EASA Part 66 - Module 6 - Materials and Hardware

uk engineering

JAR 66 CATEGORY B1 MODULE 6 MATERIALS AND HARDWARE

CONTENTS 1 2

INTRODUCTION ......................................................................................... 1-1 PROPERTIES OF METALS ........................................................................ 2-1 2.1 2.2 2.3 2.4 2.5 2.6 2.7 2.8 2.9 2.10

3

BRITTLENESS .......................................................................................... 2-1 CONDUCTIVITY ........................................................................................ 2-1 DUCTILITY ................................................................................................ 2-1 ELASTICITY .............................................................................................. 2-1 HARDNESS ............................................................................................... 2-1 MALLEABILITY.......................................................................................... 2-1 PLASTICITY .............................................................................................. 2-1 TENACITY ................................................................................................. 2-1 TOUGHNESS ............................................................................................ 2-2 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

TESTING OF MATERIALS ......................................................................... 3-1 3.1 3.2

3.3 3.4 3.5 3.6

3.7 3.8

3.9 3.10 3.11

TENSILE TESTING ................................................................................... 3-1 3.1.1 Tensile Strength ........................................................................... 3-1 LOAD/EXTENSION DIAGRAMS ............................................................... 3-4 3.2.1 Ductility ........................................................................................ 3-7 3.2.2 Proof Stress ................................................................................. 3-7 STIFFNESS ............................................................................................... 3-9 TENSILE TESTING OF PLASTICS ........................................................... 3-9 COMPRESSION TEST ............................................................................ 3-10 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 IMPACT TESTING................................................................................... 3-13 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 S-N CURVES .......................................................................................... 3-18 CAUSES OF FATIGUE FAILURE ............................................................ 3-20 VIBRATION ............................................................................................. 3-20

EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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3.14 3.15 3.16

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FATIGUE METALLURGY ........................................................................ 3-21 FATIGUE PROMOTERS ......................................................................... 3-22 3.13.1 Design........................................................................................ 3-22 3.13.2 Manufacture ............................................................................... 3-23 3.13.3 Environment ............................................................................... 3-23 FATIGUE PREVENTERS ........................................................................ 3-23 3.14.1 Cold Expansion (Broaching) ....................................................... 3-24 DO'S AND DONT'S – PREVENTING FATIGUE FAILURES .................... 3-25 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

AIRCRAFT MATERIALS - FERROUS ........................................................ 4-1 4.1

4.2

4.3

4.4

4.5 4.6

4.7 4.8 4.9 4.10 4.11 4.12 4.13

IRON ......................................................................................................... 4-1 4.1.1 Cast Iron ...................................................................................... 4-1 4.1.2 Nodular Cast Iron ......................................................................... 4-1 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 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 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 ALLOYING ELEMENTS IN STEEL ............................................................ 4-9 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 SULPHUR ................................................................................................. 4-9 SILICON .................................................................................................... 4-9 PHOSPHORUS ....................................................................................... 4-10 NICKEL ................................................................................................... 4-10 4.10.1 Nickel Alloys............................................................................... 4-10 CHROMIUM (CHROME) ......................................................................... 4-11 4.11.1 Nickel-Chrome Steel and its Alloys ............................................ 4-11 COBALT .................................................................................................. 4-11 VANADIUM.............................................................................................. 4-12


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MANGANESE .......................................................................................... 4-12 MOLYBDENUM ....................................................................................... 4-12 CHROME AND MOLYBDENUM .............................................................. 4-12 TUNGSTEN ............................................................................................. 4-13 MARAGING STEELS............................................................................... 4-13

AIRCRAFT MATERIALS - NON-FERROUS ............................................... 5-1 5.1

5.2 5.3 5.4 5.5

5.6 5.7 5.8 5.9 5.10 5.11 5.12

6


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 ALUMINIUM ALLOYS ................................................................................ 5-3 IDENTIFICATION OF ELEMENTS IN ALUMINIUM ALLOYS .................... 5-3 CLAD MATERIALS .................................................................................... 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 IDENTIFICATION OF HEAT-TREATED ALUMINIUM ALLOYS ................. 5-9 MARKING OF ALUMINIUM ALLOY SHEETS.......................................... 5-10 CAST ALUMINIUM ALLOYS ................................................................... 5-11 MAGNESIUM ALLOYS ............................................................................ 5-11 COPPER ALLOYS................................................................................... 5-12 TITANIUM ALLOYS ................................................................................. 5-13 WORKING WITH TITANIUM AND TITANIUM ALLOYS........................... 5-13 5.12.1 Drilling Titanium ......................................................................... 5-14

METHODS USED IN SHAPING METALS .................................................. 6-1 6.1

6.2

6.3 6.4 6.5 6.6 6.7

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 FORGING .................................................................................................. 6-5 6.2.1 Drop-Stamping ............................................................................. 6-6 6.2.2 Hot-Pressing ................................................................................ 6-6 6.2.3 Upsetting ...................................................................................... 6-6 ROLLING ................................................................................................... 6-7 DRAWING ................................................................................................. 6-7 DEEP DRAWING/PRESSING ................................................................... 6-7 PRESSING ................................................................................................ 6-7 STRETCH-FORMING ................................................................................ 6-7


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7

7.2 7.3 7.4 7.5

7.6

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 PRIMARY ADVANTAGES OF PLASTICS ................................................. 7-7 PRIMARY DISADVANTAGES OF PLASTICS ........................................... 7-7 PLASTIC MANUFACTURING PROCESSES ............................................. 7-8 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 NON-METALLIC COMPONENTS ............................................................ 7-13 7.6.1 Seals .......................................................................................... 7-13

DETECTING DEFECTS IN COMPOSITE MATERIALS.............................. 8-1 8.1 8.2 8.3

8.4

9

RUBBER-PAD FORMING ......................................................................... 6-7 EXTRUDING ............................................................................................. 6-8 6.9.1 Impact-Extrusion .......................................................................... 6-8 SINTERING ............................................................................................... 6-8 SPINNING ................................................................................................. 6-9 CHEMICAL MILLING ................................................................................. 6-9 ELECTRO-CHEMICAL MACHINING ......................................................... 6-9 ELECTRO-DISCHARGE MACHINING E.D.M. ........................................ 6-10 CONVENTIONAL MACHINING ............................................................... 6-11 SUPERPLASTIC FORMING .................................................................... 6-12

AIRCRAFT MATERIALS - COMPOSITE AND NON-METALLIC ................ 7-1 7.1

8


CAUSES OF DAMAGE.............................................................................. 8-1 TYPES OF DAMAGE................................................................................. 8-1 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 ASSESSMENT OF DAMAGE .................................................................... 8-4

BASIC COMPOSITE REPAIRS .................................................................. 9-1 9.1 9.2 9.3 9.4

REPAIR OF A SIMPLE COMPOSITE PANEL ........................................... 9-2 REPAIR OF A SANDWICH PANEL ........................................................... 9-3 GLASS FIBRE REINFORCED COMPOSITE REPAIRS ............................ 9-5 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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9.9 9.10 9.11


POT LIFE................................................................................................... 9-7 CURING .................................................................................................... 9-7 GEL COAT ................................................................................................ 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 PREPARATION FOR REPAIR .................................................................. 9-9 9.9.1 Surface Preparation ................................................................... 9-11 TECHNIQUES OF LAMINATING GLASS FIBRE..................................... 9-11 PRE-WETTING GLASS FIBRE ............................................................... 9-12

10 ADHESIVES AND SEALANTS ................................................................. 10-1 10.1

10.2

10.3

10.4

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 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 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 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

