASTICS ... 3-9 3.5 COMPRESSION TEST ... 3-10 3.6 HARDNESS TESTING ... 3-10 3.6.1 Brinell Test ... 3-10 3.6.2 Vickers Test ... 3-11 3.6.3 Rockwell Test ... 3-11 3.6.4 Hardness Testing on Aircraft ... 3-12 3.7 IMPACT TESTING ... 3-13 3.8 OTHER FORMS OF MATERIAL TESTING ... 3-14 3.8.1 Creep ... 3-14 3.8.2 Creep in Metals ... 3-14 3.8.3 Effect of Stress and Temperature on Creep ... 3-15 3.8.4 The Effect of Grain Size on Creep ... 3-16 3.8.5 Creep in Plastics ... 3-16 3.8.6 Fatigue ... 3-16 3.8.7 Fatigue Testing ... 3-17 3.9 S-N CURVES ... 3-18 3.10 CAUSES OF FATIGUE FAILURE ... 3-20

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3.12 FATIGUE METALLURGY ... 3-21 3.13 FATIGUE PROMOTERS ... 3-22 3.13.1 Design ... 3-22 3.13.2 Manufacture ... 3-23 3.13.3 Environment ... 3-23 3.14 FATIGUE PREVENTERS ... 3-23 3.14.1 Cold Expansion (Broaching) ... 3-24 3.15 DO'S AND DONT'S – PREVENTING FATIGUE FAILURES ... 3-25 3.16 STRUCTURAL HEALTH MONITORING (SHM) ... 3-25 3.16.1 Fatigue Meters ... 3-25 3.16.2 Strain Gauges ... 3-25 3.16.3 Fatigue Fuses ... 3-25 3.16.4 Intelligent Skins Development ... 3-25

4 AIRCRAFT MATERIALS - FERROUS ... 4-1

4.1 IRON ... 4-1 4.1.1 Cast Iron ... 4-1 4.1.2 Nodular Cast Iron ... 4-1 4.2 STEEL ... 4-1 4.2.1 Classification of Steels ... 4-2 4.2.2 Metallurgical Structure of Steel ... 4-3 4.2.3 Structure and Properties – Slow-Cooled Steels ... 4-3 4.2.4 Effects of Cooling Rates on Steels ... 4-4 4.3 HEAT-TREATMENT OF CARBON STEELS ... 4-4 4.3.1 Associated Problems - Hardening Process ... 4-5 4.3.2 Tempering ... 4-6 4.3.3 Annealing ... 4-6 4.3.4 Normalising ... 4-6 4.4 SURFACE HARDENING OF STEELS ... 4-7 4.4.1 Carburising... 4-7 4.4.2 Nitriding ... 4-8 4.4.3 Flame/Induction Hardening ... 4-8 4.4.4 Other Surface Hardening Techniques ... 4-8 4.5 ALLOYING ELEMENTS IN STEEL ... 4-9 4.6 CARBON ... 4-9 4.6.1 Low-Carbon Steel ... 4-9 4.6.2 Medium-Carbon Steel ... 4-9 4.6.3 High-Carbon Steel ... 4-9 4.7 SULPHUR ... 4-9 4.8 SILICON ... 4-9 4.9 PHOSPHORUS ... 4-10 4.10 NICKEL ... 4-10 4.10.1 Nickel Alloys ... 4-10 4.11 CHROMIUM (CHROME) ... 4-11 4.11.1 Nickel-Chrome Steel and its Alloys ... 4-11 4.12 COBALT ... 4-11 4.13 VANADIUM ... 4-12

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4.14 MANGANESE ... 4-12 4.15 MOLYBDENUM ... 4-12 4.16 CHROME AND MOLYBDENUM ... 4-12 4.17 TUNGSTEN ... 4-13 4.18 MARAGING STEELS... 4-13

5 AIRCRAFT MATERIALS - NON-FERROUS ... 5-1

5.1 PURE METALS ... 5-1 5.1.1 Pure Aluminium ... 5-1 5.1.2 Pure Copper ... 5-2 5.1.3 Pure Magnesium ... 5-2 5.1.4 Pure Titanium ... 5-2 5.2 ALUMINIUM ALLOYS ... 5-3 5.3 IDENTIFICATION OF ELEMENTS IN ALUMINIUM ALLOYS ... 5-3 5.4 CLAD MATERIALS ... 5-5 5.5 HEAT-TREATMENT OF ALUMINIUM ALLOYS ... 5-5 5.5.1 Solution Treatment ... 5-6 5.5.2 Age-Hardening ... 5-7 5.5.3 Annealing ... 5-7 5.5.4 Precipitation Treatment ... 5-8 5.6 IDENTIFICATION OF HEAT-TREATED ALUMINIUM ALLOYS ... 5-9 5.7 MARKING OF ALUMINIUM ALLOY SHEETS ... 5-10 5.8 CAST ALUMINIUM ALLOYS ... 5-11 5.9 MAGNESIUM ALLOYS ... 5-11 5.10 COPPER ALLOYS ... 5-12 5.11 TITANIUM ALLOYS ... 5-13 5.12 WORKING WITH TITANIUM AND TITANIUM ALLOYS ... 5-13 5.12.1 Drilling Titanium ... 5-14

6 METHODS USED IN SHAPING METALS ... 6-1

6.1 CASTING ... 6-1 6.1.1 Sand-Casting ... 6-1 6.1.2 Advantages/Disadvantages of Sand-Casting ... 6-3 6.1.3 Typical Casting Defects... 6-3 6.1.4 Shell-Moulding ... 6-3 6.1.5 Centrifugal-Casting ... 6-3 6.1.6 Die-Casting ... 6-4 6.1.7 Investment-Casting (Lost Wax) ... 6-4 6.2 FORGING ... 6-5 6.2.1 Drop-Stamping ... 6-6 6.2.2 Hot-Pressing ... 6-6 6.2.3 Upsetting ... 6-6 6.3 ROLLING ... 6-7 6.4 DRAWING ... 6-7 6.5 DEEP DRAWING/PRESSING ... 6-7 6.6 PRESSING ... 6-7

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6.8 RUBBER-PAD FORMING ... 6-7 6.9 EXTRUDING ... 6-8 6.9.1 Impact-Extrusion ... 6-8 6.10 SINTERING ... 6-8 6.11 SPINNING ... 6-9 6.12 CHEMICAL MILLING ... 6-9 6.13 ELECTRO-CHEMICAL MACHINING ... 6-9 6.14 ELECTRO-DISCHARGE MACHINING E.D.M. ... 6-10 6.15 CONVENTIONAL MACHINING ... 6-11 6.16 SUPERPLASTIC FORMING ... 6-12

7 AIRCRAFT MATERIALS - COMPOSITE AND NON-METALLIC ... 7-1

7.1 PLASTICS ... 7-1 7.1.1 Thermoplastic Materials ... 7-2 7.1.2 Thermosetting Materials ... 7-3 7.1.3 Resins ... 7-4 7.1.4 Elastomers ... 7-6 7.2 PRIMARY ADVANTAGES OF PLASTICS ... 7-7 7.3 PRIMARY DISADVANTAGES OF PLASTICS ... 7-7 7.4 PLASTIC MANUFACTURING PROCESSES ... 7-8 7.5 COMPOSITE MATERIALS ... 7-9 7.5.1 Glass Fibre Reinforced Plastic (GFRP) ... 7-9 7.5.2 Carbon Fibre Reinforced Plastic (CFRP) ... 7-10 7.5.3 Aramid Fibre Reinforced Plastic (AFRP) ... 7-11 7.5.4 General Information ... 7-11 7.5.5 Laminated, Sandwich and Monolithic Structures ... 7-12 7.6 NON-METALLIC COMPONENTS ... 7-13 7.6.1 Seals ... 7-13

8 DETECTING DEFECTS IN COMPOSITE MATERIALS... 8-1

8.1 CAUSES OF DAMAGE ... 8-1 8.2 TYPES OF DAMAGE... 8-1 8.3 INSPECTION METHODS ... 8-3 8.3.1 Visual Inspection ... 8-3 8.3.2 Ring or Percussion Test ... 8-3 8.3.3 Ultrasonic Inspection ... 8-3 8.3.4 Radiography ... 8-3 8.4 ASSESSMENT OF DAMAGE ... 8-4