11.2

11.3

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 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 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 11.3.9 Pitting Corrosion .......................................................................11-12


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11.4

11.5

11.6

11.7

11.8


11.3.10 Corrosion Fatigue......................................................................11-13 11.3.11 Microbiological Contamination...................................................11-13 11.3.12 Hydrogen Embrittlement of Steels .............................................11-13 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 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 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 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 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 12.2 12.3 12.4 12.5

TEMPORARY JOINTS ............................................................................ 12-1 PERMANENT JOINTS............................................................................. 12-1 FLEXIBLE JOINTS .................................................................................. 12-1 SCREW THREADS ................................................................................. 12-2 12.4.1 The Inclined Plane and the Helix ................................................ 12-2 SCREW THREAD TERMINOLOGY......................................................... 12-4


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12.6

12.7 12.8

12.9

12.10

12.11

12.12

12.13

12.14


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 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 NUTS......................................................................................................12-18 12.7.1 Stiffnuts and Anchor Nuts..........................................................12-19 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 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 THREAD INSERTS ................................................................................12-27 12.10.1 Wire Thread Inserts...................................................................12-27 12.10.2 Thin Wall Inserts .......................................................................12-28 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 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 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 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 12.14.7 Keys and Keyways ....................................................................12-43


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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 13.2

13.3

13.4

13.5

13.6

SOLID RIVETS ........................................................................................ 13-1 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 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 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 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 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 14.2

FORCES EXERTED ON, AND APPLIED BY, SPRINGS ......................... 14-1 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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14.4 14.5


14.2.6 Belleville (Coned Disc) Springs .................................................. 14-2 14.2.7 Torsion-Bar Springs ................................................................... 14-2 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 CHARACTERISTICS OF TYPICAL AEROSPACE SPRINGS.................. 14-5 APPLICATIONS OF SPRINGS IN AIRCRAFT ENGINEERING ............... 14-6

15 PIPES AND UNIONS ................................................................................ 15-1 15.1 15.2

15.3

15.4

15.5

RIGID PIPES ........................................................................................... 15-1 SEMI-RIGID FLUID LINES (TUBES) ....................................................... 15-2 15.2.1 Flared End-Fittings ..................................................................... 15-2 15.2.2 Flare-Less Couplings ................................................................. 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 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 QUICK-RELEASE COUPLINGS .............................................................. 15-8

16 BEARINGS................................................................................................ 16-1 16.1

16.2

16.3

16.4

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 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 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 BEARING MAINTENANCE ...................................................................... 16-5 16.4.1 Lubrication ................................................................................. 16-5 16.4.2 Inspection .................................................................................. 16-5

17 TRANSMISSIONS .................................................................................... 17-1 17.1 17.2

BELTS AND PULLEYS ............................................................................ 17-1 GEARS .................................................................................................... 17-3 17.2.1 Gear Trains and Gear Ratios ..................................................... 17-3 17.2.2 Spur Gears................................................................................. 17-4


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17.3 17.4


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 CHAINS AND SPROCKETS .................................................................... 17-8 17.3.1 Typical Arrangements - Chain Assemblies ................................. 17-9 MAINTENANCE INSPECTIONS.............................................................17-10

18 CONTROL CABLES.................................................................................. 18-1 18.1 18.2

18.3

TYPES OF CABLES ................................................................................ 18-1 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 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 19.2 19.3 19.4 19.5 19.6 19.7 19.8 19.9 19.10 19.11 19.12 19.13 19.14 19.15 19.16 19.17 19.18

CABLE SPECIFICATION......................................................................... 19-1 CABLE IDENTIFICATION ........................................................................ 19-1 DATA BUS CABLE .................................................................................. 19-5 CONDUCTOR MATERIAL & INSULATION ............................................. 19-6 WIRE SIZE .............................................................................................. 19-7 WIRE RESISTANCE ............................................................................... 19-8 CURRENT CARRYING CAPABILITY ...................................................... 19-8 VOLTAGE DROP ...................................................................................19-10 WIRE IDENTIFICATION .........................................................................19-11 WIRE INSTALLATION AND ROUTING ..................................................19-12 OPEN WIRING .......................................................................................19-12 WIRE & CABLE CLAMPING ...................................................................19-13 CONDUIT ...............................................................................................19-14 CONNECTORS ......................................................................................19-16 CRIMPING..............................................................................................19-19 CRIMPING TOOLS.................................................................................19-20 WIRE SPLICING.....................................................................................19-21 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 heattreatment. 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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INTENTIONALLY BLANK


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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

The ability of a metal to withstand a crushing force.


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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 crosssectional area (c.s.a.) of the specimen. The value derived from this simple calculation is called STRESS.

Stress =

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

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.


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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. Stress =

Load Area

400 400 = = 20 ⋅ 37 N / mm 2 2 2 Πr Π× 2⋅5

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)

Stress =

Load Area

1N/mm 2 = 1 MN/m 2

10 = 20MN / m 2 0⋅5 So Stress = 20 N/mm 2


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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.

Strain =

Extension Original Length

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. Extension = 21⋅ 15 - 20 = 1 ⋅ 15 mm Strain =

Extension 1 ⋅ 15 = = 0 ⋅ 0575 (no units) Original Length 20

Exercise 3 A tie rod 1.5m long under a tensile load of 500 N extends by 12 mm. Calculate the 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. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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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.

UTS

Point B Yield Point

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.


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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.

Point of Fracture

Plastic Region

Small Elongation

Zero Elongation

Large Elongation (a)

(b)

(c)

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 (highcarbon steel (c) represents a material with some elasticity and good plasticity (e.g. soft aluminium).


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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. Percentage elongation

Final Extension × 100 Original Gauge Length

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?

Elongation =

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

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 nonproportional 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.


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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. ie .

Stress is a measure of stiffness Strain

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 È' 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.


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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 È' 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.

Diameter (Area) of resulting Indentation

The Brinell Hardness Test Fig. 5


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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.

The Rockwell Hardness Test Fig. 7 EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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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 Aluminium alloy 100 Mild steel 130 Cutting tools 650

HV 100 130 697

ROCKWELL B 57 B 73 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

Test Piece 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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Three Stages of Creep Fig. 9 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 longtime 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.


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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, tensioncompression 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.) Stress (S)

Fatigue Limit

40 – 60 % UTS

Number of Cycles (N) A S-N Curve for a Steel Material Fig. 11


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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.

An S-N Curve for an Aluminium Alloy Fig. 12 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 highcycle fatigue mode, a mere 10% improvement in fatigue strength can result in a 100-times life improvement.


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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.

Nucleation

Propagation (crack growth)

Ultimate (rapid) fracture

The Three Stages of Fracture Fig. 13 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.


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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 • using generous blend radii or chamfers to eliminate sharp corners EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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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 welldefined 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 (e.g. shot peen or cold expand).


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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

Area around hole pre-stressed in compression

Tapered Mandrel pulled through fastener hole Cold Expansion of Fastener Hole Fig.14


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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

Modern developments in aircraft structures will allow the structures to be designed and built with a variety of sensors and systems embedded into the EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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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 EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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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.

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

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.


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If carbon steel is heated to just above its Upper Critical Point the structure is called ‘Austenitic’. This structure is a solid solution of carbon in iron (i.e. all the carbon is uniformly distributed throughout the iron). If the steel contains above 0.3% carbon, and it is rapidly cooled (i.e. quenched) from above the Upper Critical Point it becomes hardened. The more carbon present, the harder the steel will be after quenching. This rapid cooling causes a change in the metallurgical structure and is called ‘Martensite’. Martensite is extremely hard but is not suitable for most engineering purposes due to it being very brittle. For most applications it is necessary to carry out a further heat-treatment to reduce the brittleness of the steel, and this is called ‘tempering’. To temper hardened carbon steel it is necessary to heat it to a suitable temperature below its Lower Critical Point followed by cooling (usually quenching). The effect of this heat-treatment is to slightly reduce the hardness whilst at the same time greatly increasing the toughness. The actual tempering temperature used depends on the requirements of strength, hardness and toughness. The higher the tempering temperature, the lower will be strength and hardness, but the toughness will be greater. The maximum tensile strength of hardened carbon steel is achievable when 0.83% carbon is present. If an even greater amount of carbon is present, the hardness continues to increase but strength will decrease. 4.3.1 ASSOCIATED PROBLEMS - HARDENING PROCESS

The effective hardening of carbon steels depends not only on the amount of carbon present but also on very rapid cooling from high temperature. The cooling rate mainly depends on the cooling medium, the size of tank, and the mass of the object to be cooled. Agitation in the cooling bath can also speed up the cooling rate and, in terms of cooling severity, brine is more effective than water, followed by oil and finally air. Carbon steels require an extremely rapid cooling phase, so brine or water is normally used, whereas oil or air-cooling is used on certain alloy steels. The rapid cooling rates, involved in the hardening of carbon steel, cause enormous thermal stresses in the component and distortion is commonplace. Cracking may also occur in some cases. To achieve relatively uniform cooling it is sometimes necessary to immerse the object in a specific way because of its shape and mass.


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4.3.2 TEMPERING

As already stated, tempering is carried out to improve the toughness of hardened steel whilst suffering only a modest drop in strength. Accurate temperature measuring equipment, in addition to well equipped facilities, are required to do these procedures on aerospace metals. Table 3,however, shows that, when carbon steel is polished to a bright, clean surface, and then slowly heated, a range of colours appears, due to a thin oxide film forming during the heating process. These colours are related fairly closely to temperatures. The higher temperature achieved during the tempering process, the softer (and tougher) the material will become and vice-versa. Table 3 COLOUR/TEMPERATURE RELATIONSHIP OF CARBON STEELS COLOUR TEMPERATURE Straw 230/240°c Purple 270°C Blue 300°C Dark red 500°C 4.3.3 ANNEALING

The annealing of steel may be for one of the following purposes: • To soften the steel for forming or to improve machinability. • To relieve internal stresses induced by a previous process (rolling, forging, or unequal cooling). • To remove coarseness of grain. Annealing is normally achieved on carbon steel by heating to just above the Upper Critical Limit followed by very slow cooling. In practise the slow cooling rates are achieved by cooling in the furnace or by immersing in a poor thermal conductor such as ashes. The end result is a stress-free, fully softened material, suitable for major forming operations such as deep pressing, drawing, extruding etc. 4.3.4 NORMALISING

This process is similar to annealing, except that the cooling is done in still air. The end result, again, is a stress-free, soft material with uniform fine grain structure. Normalising is commonly used on actual components after heavy machining operations (or welding), prior to the final hardening and tempering processes.


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4.4 SURFACE HARDENING OF STEELS • Unlike conventional through-hardening of steel, it is sometimes desirable to retain a relatively tough (relatively less brittle) inner core, coupled with a very hard surface. This would, typically, be required of a component, which is subjected to high dynamic stresses, yet also has to resist surface wear and would include: • gears (where the teeth need to be hardened) • camshafts and crankshafts (bearing surfaces) • cylinder barrels of piston engines (or landing gear legs). Some materials can be ‘case-hardened’ to achieve this aim. Several methods are used, depending on the parent material and the particular application. 4.4.1 CARBURISING

This is the most common method of case-hardening low-carbon steels and, basically, consists of heating the metal to approximately 900°C, while the component is in contact with a carbon-rich medium followed by a suitable heattreatment. Carbon is generally absorbed into the surface of the heated steel and the rate of penetration is approximately 1mm in 5-6 hours. Low-carbon steels are particularly suited to this type of treatment, as it increases the carbon content and hence the hardness locally. Various methods of carburising are used, the most common ones being: • Pack Carburising. The object is sealed in a container containing a carbonrich (charcoal based) powder and heated in a furnace. The metal is next quenched in oil (not water-which would cause the hard case to peel off). The depth of the hard skin depends on the length of time that the metal is heated. • Gas Carburising. The object is placed in a basket in a furnace, through which is passed a suitable, carbon-rich gas (e.g. methane, propane). • Liquid Carburising. The object is heated to a suitable temperature and then immersed in a hot, salt bath at 900°C. The salts are usually based on sodium cyanide and the process is often called ‘cyanide hardening’. The metal is quenched in water (not oil-which would react unfavourably with the salts).


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4.4.2 NITRIDING

This process involves the absorption of nitrogen (instead of carbon) into the surface of the steel. Suitable "Nitralloy" steels are necessary for this process and they usually contain 1% Aluminium, 1.5% Chromium and 0.2% Molybdenum. A special furnace is used and ammonia gas is circulated through it. The furnace temperature of 500°C converts the ammonia into a nitrogen-rich gas and forms hard iron nitride in the surface of the steel. The case depth, achievable by this process, is less than that by pack carburising, but the major advantage of nitriding is that no hardening or tempering is necessary to achieve the final hardness, and no finish machining is required after nitriding. This, relatively low-temperature process, results in negligible distortion and is much cleaner than the carbon methods. Aircraft piston engine cylinder barrels are particularly suitable for nitriding, as are some crankshaft bearing surfaces and the stems of some aero-engine induction and exhaust valves. Nitrided surfaces must be protected against pitting corrosion, usually (as with engine gears and shafts) by keeping the surface oiled. Note: If certain surfaces of a component are not to be case-hardened, it is necessary to protect them during the carburising or nitriding processes, to locally prevent the hardening agent from being absorbed. Copper plating, nickel plating or a proprietary paste are generally used in such areas. 4.4.3 FLAME/INDUCTION HARDENING

Unlike carburising and nitriding, flame and induction hardening do not add a hardening agent into the surface of a basically softer material. Instead, they are merely techniques for hardening the surface of material by a `local heattreatment'. Steels suitable for these processes already contain sufficient carbon (or other elements) to attain a high degree of hardness if heated and quenched. Only the surface is locally heated (by a flame or electrical induction coil), and the heated surface is then immediately quenched by water jets. The flame or induction coil is positioned so that it only heats the area required to be hardened. 4.4.4 OTHER SURFACE HARDENING TECHNIQUES

In addition to case-hardening, there are other methods of producing hard surfaces on metals, such as by electro-plating, welding, bonding, and metal spraying. All usually involve adding a harder surface metal to the parent material.


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4.5 ALLOYING ELEMENTS IN STEEL As discussed earlier, iron has few practical uses in its pure state. Adding small amounts of other materials to molten iron, however, dramatically changes its properties. Some of the more common alloying elements include carbon, sulphur, silicon, phosphorus, nickel and chromium (also referred to as chrome). 4.6 CARBON Carbon is the most common alloying element found in steel. When mixed with iron, compounds of iron carbide form and it is the carbon in steel that allows it to be heat-treated to obtain varying degrees of hardness, strength and toughness. The greater the carbon content, then the more receptive the steel becomes to heat-treatment and, while its strength and hardness increases, its malleability and weldability decreases. 4.6.1 LOW-CARBON STEEL

Low-carbon steels contain between 0.1% and 0.3 % carbon and are classified as SAE 1010 to SAE 1030 steels. They are used in such items as locking wire and cable bushings and, in sheet form, they are used for low-load applications. Low-carbon steels weld easily but do not accept heat-treatment very well. 4.6.2 MEDIUM-CARBON STEEL

These steels contain between 0.3% and 0.7 % carbon. The increased carbon assists in heat-treatment while still retaining reasonable ductility. Mediumcarbon steels are used for machining or forging and where surface hardness is required. 4.6.3 HIGH-CARBON STEEL

The carbon content of these steels, ranges between 0.5% and 1.5 % and this makes them very hard. High-carbon steels are primarily used in springs, files and in most cutting tools. 4.7 SULPHUR Sulphur causes steel to be brittle when rolled or forged and so it must be removed during the refining process. If it proves impossible to remove all of the sulphur, then manganese, which is harmless to the steel can be added to the metal (to form manganese sulphide),. The manganese also improves forging by making the steel less brittle during the forming processes. 4.8 SILICON When silicon is alloyed with steel, it acts as a hardener and, used in small quantities, it also improves ductility.