9 BASIC COMPOSITE REPAIRS ... 9-1

9.1 REPAIR OF A SIMPLE COMPOSITE PANEL ... 9-2 9.2 REPAIR OF A SANDWICH PANEL ... 9-3 9.3 GLASS FIBRE REINFORCED COMPOSITE REPAIRS ... 9-5 9.4 TYPES OF GLASS REINFORCEMENT ... 9-5 9.4.1 Uni-Directional Cloth ... 9-5 9.4.2 Bi-directional Cloth ... 9-6 9.4.3 Chopped Strand Mat ... 9-6 9.4.4 Resin ... 9-6

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9.5 POT LIFE... 9-7 9.6 CURING ... 9-7 9.7 GEL COAT ... 9-8 9.8 STORAGE OF GFRP MATERIALS ... 9-8 9.8.1 Storing Resin ... 9-8 9.8.2 Storing Hardener ... 9-8 9.8.3 Storing Fabrics ... 9-8 9.9 PREPARATION FOR REPAIR ... 9-9 9.9.1 Surface Preparation ... 9-11 9.10 TECHNIQUES OF LAMINATING GLASS FIBRE ... 9-11 9.11 PRE-WETTING GLASS FIBRE ... 9-12

10 ADHESIVES AND SEALANTS ... 10-1

10.1 THE MECHANICS OF BONDING ... 10-1 10.1.1 Stresses on a Bonded Joint ... 10-1 10.1.2 Advantages of Adhesives ... 10-3 10.1.3 Disadvantages of Adhesives ... 10-3 10.1.4 Strength of Adhesives ... 10-4 10.2 GROUPS AND FORMS OF ADHESIVES ... 10-4 10.2.1 Flexible Adhesives ... 10-4 10.2.2 Structural Adhesives ... 10-4 10.2.3 Adhesive Forms ... 10-4 10.3 ADHESIVES IN USE ... 10-5 10.3.1 Surface Preparation ... 10-5 10.3.2 Final Assembly ... 10-5 10.3.3 Typical (Abbreviated) Process ... 10-6 10.4 SEALING COMPOUNDS ... 10-6 10.4.1 One-Part Sealants ... 10-7 10.4.2 Two-Part Sealants ... 10-7 10.4.3 Sealant Curing ... 10-7

11 CORROSION ... 11-1

11.1 CHEMICAL (OXIDATION) CORROSION... 11-1 11.1.1 Effect of Oxide Thickness ... 11-2 11.1.2 Effect of Temperature ... 11-3 11.1.3 Effect of Alloying ... 11-4 11.2 ELECTROCHEMICAL (GALVANIC) CORROSION ... 11-5 11.2.1 The Galvanic Cell ... 11-5 11.2.2 Factors Affecting the Rate of Corrosion in a Galvanic Cell. ... 11-6 11.3 TYPES OF CORROSION ... 11-8 11.3.1 Surface Corrosion ... 11-8 11.3.2 Dissimilar Metal Corrosion ... 11-8 11.3.3 Intergranular Corrosion ... 11-9 11.3.4 Exfoliation Corrosion ...11-10 11.3.5 Stress Corrosion ...11-10 11.3.6 Fretting Corrosion ...11-11 11.3.7 Crevice Corrosion ...11-11 11.3.8 Filiform Corrosion ...11-11

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11.3.10 Corrosion Fatigue...11-13 11.3.11 Microbiological Contamination...11-13 11.3.12 Hydrogen Embrittlement of Steels ...11-13 11.4 FACTORS AFFECTING CORROSION ...11-14 11.4.1 Climatic ...11-14 11.4.2 Size and Type of Metal ...11-14 11.4.3 Corrosive Agents...11-14 11.5 COMMON METALS AND CORROSION PRODUCTS ...11-15 11.5.1 Iron and Steel ...11-15 11.5.2 Aluminium Alloys ...11-15 11.5.3 Magnesium Alloys ...11-16 11.5.4 Titanium ...11-16 11.5.5 Copper Alloys ...11-16 11.5.6 Cadmium and Zinc ...11-16 11.5.7 Nickel and Chromium ...11-17 11.6 CORROSION REMOVAL ...11-17 11.6.1 Cleaning and Paint Removal. ...11-17 11.6.2 Corrosion of Ferrous Metals ...11-18 11.6.3 High-Stressed Steel Components ...11-18 11.6.4 Aluminium and Aluminium Alloys ...11-18 11.6.5 Alclad ...11-19 11.6.6 Magnesium Alloys ...11-19 11.6.7 Acid Spillage ...11-20 11.6.8 Alkali Spillage ...11-20 11.6.9 Mercury Spillage ...11-21 11.7 PERMANENT ANTI-CORROSION TREATMENTS ...11-22 11.7.1 Electro-Plating ...11-22 11.7.2 Sprayed Metal Coatings ...11-22 11.7.3 Cladding ...11-22 11.7.4 Surface Conversion Coatings ...11-23 11.8 LOCATIONS OF CORROSION IN AIRCRAFT ...11-23 11.8.1 Exhaust Areas ...11-23 11.8.2 Engine Intakes and Cooling Air Vents ...11-23 11.8.3 Landing Gear ...11-24 11.8.4 Bilge and Water Entrapment Areas ...11-24 11.8.5 Recesses in Flaps and Hinges ...11-24 11.8.6 Magnesium Alloy Skins ...11-24 11.8.7 Aluminium Alloy Skins ...11-24 11.8.8 Spot-Welded Skins and Sandwich Constructions ...11-25 11.8.9 Electrical Equipment ...11-25 11.8.10 Miscellaneous Items ...11-25 12 AIRCRAFT FASTENERS ... 12-1 12.1 TEMPORARY JOINTS ... 12-1 12.2 PERMANENT JOINTS ... 12-1 12.3 FLEXIBLE JOINTS ... 12-1 12.4 SCREW THREADS ... 12-2 12.4.1 The Inclined Plane and the Helix ... 12-2 12.5 SCREW THREAD TERMINOLOGY ... 12-4

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12.5.1 Screw Thread Forms ... 12-6 12.5.2 Other Thread Forms ... 12-8 12.5.3 Classes of Fit ... 12-8 12.5.4 Measuring Screw Threads ... 12-9 12.6 BOLTS ...12-10 12.6.1 British Bolts ...12-10 12.6.2 Identification of BS Unified Bolts ...12-10 12.6.3 American Bolts ...12-13 12.6.4 Identification of AN Standard Bolts ...12-14 12.6.5 Special-to-Type Bolts ...12-16 12.6.6 Metric Bolts ...12-17 12.7 NUTS ...12-18 12.7.1 Stiffnuts and Anchor Nuts ...12-19 12.8 SCREWS ...12-22 12.8.1 Machine Screws ...12-22 12.8.2 Structural Screws ...12-24 12.8.3 Self-Tapping Screws ...12-24 12.9 STUDS ...12-25 12.9.1 Standard Studs ...12-26 Waisted Studs ...12-26 12.9.3 Stepped Studs ...12-27 12.9.4 Shouldered Studs...12-27 12.10 THREAD INSERTS ...12-27 12.10.1 Wire Thread Inserts ...12-27 12.10.2 Thin Wall Inserts ...12-28 12.11 DOWELS AND PINS ...12-29 12.11.1 Dowels ...12-29 12.11.2 Roll Pins ...12-29 12.11.3 Clevis Pins ...12-30 12.11.4 Taper Pins...12-30 12.12 LOCKING DEVICES ...12-31 12.12.1 Spring Washers ...12-31 12.12.2 Shake-Proof Washers ...12-32 12.12.3 Tab Washers ...12-33 12.12.4 Lock Plates ...12-34 12.12.5 Split (Cotter) Pins ...12-34 12.13 LOCKING WIRE ...12-35 12.13.1 Use of Locking Wire with Turnbuckles ...12-37 12.13.2 Use of Locking Wire with Locking Tabs. ...12-37 12.13.3 Thin Copper Wire ...12-38 12.14 QUICK-RELEASE FASTENERS ...12-38 12.14.1 Dzus Fasteners ...12-38 12.14.2 Oddie Fasteners ...12-39 12.14.3 Camloc Fasteners ...12-40 12.14.4 Airloc Fasteners ...12-41 12.14.5 Pip-Pins ...12-41 12.14.6 Circlips and Locking Rings ...12-42

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12.14.8 Peening ...12-44 12.15 GLUE/ADHESIVE BONDED JOINTS ...12-45 12.15.1 Locking by Adhesives ...12-45 12.15.2 Loctite ...12-46 12.15.3 Synthetic Resin Adhesives ...12-46 12.15.4 Testing of Adhesive Joining Techniques ...12-46 12.16 METAL-TO-METAL BONDED JOINTS ...12-46 12.16.1 Welding ...12-46 12.16.2 Soft Soldering ...12-47 12.16.3 Hard Soldering ...12-47 13 AIRCRAFT RIVETS ... 13-1 13.1 SOLID RIVETS ... 13-1 13.2 RIVET IDENTIFICATION ... 13-2 13.2.1 Solid Rivets (British) ... 13-2 13.2.2 Rivet Identification (British) ... 13-3 13.2.3 Rivet Material Identification (British) ... 13-3 13.2.4 Solid Rivets (American) ... 13-5 13.2.5 Rivet Identification (American)... 13-6 13.2.6 Rivet Material Identification (American) ... 13-6 13.3 HEAT-TREATMENT/REFRIGERATION OF SOLID RIVETS ... 13-7 13.3.1 Heat-Treatment. ... 13-8 13.3.2 Refrigeration. ... 13-8 13.3.3 Use of Different Types of Rivet Head ... 13-8 13.4 BLIND AND HOLLOW RIVETS ... 13-9 13.4.1 Friction Lock Rivets ...13-10 13.4.2 Mechanical Lock Rivets...13-11 13.4.3 Hollow/Pull-Through Rivets ...13-12 13.4.4 Grip Range...13-12 13.4.5 Tucker ‘Pop’ Rivets ...13-13 13.4.6 Avdel Rivets ...13-14 13.4.7 Chobert Rivets ...13-15 13.4.8 Cherry Rivets ...13-16 13.5 MISCELLANEOUS FASTENERS ...13-16 13.5.1 Hi-Lok Fasteners ...13-16 13.5.2 Hi-Tigue Fasteners ...13-17 13.5.3 Hi-Shear Fasteners ...13-18 13.6 SPECIAL PURPOSE FASTENERS ...13-19 13.6.1 Jo-Bolts ...13-19 13.6.2 Tubular Rivets. ...13-20 13.6.3 Rivnuts ...13-21

14 SPRINGS ... 14-1

14.1 FORCES EXERTED ON, AND APPLIED BY, SPRINGS ... 14-1 14.2 TYPES OF SPRINGS ... 14-1 14.2.1 Flat Springs ... 14-1 14.2.2 Leaf Springs ... 14-2 14.2.3 Spiral Springs ... 14-2 14.2.4 Helical Compression and Tension Springs ... 14-2 14.2.5 Helical Torsion Springs ... 14-2

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14.2.6 Belleville (Coned Disc) Springs ... 14-2 14.2.7 Torsion-Bar Springs ... 14-2 14.3 MATERIALS FROM WHICH SPRINGS ARE MANUFACTURED ... 14-2 14.3.1 Steels used for Cold-Wound Springs... 14-2 14.3.2 Steels used for Hot-Wound Springs ... 14-3 14.3.3 Steels used for Cold-Rolled, Flat Springs ... 14-3 14.3.4 Non-Ferrous Metals used for Springs ... 14-3 14.3.5 Composite Materials used for Springs ... 14-4 14.4 CHARACTERISTICS OF TYPICAL AEROSPACE SPRINGS ... 14-5 14.5 APPLICATIONS OF SPRINGS IN AIRCRAFT ENGINEERING ... 14-6

15 PIPES AND UNIONS ... 15-1

15.1 RIGID PIPES ... 15-1 15.2 SEMI-RIGID FLUID LINES (TUBES) ... 15-2 15.2.1 Flared End-Fittings ... 15-2 15.2.2 Flare-Less Couplings ... 15-3 15.3 FLEXIBLE PIPES (HOSES) ... 15-4 15.3.1 Low-Pressure Hoses ... 15-5 15.3.2 Medium-Pressure Hoses ... 15-5 15.3.3 High-Pressure Hoses ... 15-6 15.4 UNIONS AND CONNECTORS ... 15-7 15.4.1 Aircraft General Standards (AGS) ... 15-8 15.4.2 Air Force and Navy (AN) ... 15-8 15.4.3 Military Standard (MS) ... 15-8 15.5 QUICK-RELEASE COUPLINGS ... 15-8 16 BEARINGS ... 16-1 16.1 BALL BEARINGS ... 16-2 16.1.1 Radial Bearings ... 16-2 16.1.2 Angular-Contact Bearings ... 16-2 16.1.3 Thrust Bearings ... 16-2 16.1.4 Instrument Precision Bearings... 16-2 16.2 ROLLER BEARINGS ... 16-3 16.2.1 Cylindrical Roller Bearings ... 16-3 16.2.2 Spherical Roller Bearings ... 16-3 16.2.3 Tapered Roller Bearings ... 16-3 16.3 BEARING INTERNAL CLEARANCE ... 16-4 16.3.1 Group 2 (‘One Dot’) Bearings ... 16-4 16.3.2 Normal Group (‘Two Dot’) Bearings ... 16-4 16.3.3 Group 3 (‘Three Dot’) Bearings ... 16-4 16.3.4 Group 4 (‘Four Dot’) Bearings ... 16-4 16.4 BEARING MAINTENANCE ... 16-5 16.4.1 Lubrication ... 16-5 16.4.2 Inspection ... 16-5

17 TRANSMISSIONS ... 17-1

17.1 BELTS AND PULLEYS ... 17-1 17.2 GEARS ... 17-3

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17.2.3 Helical Gears ... 17-4 17.2.4 Bevel Gears ... 17-4 17.2.5 Worm and Wheel Gears ... 17-4 17.2.6 Planetary (Epicyclic) Reduction Gear Train ... 17-5 17.2.7 Spur and Pinion Reduction Gear Train ... 17-6 17.2.8 Accessory Unit Drives ... 17-6 17.2.9 Meshing Patterns ... 17-7 17.3 CHAINS AND SPROCKETS ... 17-8 17.3.1 Typical Arrangements - Chain Assemblies ... 17-9 17.4 MAINTENANCE INSPECTIONS ...17-10

18 CONTROL CABLES... 18-1

18.1 TYPES OF CABLES ... 18-1 18.2 CABLE SYSTEM COMPONENTS ... 18-2 18.2.1 End-Fittings ... 18-2 18.2.2 Turnbuckles ... 18-3 18.2.3 Cable Tensioning Devices ... 18-4 18.2.4 Cable Fairleads ... 18-5 18.2.5 Pulleys ... 18-6 18.3 FLEXIBLE CONTROL SYSTEMS ... 18-7 18.3.1 Bowden Cables ... 18-7 18.3.2 Teleflex Control Systems ... 18-9

19 ELECTRICAL CABLES & CONNECTORS ... 19-1

19.1 CABLE SPECIFICATION ... 19-1 19.2 CABLE IDENTIFICATION ... 19-1 19.3 DATA BUS CABLE ... 19-5 19.4 CONDUCTOR MATERIAL & INSULATION ... 19-6 19.5 WIRE SIZE ... 19-7 19.6 WIRE RESISTANCE ... 19-8 19.7 CURRENT CARRYING CAPABILITY ... 19-8 19.8 VOLTAGE DROP ...19-10 19.9 WIRE IDENTIFICATION ...19-11 19.10 WIRE INSTALLATION AND ROUTING ...19-12 19.11 OPEN WIRING ...19-12 19.12 WIRE & CABLE CLAMPING ...19-13 19.13 CONDUIT ...19-14 19.14 CONNECTORS ...19-16 19.15 CRIMPING ...19-19 19.16 CRIMPING TOOLS ...19-20 19.17 WIRE SPLICING ...19-21 19.18 BEND RADIUS ...19-22

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1 INTRODUCTION

The variety of materials and hardware used in aircraft engineering is vast, and this module will only deal with a broad group of materials, their main characteristics, identification and uses. These materials can be classed into the three main categories of Ferrous Metals, Non-Ferrous Metals and Non-Metallic materials.