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4.9 PHOSPHORUS Phosphorus raises the yield strength of steel and improves a low-carbon steel’s resistance to atmospheric corrosion. The steel tends to be brittle when cold, so no more than 0.05 % phosphorus is normally used in steel production. 4.10 NICKEL Nickel is used extensively for alloying with steel as follows: • In the range of 1% - 5% there is a marked improvement in strength (and hardness) without lowering ductility. This high-strength, tough steel is widely used for highly stressed parts. • At about 25% nickel, the steel becomes highly corrosion-resistant, heatresistant and non-magnetic. • At 36% nickel, a unique steel (known under its trade name as ‘Invar’) is created. This has the lowest coefficient of expansion of any metal (1/20th that of steel) and is excellent for master gauges and instruments. • Because of the effect of such amounts of nickel on the expansion properties of steel, a range of nickel-steels can be purpose-made, to trim the coefficient of expansion to specific needs. These alloys are used in thermostats, spark plug electrodes etc. 4.10.1 NICKEL ALLOYS

When the amount of nickel present is predominant, then the material becomes known as a Nickel Alloy, many of which are widely used in industry. One of the most important nickel-based alloy groups is the nimonics. These are a family of alloys, containing 50% - 80% nickel, with the balance being mainly chromium (chrome) with some titanium and aluminium. Nimonic alloys are used in hot air control ducting, for gas turbine engine combustion chambers and turbine blades because of their extremely low coefficient of expansion at elevated temperatures. Other ranges of nickel-based alloys come under the trade names of Inconel and Hastelloy, which are also temperature-resistant and corrosion-resistant. Another common nickel alloy is Monel. This metal (68% nickel and 29% copper, with iron, manganese, silicon and carbon) has excellent resistance to both corrosion and chemical attack, is tough, ductile, reasonably strong (equivalent to mild steel) and is non-magnetic. It is used in many marine applications, for surgical apparatus and for aircraft rivets. Normally Monel does not respond to heat treatment but, when alloyed with a small amount of aluminium (2% - 4%), it can be hardened to double its strength. This version is known as ‘K-Monel’.


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Nickel adds strength and hardness to steel as well as increasing its yield strength. By slowing the rate of hardening during heat-treatment, the depth of hardening can be increased and the steel’s grain structure made finer. SAE 2330 steel, containing 3.0 % nickel and 0.3 % carbon, is used in the manufacture of bolts, nuts, rod ends and pins. 4.11 CHROMIUM (CHROME) When small amounts of chrome are added to steel, the strength and hardness increases, but there is some loss of ductility. 1.5% chrome, in a high-carbon (1%) steel, results in a very hard material which is used extensively for instrument pivots and in ball and roller bearings. Low chrome (1.5%-3%) steels are used for high tensile fasteners and are suitable for nitriding. Chromium can also be electrolytically deposited onto metals, to provide hardwearing surfaces, such as those required in cylinder bores. Steels containing 12% or more chrome, are very corrosion-resistant. Stainless (SS) Steels or Corrosion Resistant Steels (CRS) come into this category. One particular stainless steel is designated ‘18/8 Stainless’, which contains approximately 18% chrome and 8% nickel. These stainless steels are used extensively in engine parts, particularly for hot applications and for exhaust areas where their corrosion resistance is vital. 4.11.1 NICKEL-CHROME STEEL AND ITS ALLOYS

This term is used when the amount of nickel present is greater than the chrome content. A wide range of such steels exists, but the low nickel-chrome alloys are suitable for through-hardening or case-hardening. The nickel content is around 3%-5% and the chrome ranges from 0.5%-1.5%. Crankshafts and connecting rods are often made from this group. High nickel-chrome alloys (65%-85% nickel, 15%-20% chrome) have a high electrical resistance and are often used as heater elements. By adding both metals, in appropriate percentages, steel, which is suitable for high-strength structural applications, is produced. Nickel-chromium steels are used for forged and machined parts requiring high strength, ductility, shockresistance and toughness. 4.12 COBALT Cobalt is often included in High-Speed Steel (HSS) in addition to chrome, vanadium, molybdenum, and tungsten (to improve still further the ability to cut at high working temperatures). Cobalt is included in high-strength, permanent magnets, in some of the nimonic alloys used for high-temperature components in gas turbine engines and cobalt is also found in a range of temperature-resistant alloys called ‘Stellite’ (used in piston engine valves and for cutting tools)


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4.13 VANADIUM When added to steel, vanadium improves the strength without loss of ductility, but also greatly improves its toughness and its resistance to fatigue. Because of the improved tensile and elastic properties, Valve (and many types of other) Springs, usually include vanadium. Small amounts of vanadium are included in certain nickel-chrome steels and good quality engineering tools. Vanadium, when combined with chromium, produces a strong, tough, ductile steel-alloy. Amounts of up to 0.2 % vanadium improve grain structure, ultimate tensile strength and toughness. Ball bearings are also made from chromevanadium steel. 4.14 MANGANESE When small amounts of manganese are added to steel (up to 1.5%) the result is a steel which is strong and hard (similar to nickel-chrome steel). Such steel is often used for shafts and axles 11%-14% manganese steel has very unusual properties and is extremely useful. When this material is heated to approximately 1000°C and water-quenched, its structure becomes austenitic and, although it is only moderately hard, any attempt to cut it, or abrade it, results in the local formation of hard martensite and it thus becomes highly resistant to cutting or abrasion. Because of this peculiar property, it is used extensively for rock drills, stone crushers, and railway lines at junctions etc. Small amounts of manganese are used in steel production and in welding rods since it acts as a purifying agent by reducing oxidation. 4.15 MOLYBDENUM One of the most widely used alloying elements for aircraft structural steel is molybdenum. It reduces the grain size of steel, which increases its impactstrength and elastic limit. Other advantages are an increase in wear-resistance and high fatigue-resistance, which is the reason why molybdenum-steels are found in structural members and engine parts. 4.16 CHROME AND MOLYBDENUM Chrome-molybdenum steel is, probably, the most commonly used alloy steel in the aircraft industry. Its SAE 4130 designation denotes an alloy of 1 0% molybdenum and 0.3 % carbon. It machines well and is easily welded by gas or electric arc methods, as well as responding well to heat-treatment. Its use in aircraft construction includes landing gear, engine mountings and many engine components. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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4.17 TUNGSTEN Tungsten has an extremely high melting point and adds this characteristic to the steel with which it is alloyed. Tungsten steels retain their hardness at elevated temperatures, and are typically used for contact-breaker contacts (in magnetos), and also for high-speed cutting tools. 4.18 MARAGING STEELS Conventional very high tensile steels have a high carbon content and are, thus, very hard and difficult to work and also tend to be somewhat brittle. To combat these shortcomings, maraging steels were developed. These steels are over 50% stronger than normal high tensile steels and yet are very tough and easy to machine. These properties are achieved by the almost total elimination of carbon and by alloying with nickel, cobalt and molybdenum in such a way that it can be precipitation hardened. Maraging steels can only be used for special, high-stressed applications (due to cost, which is about three times that of conventional alloy steels). They are used for some airframe and engine components and can be nitride hardened


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INTENTIONALLY BLANK


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5 AIRCRAFT MATERIALS - NON-FERROUS A metal in which there is little or no iron is said to be non-ferrous. The list of nonferrous metals is quite impressive – and their uses make very interesting reading, if it is intended to follow a career in metallurgy – but, for the purposes of this course, the topics must be confined to the more common non-ferrous metals, their qualities and their uses in aerospace engineering. 5.1 PURE METALS Certain non-ferrous metals, such as aluminium, copper and lead, are used in the commercially ‘pure’ state for engineering purposes – usually in the form of sheets, tubes, wires or as thin coatings on other metals. Cadmium, chromium, nickel, tin and zinc are also often used to provide protective coatings on other metals in order to retard the effects of corrosion. Precious metals, such as gold, platinum and silver have been used for special work in high-grade electrical instruments, aircraft windshields and, of course, space vehicles. Mercury (quicksilver) – the only metal to remain liquid at room temperature – may be found in certain types of barometers, discharge lamps, small, electrical circuit breakers, pressure gauges and vacuum pumps (it can also be found in the detonators of some explosive devices). In a similar manner to steels, it has been discovered that tremendous advantages are to be gained by alloying non-ferrous metals with each other and, indeed, with other (ferrous) metals and elements. Aluminium, copper, magnesium and titanium alloys are among the more common non-ferrous metals that are used in aircraft construction and repair. 5.1.1 PURE ALUMINIUM

Pure aluminium is extracted from the mineral rock bauxite (named after the town of Les Baux, in France, where it was first found) . It is a soft, weak, ductile and malleable metal. Aluminium is approximately one third the weight of steel and has approximately one third the stiffness of steel. While its strength may be improved by cold working, it remains a low-strength material. Aluminium is highly corrosion-resistant, due to the rapid formation of a thin, but very dense oxide surface film, which limits further corrosion and it is an excellent conductor of electricity (and heat).


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5.1.2 PURE COPPER

Copper also has the ability to retard the progress of corrosion, by developing a patina of green copper carbonate (also called verdigris or aerugo) on its surface. With a conductivity (of electricity and heat) second only to silver, and having the ability to be beaten, cast, drawn, forged, pressed, rolled or spun into many different (and often complicated) shapes, copper is a very versatile metal. Despite a relative density of 8.96, copper’s ductility and malleability allow it to be used in electrical systems (in bus bars, bonding, electrical motors, wiring etc.), though neither copper, nor its alloys, find much use as structural materials in the construction of aircraft. 5.1.3 PURE MAGNESIUM

WARNING;- WATER MUST NOT BE USED TO EXTINGUISH MAGNESIUM FIRES. Two thirds the weight of aluminium (with a relative density of 1.74), no metal can be cut, drilled, filed or shaped so easily as magnesium – provided that certain precautions are taken to prevent it over-heating. Magnesium burns readily, especially in small particles and dust. Great care must be taken when filing and grinding this metal and, if a fire should occur, it must be extinguished with dry sand or an appropriate powder extinguisher but WATER MUST NOT BE USED. Magnesium is obtained primarily from electrolysis of seawater or brine from deep wells. In its pure state it lacks sufficient strength and characteristics for use as a structural metal. It can, however, be alloyed with a range of other elements to greatly improve its strength. These elements include aluminium, manganese, thorium, zirconium, and zinc. 5.1.4 PURE TITANIUM

WARNING:- TITANIUM FIRES MUST BE EXTINGUISHED WITH THE CORRECT EXTINGUISHANT (DRY ASBESTOS WOOL AND CHALK POWDER) AND NOT WATER. Pure titanium at approximately 56% the weight of stainless steel, has almost the same strength as iron. It is highly resistance to corrosion, non-magnetic and is readily shaped by all of the methods, which relate to steel. Titanium is also soft and ductile. Care should be taken when working with titanium. Titanium fires usually start through high-speed rubbing. The low thermal conductivity of titanium prevents the EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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rapid dissipation of heat, which progressively builds up locally, until ignition finally occurs. Accumulations of small particles of swarf and dust are a possible fire risk and all such accumulations should be avoided. 5.2 ALUMINIUM ALLOYS Because pure aluminium lacks sufficient strength to be used for aircraft construction and, to achieve medium/high-strength properties, aluminium must be alloyed with other elements. The most common alloying elements in the wrought aluminium alloys are copper, manganese, magnesium and zinc. A common element used when casting aluminium is silicon. Aluminium alloys may be designated as being either heat-treatable or as nonheat-treatable, though both types can be strengthened and hardened through work-hardening (or strain-hardening). This process requires mechanically working an alloy at a temperature below its critical range and can be achieved by rolling, drawing or pressing Note:- Alloys, which have aluminium or magnesium as their base elements, are referred to as Light Alloys, while the remainder are termed Heavy Alloys. 5.3 IDENTIFICATION OF ELEMENTS IN ALUMINIUM ALLOYS Various national Standards Institutions have evolved their individual systems for identifying the many variants of aluminium alloys (in a similar manner to that shown with SAE Steels). While it would be impossible (and unsafe) to attempt to memorise them all, these notes provide examples of the American system of identifying aluminium (or aluminum) alloys. American aluminium alloys are classified by a code, which refers to the element that makes up the major percentage of the alloy As previously stated, the elements most commonly used for alloying with aluminium are copper, manganese, silicon, magnesium, and zinc. Table 4 shows a four-digit number, which identifies aluminium, either in its commercially ‘pure’, or in its alloyed state. The first digit of the designating code represents the major alloying element, while the second digit of the code indicates a specific alloy modification, such as controls over impurities. The last two numbers of the 1xxx group indicate the hundredths of 1% above the 99% of pure aluminium. For example, if 75 were the last two digits, the metal would be 99.75%pure.


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The 2xxx to 8xxx groups use the last two digits to identify the different alloys in the group. Table 4 American System of Identifying Alloying Elements with Aluminium Code

Major Alloying Element

1xxx 2xxx 3xxx 4xxx 5xxx 6xxx 7xxx 8xxx

aluminium copper manganese silicon magnesium magnesium & silicon zinc other elements

In the 1xxx group, commercially ‘pure’ aluminium (over 99% pure) is good for corrosion resistance, has good electrical and thermal conduction properties, is easy to work but is not very strong. The 2xxx group uses copper as its major alloying element. The major benefit of copper is a large increase in strength, although if the alloy is not correctly heattreated, intergranular corrosion can occur between the aluminium and copper grains within the metal. These are probably the commonest aluminium alloys used in aircraft construction. The 3xxx group has manganese as its major alloying agent and it is not possible to heat-treat. The 4xxx series utilises silicon as its major element. This lowers its melting point and improves its welding and brazing capabilities. The 5xxx group has magnesium as the main alloying element. This is good for welding and corrosion resistance although, if exposed to high temperature or cold working, it can corrode quite badly. The 6xxx group has silicon and magnesium added to the aluminium. This makes the alloy heat-treatable and with good forming and corrosion resistance properties. The 7xxx alloys are made harder and stronger by the addition of zinc. These are difficult to bend and are more often used where flat plates are required.