Additionally, combinations (Composites) of many of these materials will be found, in use, in the aerospace industry.

The usefulness of any materials may be enhanced as a result of the addition of other materials that alter the basic characteristics to suit the specific requirements of the aircraft designer.

A metal’s usefulness is governed principally by the physical properties it possesses. Those properties depend upon the composition of the metal, which can be changed considerably by alloying it with other metals and by heat-treatment. The strength and hardness of steel, for example, can be intensified by increasing its carbon content, adding alloying metals such as Nickel and Tungsten, or by heating the steel until red-hot and then cooling it rapidly.

Apart from the basic requirement of more and more strength from metals, other, less obvious characteristics can also be added or improved upon, when such features as permanent magnetism, corrosion resistance and high-strength whilst operating at elevated temperatures, are desired.

Composites make up a large part of the construction of modern aircraft. In the early days, composites and plastics were limited to non-structural, internal cosmetic panels, small fairings and other minor parts. Today there are many large aircraft, which have major structural and load-carrying parts manufactured from composites. Composite materials, in addition to maintaining or increasing component strength, contribute to the important factor of weight saving. There are also many modern light aircraft that are almost totally manufactured from composites and contain little metal at all.

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INTENTIONALLYBLANK

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2 PROPERTIES OF METALS

The various properties of metals can be assessed, by accurate laboratory tests on sample pieces. The terminology, associated with these properties, is outlined in the following paragraphs.

2.1 BRITTLENESS

The tendency of the metal to shatter, without significant deformation. It will shatter under a sudden, low stress but will resist a slowly-applied, higher load.

2.2 CONDUCTIVITY

The ability of a metal to conduct heat, (thermal conductivity) and electricity. Silver and copper are excellent thermal and electrical conductors.

2.3 DUCTILITY

The property of being able to be permanently extended by a tensile force. It is measured during a tensile, or stretching, test, when the amount of stretch (elongation), for a given applied load, provides an indication of a metal’s ductility.

2.4 ELASTICITY

The ability of a metal to return to its original shape and size after the removal of any distorting force. The ‘Elastic Limit’ is the greatest force that can be applied without permanent distortion.

2.5 HARDNESS

The ability of a metal to resist wear and penetration. It is measured by pressing a hardened steel ball or diamond point into the metal’s surface. The diameter or depth of the resulting indentation provides an indication of the metal’s hardness.

2.6 MALLEABILITY

The ease with which the metal can be forged, rolled and extruded without fracture. Stresses, induced into the metal, by the forming processes, have to be subsequently relieved by heat-treatment. Hot metal is more malleable than cool metal.

2.7 PLASTICITY

The ability to retain a deformation after the load producing it has been removed. Plasticity is, in fact, the opposite of elasticity.

2.8 TENACITY

The property of a metal to resist deformation when subjected to a tensile load. It is proportional to the maximum stress required to cause the metal to fracture.

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2.9 TOUGHNESS

The ability of a metal to resist suddenly applied loads. A metal’s toughness is tested by impact with a swinging pendulum of known mass.

2.10 STRENGTH

There are several different measurements of the strength of a metal, as may be seen from the following sub-paragraphs

2.10.1 TENSILE STRENGTH

The ability to resist tension forces applied to the metal

2.10.2 YIELD STRENGTH

The ability to resist deformation. After the metal yields, it is said to have passed its yield point.

2.10.3 SHEAR STRENGTH

The ability to resist side-cutting loads - such as those, imposed on the shank of a rivet, when the materials it is joining attempt to move apart in a direction normal to the longitudinal axis of the rivet.

2.10.4 BEARING STRENGTH

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3 TESTING OF MATERIALS

The mechanical properties of a material must be known before that material can be incorporated into any design. Mechanical property data is compiled from extensive material testing. Various tests are used to determine the actual values of material properties under different loading applications and test conditions.

3.1 TENSILE TESTING

Tensile testing is the most widely-used mechanical test. It involves applying a steadily increasing load to a test specimen, causing it to stretch until it eventually fractures. Accurate measurements are taken of the load and extension, and the results are used to determine the strength of the material. To ensure uniformity of test results, the test specimens used must conform to standard dimensions and finish as laid down by the appropriate Standards Authority (BSI, DIN, ISO etc). The cross-section of the specimen may be round or rectangular, but the

relationship between the cross-sectional area and a specified "gauge length", of each specimen, is constant. The gauge length, is that portion of the parallel part of the specimen, which is to be used for measuring the subsequent extension during and/or after the test.

3.1.1 TENSILE STRENGTH

Tensile strength in a material is obtained by measuring the maximum load, which the test piece is able to sustain, and dividing that figure by the original cross-sectional area (c.s.a.) of the specimen. The value derived from this simple calculation is called STRESS.

Note: The units of Stress may be quoted in the old British Imperial (and

American) units of lbf/in2, tonf/in2 (also psi and tsi), or the European and SI units such as kN/m2, MN/m2 and kPa or MPa.

) 2 (mm c.s.a. Original (N) Load Stress =

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Example 1

A steel rod, with a diameter of 5 mm, is loaded in tension with a force of 400 N. Calculate the tensile stress.

Exercise 1

Calculate the tensile stress in a steel rod, with a cross-section of 10 mm x 4 mm, when it is subjected to a load of 100 N.

Exercise 2

Calculate the cross-sectional area of a tie rod which, when subjected to a load of 2,100N, has a stress of 60 N/mm2.

Note: When calculating stress in large structural members, it may be more

convenient to measure load in Mega-Newtons (MN, or N6) and the area in square metres (m2). When using such units, the numerical value is identical to that if the calculation had been made using Newtons and mm2.

i.e. A Stress of 1 N/mm2 = l MN/m2 Example 2

A structural member, with a cross-sectional area of 0⋅5m2, is subjected to a load of 10 MN. Calculate the stress in the member in; (a) MN/m2 and (b) N/mm2 (a) (b) Area Load Stress = 2 2 2 20 37 / 5 2 400 400 mm N r =_Π_× _⋅ = ⋅ Π Area Load Stress = 20MN/m2 5 0 10 = = = = ⋅⋅⋅⋅ 2 2 2 ₁_MN/m _So_Stress ₂₀_N/mm N/mm ₌ ₌ 1

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As the load in the tensile test is increased from zero to a maximum value, the material extends in length. The amount of extension, produced by a given load, allows the amount of induced STRAIN to be calculated. Strain is calculated by measuring the extension and dividing by the original length of the material. Note: Both measurements must be in the same units, though, since Strain is a

ratio of two lengths, it has no units.

Example 3

An aluminium test piece is marked with a 20 mm gauge length. It is subjected to tensile load until its length becomes 21⋅15 mm. Calculate the induced strain.

Exercise 3

A tie rod 1.5m long under a tensile load of 500 N extends by 12 mm. Calculate the strain. Length Original Extension Strain = mm 15 1 20 -15 21 Extension= ⋅ = ⋅ units) (no 0575 0 20 15 1 Length Original Extension Strain ==== ==== ⋅⋅⋅⋅ ==== ⋅⋅⋅⋅

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3.2 LOAD/EXTENSION DIAGRAMS

If a gradually increasing tensile load is applied to a test piece while the load and extension are continuously measured, the results can be used to produce a

Load/Extension diagram or graph (refer to Fig. 1). Obviously a number of different forms of graph may be obtained, depending on the material type and condition, but the example shows a Load/Extension diagram which typifies many metallic materials when stressed in tension.