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5.4 CLAD MATERIALS Though strong, aluminium alloys are not as resistant to corrosion as pure aluminium and, for external use such as skins, the high-strength sheet has a thin layer of pure aluminium hot-rolled onto the surfaces. These are then known as clad materials with commercial names such as Alclad, and Pureclad. Alclad is a ‘pure’ aluminium coating that is rolled onto the surface of an aluminium alloy, which may, then, be heat-treated. The thickness of the coating is approximately 5% of the material thickness on each side. For example, if an alclad sheet of aluminium alloy has a thickness of 1.2 mm (0.047”), then 0.06 mm (0.0024”) of ‘pure’ aluminium is applied to each side. This clad surface greatly increases the corrosion resistance of an aluminium alloy. If, however, the cladding is penetrated, corrosive agents can attack the alloy under the cladding. For this reason, sheet metal should be protected from scratches and abrasions. In addition to providing a starting point for corrosion, abrasions can create potential ‘stress raisers’ (points from which cracking can initiate). 5.5 HEAT-TREATMENT OF ALUMINIUM ALLOYS WARNING:- SAFETY PRECAUTIONS MUST BE OBEYED WHENEVER YOU ARE INVOLVED WITH HEAT-TREATMENTS. BATHS, OVENS AND FURNACES ALL PRESENT DANGERS – FROM CORROSIVE AGENTS, HEAT AND ELECTROCUTION – EXERCISE EXTREME CAUTION WITH THESE METHODS AND WEAR ADEQUATE PROTECTIVE CLOTHING (APRONS, FACE MASKS, GOGGLES AND GLOVES) WHERE NECESSARY AND ENSURE THE CORRECT FIRE-FIGHTING APPLIANCES ARE AVAILABLE. Heat-treatment is a series of operations involving the heating and subsequent cooling of alloys in their solid state. Its purpose is to make the metal harder, stronger and more resistant to impact but it can also make the metal softer and more ductile for working into a required shape (bending etc.). One treatment cannot give all of these properties. Some treatments are achieved at the expense of others when, for example, a hardened material usually becomes more brittle. The heating and cooling cycles occur in most treatments and it is only the time and temperatures which differ. Aluminium alloys have two main heat-treatments, which are referred to as solution heat-treatment and precipitation heat-treatment.


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The procedures for heat-treating aluminium alloys are critical if correct properties are to be obtained. Uniform heating is absolutely essential and two methods are used: • a muffle furnace • or a salt bath The muffle furnace uses hot air, which circulates around an inner chamber in which the aluminium alloy is placed. The salt bath employs molten mineral salts (water would evaporate long before the required temperatures were reached. The salts (usually nitrate of soda or similar) are solid at room temperature, but become liquid when they are electrically heated. Gradual heating of the bath is necessary to avoid spattering or spitting. The aluminium alloy (pre-dried, also to avoid spattering) can then be submerged within the heated liquid. Another precaution when using a salt bath is to avoid any adjacent flames or sparks, because the salts are inflammable. Accurate thermostatic control is vital, as narrow tolerances on temperatures are specified (typically plus or minus 5ºC). Quench tanks must be sited nearby the furnace or salt bath, to avoid delay between removing from the heating source and quenching. Most quench tanks contain cold water but hot water is sometimes specified (especially for heavy sections e.g. large forgings). Limits are also stipulated for the permissible period between heating and quenching which is known as the lag-time (typically 10 seconds max.). If these lag-times are exceeded, material properties or corrosion resistance may be adversely affected. If the cooling rate, during quenching, is too slow this may also affect the corrosion resistance. Thorough washing of the material is essential after salt bath heat-treatment to remove any salt residue. There is no limit to the number of times that heat-treatment may be carried out on normal aluminium/copper alloys but, if the material is clad with pure aluminium, for corrosion resistance (Alclad), then a maximum of three treatments is imposed. This is to limit the migration of copper, from the alloyed material, into the pure aluminium cladding, which would significantly reduce its corrosion resistance. 5.5.1 SOLUTION TREATMENT

Solution treatment is sometimes called ‘re-crystallisation H.T’. This operation serves to distribute the copper uniformly throughout the aluminium (i.e. to create a solid solution). The heating may be achieved (as previously stated) in an oven EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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or, more commonly (to obtain better overall heating), in a bath of special, molten salts. However, although the aluminium can accommodate 5% or so of copper in solid solution at high temperature, this condition is unstable at lower temperatures and, after the alloy has cooled to room temperature, most of the copper slowly comes out of solution and separates into local ìslands' of copper aluminide. By cooling the alloyed metals very quickly (quenching), the copper becomes trapped 'in solution', making the aluminium very strong. 5.5.2 AGE-HARDENING

The gradual formation of the copper alumide ‘islands’ (also referred to as ‘slip’), causes an increase in hardness and strength and these properties reach maximum values after several days (or weeks in some instances). Because of the time lapse involved, this gradual hardening is termed ‘age-hardening’. Although copper may be the major alloying element (in the ‘2000 series’ alloys) other elements, including magnesium and manganese can also be present. Although the aluminium/copper alloys are the most common age-hardened, highstrength metals, they are not unique. Aluminium, when alloyed with 5%-7% Zinc, is also able to be age-hardened. This is a more modern alloy than the aluminium/ copper type and is the highest-strength aluminium alloy in general use. This alloy is used in heavy loaded applications such as Main Spars, Landing Gear and Mainplane Attachment brackets etc.. 5.5.3 ANNEALING

Annealing, as with steel, serves to soften the aluminium alloy, to enable it to be worked without cracking. Even in this condition, ageing will gradually occur and 24 hours is the normal limit for working after annealing, although this can be extended if the material is stored under refrigerated conditions to slow the ageing process. A temperature of -5ºC will provide approximately 2 days’ delay while one of -20ºC will provide approximately 1 week’s delay in the age-hardening process The maximum for refrigeration is approximately 150 hours at -20°C. Typical annealing procedure may be achieved by raising the temperature of the alloy to between 340°C and 410°C. The alloy is then cooled slowly at about 10°C per hour (rates will differ with each particular alloy), until it reaches a predetermined temperature. At this point it is allowed to cool naturally. These, heat-treatable type, alloys must never be installed in an aircraft structure while in the annealed state, since material properties and corrosion resistance will be severely affected. Note: Alloys, in the annealed state, are very prone to corrosion.


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5.5.4 PRECIPITATION TREATMENT

Solution-treated aluminium alloys are comparatively soft, immediately following quenching although, with time, the metal gradually becomes harder and gains strength. When the alloys are left at room temperature, after quenching, the hardening process (natural ageing), and can take from several hours to several weeks. An aluminium/copper alloy, for example, is only at 90% strength within 30 minutes of quench, but is at maximum strength after four or five days. We have already discussed how the natural ageing process can be drastically retarded (allowing the metal to be kept in a soft condition until required for use), by storing the alloys at sub-zero temperatures (refrigeration) for prescribed periods of time. Alternatively, following quenching, by re-heating the metal to a lower temperature than that employed for the solution treatment and allowing it to ‘soak’ at that heat for a period of time, the ageing process (and, thus, the hardening of the alloy) can be accelerated. This process is referred to as artificial ageing or precipitation treatment.