Load/Extension Diagram Fig 1

The graph can be considered as comprising two major regions. Between points 0 and A, the material is in the Elastic region (or phase), i.e. when the load is

removed the material will return to its original size and shape. In this region, the extension is directly proportional to the applied load.

This relationship is known as ‘Hooke's Law’, which states:

Within the elastic region, elastic strain is directly proportional to the stress causing it.

Point A is the Elastic Limit. Between this point and point B, the material continues to extend until the maximum load is reached (at point B). In this region the

material is in the plastic phase. When the load is removed, the material does not return to its original size and shape, but will retain some extension. After point B, the cross-sectional area reduces and the material begins to ‘neck’. The material continues to extend under reduced load until it eventually fractures at point C.

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Aircraft structural designers’ interest in materials does not extend greatly beyond the elastic phase of materials. Production engineers, however, are greatly

interested in material properties beyond this phase, since the forming capabilities of materials are dependent on their properties in the plastic phase.

An examination of a graph, obtained from the results of a tensile test on mild steel (refer to Fig. 2), shows that considerable plastic extension occurs without any increase in load shortly after the elastic limit is reached. The onset of increasing extension, without a corresponding increase in load, at point `B', is known as the ‘yield point’ and, if this level of stress is reached, the metal is said to have ‘yielded’. This is a characteristic of mild steel and a few other, relatively ductile, materials.

Load/Extension Diagram for Mild Steel Fig. 2

If, after passing the yield point, the load is further increased, it may be seen that mild steel is capable of withstanding this increase until the Ultimate Tensile Stress (UTS) is reached. Severe necking then occurs and the material will fracture at a reduced load. The unexpected ability of mild steel to accept more load after yielding is due to strain-hardening of the material. Work-hardening of many materials is often carried out to increase their strength.

Point B Yield Point UTS

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As previously stated, various forms of load/extension curves may be constructed for other materials (refer to Fig 3), and their slopes will depend on whether the materials are brittle, elastic or plastic.

Load/Extension Graphs for Brittle, Elastic and Plastic Materials Fig. 3

(a) represents a brittle material (glass)

(b) represents a material with some elasticity and limited plasticity (high-carbon steel

(c) represents a material with some elasticity and good plasticity (e.g. soft aluminium). Zero Elongation Small Elongation Large Elongation Plastic Region Point of Fracture (a) (b) (c)

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3.2.1 DUCTILITY

After fracture of a specimen, following tensile testing, an indication of material ductility is arrived at, by establishing the amount of plastic deformation which occurred. The two indicators of ductility are:

• Elongation

• Reduction in area (at the neck)

Elongation is the more reliable, because it is easier to measure the extension of the gauge length than the reduction in area. The standard measure of ductility is to establish the percentage elongation after fracture.

Example 4

In a tensile test, on a specimen with 150.5 mm gauge length, the length over the gauge marks at fracture were 176.1 mm. What was the percentage elongation?

3.2.2 PROOF STRESS

Many materials do not exhibit a yield point, so a substitute value must be employed. The value chosen is the ‘Proof Stress’, which is defined as:

The tensile stress, which is just sufficient to produce a

non-proportional elongation, equal to a specified percentage of the original gauge length.

Usually a value of 0.1% or 0.2% is used for Proof Stress, and the Proof Stress is then referred to as the 0.1% Proof Stress or the 0.2% Proof Stress respectively. The Proof Stress may be acquired from the relevant Load/Extension graph (refer to Fig 4) as follows:

If the 0.2% Proof Stress is required, then 0.2% of the gauge length is marked on the extension axis. A line, parallel to the straight-line portion of the graph, is drawn until it intersects the non-linear portion of the curve. The corresponding load is then read from the graph. Proof Stress is calculated by dividing this load by the original cross-sectional area.

100 Length Gauge Original Extension Final elongation Percentage × 17% 17.009% 100 5 150 150.5 -176.1 100 Length Gauge Extension Final Elongation × = = ⋅ = × =

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0.1% Proof Stress will produce permanent set equivalent to one thousandth of the specimen's original length.

0.2% Proof Stress will produce permanent set equivalent to one five hundredth of the original length.

Acquiring the Proof Stress from a Load/Extension Graph Fig. 4

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3.3 STIFFNESS

Within the elastic range of a material, if the Strain is compared to the Stress causing that extension, it will provide a measure of stiffness/rigidity or flexibility.

This value, which is of great importance to designers, is known as ‘the Modulus

of Elasticity, or Young’s Modulus’, and is signified by use of the symbol E.

Thus E = Stress divided by Strain and, since Strain has no units, the unit for `E' is the same as Stress. i.e. lbf/in2, tonf/in2 (also psi and tsi), or the European and SI units such as kN/m2, MN/m2 and kPa or MPa.

The actual numerical value is usually large, as it is a measure of the stress required to theoretically double the length of a specimen (if it did not break first). A typical value of E for steel would be 30 x 106 psi. or 210,000 MN/m2

Relative stiffness values for some common materials (using Rubber as a datum), are:

• Wood 2000 x

• Aluminium 10,000 x • Steel 30,000 x • Diamond 171,000 x

3.4 TENSILE TESTING OF PLASTICS

This is conducted in the same way as for metals, but the test piece is usually made from sheet material. Although the basic load/extension curve for some plastics is somewhat similar to metal curves, changes in test temperature or the rate of loading can have a major effect on the actual results.

Even though the material under test may be in the elastic range, the specimen may take some time to return to its original size after the load is removed.

stiffness of measure a is Strain Stress . ie

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3.5 COMPRESSION TEST

Machines for compression testing are often the same as those used for tensile testing, but the test specimen is in the form of a short cylinder.

The Load/Deflection graph in the elastic phase for ductile materials is similar to that in the tensile test. The value of `E' is the same in compression as it is in tension. Compression testing is seldom used as an acceptance test for metallic or plastic materials (except for cast iron).

Compression testing is generally restricted to building materials and research into the properties of new materials.

3.6 HARDNESS TESTING

The hardness of materials is found by measuring their resistance to indentation. Various methods are used, but the most common are those of the Brinell, Vickers and Rockwell Hardness Tests.

3.6.1 BRINELL TEST

In the Brinell Hardness Test (refer to Fig. 5), a hardened steel ball is forced into the surface of a prepared specimen, using a calibrated force, for a specified time. The diameter of the resulting indentation is then measured accurately, using a graduated microscope and, thus, the area of the indentation is calculated. The hardness number is determined by reference to a Brinell Hardness Number (BHN) chart.

The Brinell Hardness Test Fig. 5

Diameter (Area) of resulting Indentation

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3.6.2 VICKERS TEST

The Vickers Hardness Test is similar to the Brinell test but uses a square-based diamond pyramid indenter (refer to Fig. 6). The diagonals, of the indentation, are accurately measured, by a special microscope, and the Hardness Value (HV) is again determined by reference to a chart.

The Vickers Hardness Test Fig. 6

3.6.3 ROCKWELL TEST

The Rockwell Hardness Test (refer to Fig. 7) also uses indentation as its basis, but two types of indenter are used. A conical diamond indenter is employed for hard materials and a steel ball is used for soft materials. The hardness number, when using the steel ball, is referred to as Rockwell B (e.g. RB 80) and the diamond hardness number is known as Rockwell C (e.g. RC 65).

Note: Whereas Brinell and Vickers hardness values are based upon the area of indentation, the Rockwell values are based upon the depth of the indentation.

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No precise relationship exists between the various hardness numbers, but approximate relationships have been compiled. Some comparative values between Brinell Vickers and Rockwell are shown in Table 1.

Table 1

COMPARATIVE HARDNESS VALUES

MATERIAL BHN HV ROCKWELL

Aluminium alloy 100 100 B 57

Mild steel 130 130 B 73

Cutting tools 650 697 C 60

Note: There is a good correlation between hardness and U.T.S. on some materials (e.g. steels)

3.6.4 HARDNESS TESTING ON AIRCRAFT

It is not normal to use Brinell, Rockwell or Vickers testing methods on aircraft in the hangar. There are, however, portable Hardness Testers, which may be used to test for material hardness on items such as aircraft wheels, after an over-heat condition, because the over-heat condition may cause the wheel material to become soft or partially annealed.

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3.7 IMPACT TESTING

The impact test (refer to Fig. 8) is designed to determine the toughness of a material and the two most commonly used methods are those using the ‘Charpy’ and ‘Izod’ impact-testing machines.

Both tests use notched-bar test pieces of standard dimensions, which are struck by a fast-moving, weighted pendulum. The energy, which is absorbed by the test piece on impact, will give a measure of toughness. A brittle material will break easily and will absorb little energy, so the swing of the pendulum (which is recorded against a calibrated scale) will not be reduced significantly. A tough material will, however, absorb considerably more energy and thus greatly reduce the recorded pendulum swing.

Most materials show a drop in toughness with a reduction in temperature, though some materials (certain steels in particular) show a rapid drop in toughness as the temperature is progressively reduced. This temperature range is called the Transition Zone, and components, which are designed for use at low

temperature, should be operated above the material’s Transition Temperature. Nickel is one of the most effective alloying elements for lowering the Transition Temperature of steels

Impact Test Fig. 8 .

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3.8 OTHER FORMS OF MATERIAL TESTING

Although some of the more important forms of material testing have already been discussed, there are several other forms of material testing to be considered, not least important of which are those associated with Creep and Fatigue Testing.

3.8.1 CREEP

Creep can be defined as the continuing deformation, with the passage of time, of materials subjected to prolonged stress. This deformation is plastic and occurs even though the acting stress may be well below the yield stress of the material. At temperatures below 0.4T (where T is the melting point of the material in Kelvin), the creep rate is very low, but, at higher temperatures, it becomes more rapid. For this reason, creep is commonly regarded as being a high-temperature phenomenon, associated with super-heated steam plant and gas turbine

technology. However, some of the soft, low-melting point materials will creep significantly at, or a little above, ambient temperatures and some aircraft materials may creep when subjected to over-heat conditions.

3.8.2 CREEP IN METALS

When a metallic material is suitably stressed, it undergoes immediate elastic deformation. This is then followed by plastic strain, which occurs in three stages (refer to Fig. 9):

• Primary Creep - begins at a relatively rapid rate, but then decreases with time as strain-hardening sets in.

• Secondary Creep - the rate of strain is fairly uniform and at its lowest value. • Tertiary Creep - the rate of strain increases rapidly, finally leading to

rupture. This final stage coincides with gross necking of the component, prior to failure. The rate of creep is at a maximum in this phase.

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3.8.3 EFFECT OF STRESS AND TEMPERATURE ON CREEP

Both stress and temperature have an effect on creep. At low temperature or very low stress, primary creep may occur, but this falls to a negligible value in the secondary stage, due to strain-hardening of the material. At higher stress and/or temperature, however, the rate of secondary creep will increase and lead to tertiary creep and inevitable failure.

It is clear, from the foregoing, that short-time tensile tests do not give reliable information for the design of structures, which must carry static loads over long periods of time, at elevated temperatures. Strength data, determined from long- time creep tests (up to 10,000 hours), are therefore essential.

Although actual design data are based on the long-time tests, short-time creep tests are sometimes used as acceptance tests.

Three Stages of Creep Fig. 9

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3.8.4 THE EFFECT OF GRAIN SIZE ON CREEP

Since the creep mechanism is partly due to microscopic flow along the grain boundaries, creep resistance is improved by increased grain size, due to the reduced grain boundary region per unit volume. It is mainly for this reason that some modern, high-performance turbine blades are being made from directionally solidified (and, alternatively, improved single-crystal) castings.

3.8.5 CREEP IN PLASTICS

Plastics are also affected by creep and show similar, though not identical,

behaviour to that described for metals. Since most plastics possess lower thermal properties than metals, the choice of plastic for important applications, particularly at elevated temperature, must take creep considerations into account.

3.8.6 FATIGUE

An in-depth survey, in recent years, revealed that over 80 percent of failures of engineering components were caused by fatigue. Consequences of modern engineering have led to increases in operating stresses, temperatures and speeds. This is particularly so in aerospace and, in many instances, has made the fatigue characteristics of materials more significant than their ordinary, static strength properties.

Engineers became aware that alternating stresses, of quite small amplitude, could cause failure in components, which were capable of safely carrying much greater, steady loads. This phenomenon of small, alternating loads causing failure was likened to a progressive weakening of the material, over a period of time and hence the attribution of the term ‘fatigue’. Very few constructional members are immune from it, and especially those operating in a dynamic environment.

Experience in the aircraft industry has shown that the stress cycles, to which aircraft are subjected, may be very complex, with occasional high peaks, due to gust loading of aircraft wings. For satisfactory correlation with in-service

behaviour, full-size or large-scale mock-ups must be tested in conditions as close as possible to those existing in service.

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3.8.7 FATIGUE TESTING

An experiment, conducted in 1861, found that a wrought iron girder, which could safely sustain a mass of 12 tons, broke when a mass of only 3 tons was raised and lowered on the girder some 3x106 times.

It was also found that there was some mass, below 3 tons, which could be raised and lowered on to the beam, a colossal number (infinite) of times, without causing any problem.

Some years later, a German engineer (Wohler), did work in this direction and eventually developed a useful fatigue-testing machine which bears his name and continues to be used in industry. The machine uses a test piece, which is rotated in a chuck and a force is applied at the free end, at right angles to the axis of rotation (refer to Fig. 10). The rotation thus produces a reversal of stress for every revolution of the test piece.

Various other types of fatigue testing are also used e.g. cyclic-torsional, tension-compression etc. Exhaustive fatigue testing, with various materials, has resulted in a better understanding of the fatigue phenomenon and its implications from an engineering viewpoint.

Test Piece made to vibrate or oscillate against load (Stress Cycles).

Test Piece

Load

Simple Fatigue Testing Fig 10

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3.9 S-N CURVES

One of the most useful end-products, from fatigue testing, is an S-N curve, which shows, graphically, the relationship between the amount of stress (S), applied to a material, and the number of stress cycles (N), which can be tolerated before failure of the material.

Using a typical S-N curve, for a steel material (refer to Fig. 11), it can be seen that, if the stress is reduced, the steel will endure a greater number of stress cycles. The graph also shows that a point is eventually reached where the curve becomes virtually horizontal, thus indicating that the material will endure an infinite number of cycles at a particular stress level.

This limiting stress is called the ‘Fatigue Limit’ and, for steels, the fatigue limit is generally in the region of 40% to 60% of the value of the static, ultimate tensile strength (U.T.S.)

A S-N Curve for a Steel Material Fig. 11 Fatigue Limit 40 – 60 % UTS Number of Cycles (N) Stress (S)

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Many non-ferrous metals, however, show a different characteristic from steel (refer to Fig. 12). In this instance there is no fatigue limit as such and it can be seen that these materials will fail if subjected to an appropriate number of stress reversals, even at very small stresses. When materials have no fatigue limit an endurance limit together with a corresponding number of cycles is quoted instead.

It follows that components made from such materials must be designed with a specific life in mind and removed from service at the appropriate time. The

service fatigue lives of complete airframes or airframe members are typical examples of this philosophy.

Non-metallic materials are also liable to failure by fatigue. As is the case with metals, the number of stress cycles, required to produce a fatigue failure, will increase as the maximum stress in the loading cycle decreases. There is, however, generally no fatigue limit for these materials and some form of endurance limit must be applied.

The importance of fatigue strength can be illustrated by the fact that, in a high- cycle fatigue mode, a mere 10% improvement in fatigue strength can result in a 100-times life improvement.

An S-N Curve for an Aluminium Alloy Fig. 12

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3.10 CAUSES OF FATIGUE FAILURE

As the fatigue characteristics of most materials are now known (or can be ascertained), it would seem reasonable to suppose that fatigue failure, due to lack of suitable allowances in design, should not occur.

Nevertheless, fatigue cracking occurs frequently, and even the most

sophisticated engineering product does not possess immunity from this mode of failure. Such failures are often due to unforeseen factors in design, environmental or operating conditions, material, and manufacturing processes.