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IDENTIFICATION OF HEAT-TREATED ALUMINIUM ALLOYS

Aluminium alloys that have been subjected to heat-treatment are usually identified by markings that indicate the heat-treatments involved. Three typical identification systems are those of the British Standards Institute (BS), the Ministry of Supply (MoS), and the American systems as can be seen in Table 5. Table 5 IDENTIFICATION MARKINGS OF HEAT-TREATED ALUMINIUM ALLOYS BS System

M O OD T W WP MoS System

A N W WP American System

T3 T4 T6 T8 T9

Meaning

As manufactured state Annealed state Annealed and lightly drawn Solution-treated, no precipitation required Solution-treated, can be precipitated Solution-treated and precipitation treated Meaning

Annealed state Solution-treated, no precipitation required Solution-treated, and requires precipitation Solution-treated and precipitation treated Meaning

Solution-treated and cold worked Solution-treated only (naturally aged) Solution-treated and artificially aged Solution-treated, cold worked and artificially aged Solution-treated, artificially aged and cold worked

An example of one of these marking systems would be an alloy with the designation 2024-T4, which indicates an aluminium/copper alloy that has been solution-treated only, and then naturally aged Apart from these systems, many other exist world-wide, but the British systems are, broadly, confined to three basic ones for light alloys. • British Standards for general engineering use BS 1470 -1475. In this series the prefix N is used to denote non-heat-treatable aluminium alloys and prefix H for the heat-treatable alloys. • British Standards for aerospace use: BS X LXX. (The "L" series) e.g. BS 3 L72 indicates the 3rd amendment to the basic L 72 spec. LM - indicates a cast material. The wrought materials are commonly abbreviated to L71, L72, L 73 etc. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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Examples of some of these aircraft BS codes are: a) L159 DURAL* Solution-Treated - Artificially aged b) L163 ALCLAD Solution-Treated - Naturally aged *DURAL is, actually, a Trade name for an Al/Cu/Mg/Si/Mn alloy, originally manufactured by the Duren Aluminium Company (Germany), but it tends to be used as a generic name for similar alloys, regardless of source of manufacture. • D.T.D. Specifications: - these are material identification numbers issued by the Directorate of Technical Development (a Ministry Department) for specialised applications. i.e. when widespread use is not anticipated. If such a material finally becomes commonly used, a British Standards specification is compiled and issued. 5.7 MARKING OF ALUMINIUM ALLOY SHEETS Sheet material, for aerospace use, is marked ‘all over’ with the specification identification, in regular lines, usually in a blue (or green) ink e.g. ‘7075 - T6’, along with a batch number and its thickness, to avoid confusion with similar looking metals. Some sheets may also have alternate lines of red numbers/letters, which indicate that heat-treatment is needed before assembly. These red numbers/letters then disappear when the necessary heat-treatment is done. It is imperative that only the correct specifications and thicknesses of materials are used in the construction of aircraft structures.


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5.8 CAST ALUMINIUM ALLOYS These are not used extensively on airframes mainly due to their lack of strength, poor fatigue characteristics and lack of elasticity when compared to the wrought aluminium alloys. The lack of elasticity is particularly relevant, as the very nature of an airframe structure requires the ability to flex considerably without cracking. Although their use is obviously limited on airframes, cast aluminium alloys are used extensively on engines, where there is a need to produce complex cored shapes such as crankcases, drive casings, cylinder heads etc. No other method than casting would be viable for such items. The stresses can be kept to a modest level on these parts by producing robust castings of adequate stiffness. Very few non-heat-treatable cast alloys are used in aerospace applications and, for high-duty engine casings and pistons, some very strong, temperatureresistant alloys exist. One of the most common in the category is RR 58 (sometimes known as `Y' Alloy), which is an age-hardening material containing approximately 2½% copper, l½% magnesium, 1½% nickel, and l% iron. A derivative of this material was also used (in wrought form) for the skin of the supersonic Concord aircraft, due to the high metal temperatures encountered. Cast aluminium alloys often contain silicon, which creates high fluidity and, thus, is good for producing complex shapes. It also reduces the coefficient of linear expansion, so is often included in piston castings. 5.9

MAGNESIUM ALLOYS

WARNING;- WATER MUST NOT BE USED TO EXTINGUISH MAGNESIUM ALLOY FIRES. Magnesium alloys are used for castings and, in their wrought form, are available as sheet, bar, tubing and extrusions. They are among the lightest metals having sufficient strength and suitable working characteristics for use in aircraft structures. There are some serious disadvantages to using magnesium alloys in aircraft construction. These include a high susceptibility to corrosion and cracking. The corrosion problem is minimised by treating the surface of the metal with chemicals, which form an oxide film, to prevent oxygen reaching the metal. Another way of minimising corrosion is to use hardware such as rivets, nuts, bolts and screws that are made from compatible materials. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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The cracking problem contributes to the difficulty in shaping magnesium alloys and, thus, limits its use. One method used to overcome the tendency for cracking is to form the metal whilst it is hot. Magnesium alloys can also be solution heat-treated, which will improve their tensile strength, ductility and resistance to shock. To improve their hardness and yield strength they can also be precipitation heat-treated after the solution heattreatment. 5.10 COPPER ALLOYS Of those (Heavy) alloys that use copper as a base; brasses, and various bronzes are the primary types used on aircraft. Brasses may contain zinc and small amounts of aluminium, iron, lead and other elements such as manganese, nickel (and even very small amounts of tin!). Depending on the percentage content of zinc, brass can be made ductile (30%35% Zn) or strong (45% Zn). Bronze is a copper alloy that contains comparatively higher percentages of tin and is usually found in the form of castings. A true bronze contains up to 25% tin, and bronze, along with brass, is used in bushings, bearings, valves and valve seats. Bronzes with less than 11% tin are normally used for tubes and pipes. There are other copper alloys that contain practically no tin and yet are still referred to as ‘bronzes’. High-Tensile Brass, for instance, because of its manganese content is called ‘Manganese Bronze’, while Phosphor and Silicon bronzes also contain practically no tin. Wrought aluminium bronzes are almost as strong as medium-carbon steel while cast aluminium bronzes are found in bearings and pump parts Probably, the most common of these is Beryllium Bronze. This contains 97% copper, 2% beryllium and small amounts of nickel to increase its strength. Once it has been heat-treated, beryllium bronze is very strong (300-400 Brinell) and is used for diaphragms, precision bearings and bushings, ball bearing cages and spring washers. Leaded Bronze is found in the bearings of some aero engines. The very high pressures (and speeds) tend to squeeze the lubricant out of normal journal bearings, so the addition of lead acts as a sort of lubricant in the event of the oil film breaking down. Solder is a general term frequently used for joining metals together. The principal types are ‘soft solder’ (which is a mainly lead-tin alloy), and ‘hard solder’ which is an alloy of copper, silver and zinc.


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5.11 TITANIUM ALLOYS WARNING:- TITANIUM ALLOY FIRES MUST BE EXTINGUISHED WITH THE CORRECT EXTINGUISHANT (DRY ASBESTOS WOOL AND CHALK POWDER) AND NOT WATER. Titanium alloys, apart from being light and strong, also have excellent corrosion resistance, particularly in a salt-laden atmosphere. To prevent reaction with oxygen and nitrogen, in its pure form, titanium is treated with chlorine gas and a reducing agent, to produce a coating of titanium dioxide. There are three types of titanium, which are called alpha, alpha-beta and beta. They have different strength and forming properties, depending on their heattreatments. Commercially pure titanium is ‘non-heat-treatable’ (It can be annealed, but its strength/hardness cannot be improved by heat-treatment.). When suitably alloyed, titanium based materials are heat-treatable. The strengthening is immediate i.e. it is not an age-hardening material. Titanium alloys are used extensively in aerospace gas turbines, but their use is limited on subsonic civil airframes to fasteners, and high temperature areas such as engine bays, heat shields, hot zone bulkheads, air ducts etc. In appearances titanium is similar to 18/8 stainless steel. Two practical methods of identification apart from weight are: • spark test - a light touch of a grinding wheel will produce a brilliant white trace, ending in a brilliant white burst. • moisten the titanium and draw a line on a piece of glass - this will leave a dark line similar to a pencil mark. 5.12 WORKING WITH TITANIUM AND TITANIUM ALLOYS CAUTION: DO NOT STAND ON, OR PUSH AGAINST, THIN-WALLED TITANIUM STRUCTURES. LOCAL HARDENING WILL CREATE STRESS RAISERS WHICH COULD LEAD TO STRUCTURAL FAILURE. DO NOT ALLOW HOT OIL TO DRIP ONTO TITANIUM. IT WLL CAUSE HYDROGEN EMBRITTLEMENT. THERE IS NO REPAIR PROCEDEURE FOR THIS TYPE OF CONTAMINATION OF TITANIUM. AN INCORRECT CLEANING AGENT CAN ALSO CAUSE HYDROGEN EMBRITTLEMENT IN TITANIUM COMPONENTS Titanium materials are, generally, not susceptible to normal corrosion attack, but it has been established that stress corrosion cracking can take place in some welded structures which are exposed to trichloroethylene and other chlorinated