Two essential requirements for fatigue development in a material are: • An applied stress fluctuation of sufficient magnitude (with or without an

applied steady stress).

• A sufficient number of cycles of that fluctuating stress.

The stress fluctuations may be separated by considerable time intervals, as experienced in aircraft cabin pressurisation, during each take-off (e.g. daily), or they may have a relatively short time interval, such as encountered during the aerodynamic buffeting/vibration of a wing panel. The former example would be considered to be low-cycle fatigue and the latter to be high-cycle fatigue.

In practice, the level of the fluctuating stress, and the number of cycles to cause cracking of a given material, are affected by many other variables, such as stress concentration points (stress raisers), residual internal stresses, corrosion, surface finish, material imperfections etc.

3.11 VIBRATION

Vibration has already been quoted as being a cause of high-cycle fatigue and, because most dynamic structures are subjected to vibration, this is undoubtedly the most common origin of fatigue. All objects have their own natural frequency at which they will freely vibrate (the resonant frequency). Large, heavy, flexible components vibrate at a low frequency, while small, light, stiff components vibrate at a high frequency.

Resonant frequencies are undesirable (and in some cases could be disastrous), so it is important to ensure that, over their normal operating ranges, critical components are not vibrated at their natural frequencies and so avoid creating resonance.

The resonant frequency, of a component, is governed by its mass and stiffness and, on certain critical parts, it is often necessary to do full-scale fatigue tests to confirm adequate fatigue life before putting the product into service.

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3.12 FATIGUE METALLURGY

Under the action of fatigue stresses, minute, local, plastic deformation on an atomic scale, takes place along slip planes within the material grains. If the fatigue stresses are continued, then micro cracks are formed within the grains, in the area of the highest local stress, (usually at or near the surface of the

material). The micro cracks join together and propagate across the grain boundaries but not along them.

A fatigue fracture generally develops in three stages (refer to Fig. 13): • Nucleation

• Propagation (crack growth) • Ultimate (rapid) fracture.

The resultant fractured surface often has a characteristic appearance of:

• An area, on which a series of curved, parallel, relatively smooth ridges are present and are centred around the starting point of the crack. These ridges are sometimes called conchoidal lines or beach marks or arrest lines.

• A rougher, typically crystalline section, which is the final rapid fracture when the cross-section is no longer capable of carrying its normal, steady load.

Nucleation Propagation (crack growth) Ultimate (rapid) fracture The Three Stages of Fracture

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The arrest lines are, normally, formed when the loading is changed, or the loading is intermittent. However, in addition to these characteristic and informative marks, there are similar, but much finer lines (called ‘striations’), which literally show the position of the crack front after each cycle. These striations are obviously of great importance to metallurgists and failure

investigators when attempting to estimate the crack initiation and/or propagation life. The striations are often so fine and indistinct that electron beam microscopes are required to count them.

In normal circumstances, a great deal of energy is required to `weaken' the material sufficiently to initiate a fatigue crack, and it is not surprising, therefore, to find that the nucleation phase takes a relatively long time.

However, once the initial crack is formed, the extremely high stress concentration (present at the crack front) is sufficient to cause the crack to propagate relatively quickly, and gaining in speed as the crack front not only increases in size, but also reduces the component cross-sectional area.

A point is eventually reached (known as the 'critical crack length') at which the remaining cross-section is sufficiently reduced to cause a gross overloading situation, and a sudden fracture finally occurs.

It is not unusual for the crack initiation phase to take 90% of the time to failure, with the propagation phase only taking the remaining 10%. This is one of the major reasons for operators of equipment being relatively unsuccessful in detecting fatigue cracks in components before a failure occurs.

3.13 FATIGUE PROMOTERS

As fatigue cracks initiate at locations of highest stress and lowest local strength, the nucleation site will be:

• dictated largely by geometry and the general stress distribution • located at or near the surface or

• centred on surface defects/imperfections, such as scratches, pits, inclusions, dislocations and the like

3.13.1 DESIGN

Apart from general stressing, the geometry of a component has a considerable influence on its susceptibility to fatigue. A good designer will therefore minimise stress concentrations by:

• avoiding rapid changes in section and

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3.13.2 MANUFACTURE

While the designer may specify adequate blend radii, the actual product may still be prone to fatigue failure if the manufacturing stage fails to achieve this

sometimes-seemingly unimportant drawing requirement.

Several other manufacturing-related causes of premature fatigue failure exist, the most common of which are:

• Inherent material faults: e.g. cold shuts, pipe, porosity, slag inclusions etc. • Processing faults: e.g. bending, forging, grinding, shrinking, welding, etc. • Production faults: e.g. incorrect heat-treatment, inadequate surface

protection, poor drilling procedures, undue force used during assembly, etc • In-service damage: e.g. dents, impact marks, scratches, scores, tooling

marks etc.

3.13.3 ENVIRONMENT

One of the most potent environmental promoters of fatigue occurs when the component is operating in a corrosive medium. Steel (normally), has a well-defined fatigue limit on the S-N curve but, if a fatigue test is conducted in a corrosive environment, not only does the general fatigue strength drop

appreciably, but the curve also resembles the aluminium alloy curve (e.g. the fatigue failure stress continues to fall as the number of cycles increases). Other environmental effects such as fretting and corrosion pitting, erosion or elevated temperatures will also adversely affect fatigue strength.

3.14 FATIGUE PREVENTERS

If a component is prone to fatigue failure in service, then several methods of improvement are available, in the form of:

• Quality. Correct and eliminate any failure-related manufacturing or processing shortcomings.

• Material. Select a material with a significantly better fatigue strength, or corrosion-resistance or corrosion-protection if relevant.

• Geometry.

a) Increase the size (c.s.a.) to reduce the general stress level or modify the local geometry to reduce the change in section (large radius).

b) Modify the geometry to change the vibration frequency or

introduce a damping feature, to reduce the vibration amplitudes. c) Improve the surface finish or put a compressive stress in the skin

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3.14.1 COLD EXPANSION (BROACHING)

Most fatigue failures occur whilst a material is subject to a tensile, alternating stress. If the most fatigue-prone areas, such as spar fastener holes, have a compression stress applied (refer to Fig. 14), they are significantly more resistant to fatigue failure.

The fastener hole is initially checked for defects (using, usually, an Eddy Current NDT procedure) and the surface finish is further improved by reaming (and checked once again).

A tapered mandrel is then pulled through the hole, resulting in a localised area of residual (compressive) stress which will provide a neutral or, at least, a significantly reduced level of fatigue in the area around the fastener hole

Cold Expansion of Fastener Hole Fig.14

Area around hole pre-stressed in compression

Tapered Mandrel pulled through fastener hole

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3.15 DO'S AND DONT'S – PREVENTING FATIGUE FAILURES DO

• Be careful not to damage the surface finish of a component by mishandling. • Use the right tools for assembling press-fit components etc.

• Maintain drawing sizes and tolerances.

• Keep the correct procedures (e.g. don't overheat when welding).

• Avoid contact or near contact of components that might cause fretting when touching.

DON'T

• Leave off protective coverings - plastic end caps etc. • Score the surface.

• Leave sharp corners or ragged holes. • Force parts unnecessarily to make them fit.

• Work metal unless it is in the correct heat-treated state.

3.16 STRUCTURAL HEALTH MONITORING (SHM)

Obviously it is extremely important, that the level of fatigue, imposed on an aircraft structure (and associated components), be monitored and recorded so that the respective fatigue lives are not exceeded. Several methods have been developed to assist in the vital tasks involved with SHM

3.16.1 FATIGUE METERS

Fatigue meters are used to check overall stress levels on aircraft and to monitor the fatigue history of the aircraft. Fatigue meters also allow a check to be made on the moment in time when the aircraft exceeds the design limits imposed on it.

3.16.2 STRAIN GAUGES

Strain gauges may be used to monitor stress levels on specific aircraft structures. Strain gauges are thin-foil, electrical, resistor elements, bonded to the aircraft structure. Their resistance varies proportional to the applied stress loading.

3.16.3 FATIGUE FUSES

Fatigue fuses are metallic fuses, which are bonded to the structure and which fail at different fatigue stresses. The electrical current, flowing through the fuse, will vary and thus, provide an indication of the stress level.

3.16.4 INTELLIGENT SKINS DEVELOPMENT

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structure and skin. This would mainly be restricted to structures manufactured from composite materials. One major benefit of this is to allow the structure to monitor it's own loads and fatigue life.

3.16.4.1 Smart Structures

The generic heading ‘Smart Structures’ actually covers three areas of development:

• Smart Structures. These are structures, which have sensors, actuators, signal-processing and adaptive control systems built in

• Smart Skins. These have radar and communications antennae embedded in, or beneath, the structural skin

• Intelligent Skins. Skin embedded with fibre optic sensors Smart Structures perceived benefits include:

• Self-diagnostic in the monitoring of structural integrity • Reduced life cycle costs

• Reduced inspection costs

• Potential weight saving/performance improvements derived from increased knowledge of composite material characteristics

• From a military point of view – an improvement in ‘Stealth’ characteristics. A fully monitored and self-diagnostic system could:

• Assess structural integrity. • Pinpoint structural damage. • Process flight history.

Composite laminates, containing embedded fibre optic sensors can be used for SHM, including fatigue monitoring and flight envelope exceedance monitoring and their advantages include:

• Cover a greater area of structure •

• Not prone to electrical interference

• Less vulnerable to damage when embedded in the plies Increased knowledge of structural loads aids designers

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4 AIRCRAFT MATERIALS - FERROUS

Any alloy containing iron as its main constituent is called a ferrous metal. The most common ferrous metal, in aircraft construction is steel, which is an alloy of iron with a controlled amount of carbon added.

4.1 IRON

Iron is one of the most common elements in the Earth's crust. It comprises approximately 5% compared with aluminium at 8%. Iron is never found naturally in its metallic state, but as iron ores which only contain in the range of 25% to 60% iron and are mined in open-cast or open-pit mines. Iron has a great affinity for oxygen.

Iron is a chemical element that is fairly soft, malleable and ductile in its pure form. It is silvery-white in colour and quite heavy, having a density of 7870 kgm-3. Unfortunately it combines well with oxygen, producing iron oxide, which is more commonly known as rust. Iron usually has other materials added to improve its properties.

The first smelt from the raw ore is poured into troughs (which are said to resemble piglets suckling on a sow) and the iron is referred to as ‘pig iron’. The pig iron is then re-melted to give cast irons.

4.1.1 CAST IRON

Cast Iron normally contains over two percent carbon and some silicon. It has few aircraft applications, excepting where its hardness and porosity are required, such as in piston rings and valve guides.

4.1.2 NODULAR CAST IRON

This is a more modern development and is sometimes known as ‘Spheroidal Graphite Iron’. It is produced by adding magnesium and nickel (or magnesium, copper and silicon) and is a tough, strong, hard-wearing material which can be used in applications where only wrought materials were used in the past (a classic example being piston engine crankshafts).

4.2 STEEL

Steel is essentially an alloy of iron and less than 2.5% carbon, usually with a few impurities. (In practice most steels do not have more than 1.5% carbon).

Steel is produced by refining pig iron (removing excess carbon and other unwanted impurities). The excess carbon is extracted by blowing oxygen or air through the molten metal, and/or adding iron oxide. Slag, containing other impurities, is skimmed off. The most common furnace used for this process was the ‘Bessemer Converter’, developed in 1856. It reduced the cost of steel to one

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fifth of its original cost. Bessemer converters were loaded with 20 - 50 tons of pig iron and air was blown from the bottom for approximately 15 minutes.

The high quality steels, used in aircraft construction, are usually produced in electric furnaces, which allow better control, than do gas furnaces, when alloying. The carbon electrodes produce an intense arc and the steel, when molten, can have impurities removed and measured amounts of alloying materials added.

4.2.1 CLASSIFICATION OF STEELS

When carbon is alloyed with iron, the hardness and strength of the metal

increases. The effect of varying amounts of carbon is truly dramatic. If carbon is progressively added to pure iron the following occurs:

• Initially, the strength and hardness increases - (Steel containing 0.4% carbon has twice the strength of pure iron.

• When 1% of carbon is added, the strength and hardness show a further increase but ductility is reduced.

• If 1% to 1.5% of carbon is added, the hardness continues to increase, but there is no further increase in strength and there is even less ductility. Steels containing such high amounts of carbon are seldom used for anything except cutting implements e.g. razor blades and scissors

The (American) Society of Automotive Engineers (SAE) has classified steel alloys with a four-digit numerical index system. A small extract from the SAE

classification system is shown in Table 2, where it can be seen, for example, that one common steel alloy is identified by the designation SAE 1030. The first digit identifies it as a Carbon-Steel, while the second digit shows that it is a Plain Carbon-Steel. The last two digits denote the percentage of carbon in the steel (0.30%).

It should be noted that the British Standards Institute (BS) has a different classification system.

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Table 2

EXTRACT FROM THE SAE CLASSIFICATION FOR STEEL ALLOYS

1xxx Carbon Steels

10xx Plain Carbon Steels

2xxx Nickel Steels

3xxx Nickel Chromium Steels

40xx Molybdenum Steels

41xx Chromium Molybdenum Steels

5xxx Chromium Steels

6xxx Chromium Vanadium

4.2.2 METALLURGICAL STRUCTURE OF STEEL

The amount of carbon present in steel has a major effect on the mechanical properties. The form in which the carbon is present is also important.

4.2.3 STRUCTURE AND PROPERTIES – SLOW-COOLED STEELS

Carbon can be present in these steels in the following forms:

• When the carbon is fully dissolved and, consequently, uniformly distributed in a solid solution, the metallurgical structure is called ferrite. At room temperature only a very small amount of carbon (0.006%) can be contained in solid solution, therefore this ferrite structure is almost pure iron. It is (not surprisingly) soft, weak and ductile.

• When 1 carbon atom chemically combines with 3 iron atoms the result is called cementite or iron carbide. It is very hard and brittle.

• Cementite can be present either as free cementite or laminated with ferrite (in alternate layers) to produce a metallurgical structure called pearlite. As pearlite is half cementite and half ferrite, it is not surprising to find that pearlite combines the properties of ferrite and cementite I.e. Whereas ferrite was too soft and weak - and cementite was basically strong but too hard and brittle - pearlite is strong without being brittle.

The amount of carbon necessary to produce a totally pearlite structure is 0.83% but this material is a little too hard for general structural use. If the carbon content exceeds this value, the excess carbon forms carbon-rich cementite areas along the grain boundaries, and this is known as free cementite. Such high-carbon steels as already stated are very hard and strong but very brittle.

Mild steel has a metallurgical structure comprising approximately one third pearlite and two thirds ferrite.

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4.2.4 EFFECTS OF COOLING RATES ON STEELS

Previously the effect of carbon on the properties of a slowly cooled steel has been considered. If such steels are, however, rapidly cooled from relatively high temperature the metallurgical structure and properties can be somewhat different.

4.3 HEAT-TREATMENT OF CARBON STEELS

If a ‘straight’ carbon steel is progressively heated from cold, a steady rise in temperature occurs. However, at approximately 700°C, there is a reduction in the rate of temperature rise (a ‘hesitation’), even though the heating is continued (refer to Fig. 15). This hesitation starts at 700°C and finishes at up to 200°C higher (depending on the percentage of carbon present) and, eventually, the temperature rise speeds up and the rate of rise is similar to that which occurred before the hesitation.

The start of the hesitation is known as the ‘lower critical point’ and the end is called the ‘upper critical point’, and the phenomenon of the temperature

response is due to a change in the crystalline structure of the steel in between the two critical points.

Temperature/Time Graph for Steel Heat-Treatments Fig. 15

EASA Part 66 - Module 6 - Materials and Hardware

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CONTENTS

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1 INTRODUCTION

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2 PROPERTIES OF METALS

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3 TESTING OF MATERIALS

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