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hydro-carbons (the alloys most affected in practise being the titanium-aluminiumtin family). Titanium may also show evidence of deterioration in the presence of salt deposits or metal impurities, especially at high temperatures. It is, therefore prohibited to use steel wool, iron scrapers or steel brushes for the cleaning of, or for the removal of corrosion from, titanium components If titanium surfaces need cleaning, then hand-polishing, or the use of soft bristle fibre brushes, with aluminium oxide compound or a mild abrasive may be permissible. Use only the recommended procedures outlined in the relevant Maintenance or Overhaul Manual When it is necessary to machine a welded titanium structure, or doubt exists regarding the use of cutting fluids with a particular titanium alloy, the material manufacturer should be consulted 5.12.1 DRILLING TITANIUM

Rigidity is essential when drilling titanium and titanium alloys so that thin-wall structures must always have a backing support. Centre drilling should always be used, instead of centre punching, as the local work-hardening caused by centre punching will cause difficulty in starting the drill and will also tend to make the drill wander as well as blunt the drill point. A High-Speed Steel (HSS) drill, having a point angle of 105º to 120º, with a helix angle of 38º and a thickened web is recommended. It is important that a stub (i.e. short) drill should be used. For holes of more than 6 mm (¼ inch) diameter, a 90º or ‘double-angled’ point is better. Drills must be precision ground and special care must be taken to ensure that the drill tip is completely central, as any off-set of the tip will cause work hardening as a result of friction of the non-cutting edge. Flood lubrication with a cutting fluid of low viscosity helps to reduce frictional troubles. High quality soluble oils, used in the diluted form recommended by the manufacturers, or chlorinated or sulphured oils, should be used in generous quantities for all machining operations. Chlorinated solvents should be removed, after machining. For satisfactory drill life, drill surface speeds within 3 to 13 metres (10 to 40 feet) per minute are used, otherwise work hardening is likely to result. A continuous feed of 0.05 to 0.1mm (0.002 to 0.005 inch) per revolution for holes below 6 mm.(0.25 inch) diameter, and of 0.1 to 0.2 mm (0.005 to 0.010 inch) per revolution for larger holes is recommended. Positive power feed must be employed whenever possible. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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6 METHODS USED IN SHAPING METALS There are four basic methods of converting raw material into the required manufactured shape whilst also achieving the desired material structure. They are casting, deformation, machining, and various forms of fabrication (i.e. the joining together of smaller pieces or particles of material to form a larger object). Welding, adhesive bonding, mechanical fasteners or even powder metallurgy come under this latter heading. Casting exploits the fluidity of a liquid as it takes shape and solidifies in a mould. Deformation exploits the remarkable property of materials (mostly metals) to flow plastically in the solid state without deterioration of their properties. Processes such as these, result in a minimum of material waste. Machining processes provide excellent precision, but the process generates a large amount of waste material. Fabrication techniques enable complex shapes to be constructed from simpler particles or units. 6.1 CASTING This involves the pouring of molten material into a shaped mould and allowing it to solidify to that shape. It is an ancient process, which enables complex shapes to be produced in a wide range of materials in a single-step operation. Cast components can range in size from the small teeth of a zip, to large casings of several metres in diameter. Ocean-going ships’ propellers, up to 10 metres in diameter, are produced this way. Modern casting techniques have resulted in: • high quality (i.e. minimum porosity and reasonably defect-free products) • high production rates • good surface finish • small dimensional tolerances • the ability to cast a very wide range of materials. Moulds are made in a variety of materials including plaster and ceramics but, by far, the most widely used are those of sand and metal. 6.1.1 SAND-CASTING

The two basic types of sand-casting are: • Removable/re-usable pattern (usually wood or metal) • Disposable pattern (e.g. polystyrene patterns, which vaporise when the metal is poured).


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Although sand-casting is simple in principle, there are many vital aspects of the technique, which are necessary to produce good castings. The sand, for example, must have: •

Adequate binding qualities (to achieve this, a small percentage of clay is added).

•

Suitable porosity characteristics (to permit the escape of gas/steam, formed in the mould). There are different requirements for different metals (e.g. steel and aluminium).

•

Correct grain size and sufficient strength (the sand is graded by means of a sieve and the strength is controlled by the amount of bonding agent present).

•

Suitable temperature resistance (i.e. the sand must withstand the molten metal temperature without fusing/melting).

•

Adequate hardness (the hardness may be checked by the resistance to indentation by a spring-loaded ball).

•

Acceptable moisture content levels (this is usually in the range of 2% to 8% and is checked by weighing the sand before and after drying).

While the characteristics of the sand are important, the design of the mould must also meet certain standards, some of which are: •

The top and bottom halves of the mould (‘cope’ and ‘drag respectively), must incorporate positive alignment features.

•

The pattern must be shaped such that withdrawal from the sand leaves a perfect impression. Tapered faces are, therefore, better than perpendicular faces.

•

Suitable feed channels must be provided for the molten metal to enter the mould. These channels are called the ‘sprue’ and the ‘runners’.

•

Strategically placed reservoirs (called `risers') must be incorporated to ensure proper filling of the mould as the metal shrinks and begins to solidify. Typical steel shrinkage is around 3%-4% and aluminium shrinkage, 6%-7%.

•

The incorporation of vents, where necessary, to permit the escape of gas and steam when the molten metal contacts the sand.

•

Local ‘chills’ are sometimes included in the mould, to encourage more rapid, local solidification of the metal.


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6.1.2 ADVANTAGES/DISADVANTAGES OF SAND-CASTING

The advantages of sand-casting are that it is a simple process, which does not require elaborate equipment and is economical for small batches. It is also suitable for most metals. The major shortcomings are that the process is not very rapid, it is not particularly accurate (due to lack of sand rigidity) and it is not suitable for thin-wall sections. 6.1.3 TYPICAL CASTING DEFECTS

Casting defects vary to some extent, depending on the casting process used, but the most common ones are: • Inclusions (e.g. sand or mould lining material sticking to the surface) • Porosity (usually caused by gas/vapour, which is unable to escape before solidification) • Cold Shuts (when local areas of metal are not molecularly joined, due to solidification occurring too rapidly). • Hot Tears (where the material is cracked by excessive tensile stresses, resulting from thermal contraction). 6.1.4 SHELL-MOULDING

Shell-moulding is a process in which a thin shell is produced, by bringing a mixture of sand and a thermosetting resin into contact with a heated pattern. When a sufficiently thick shell has been produced, the shell is finally cured (backed up by sand or steel shot in a moulding box). The subsequent casting process is then the same as for normal sand-casting. The advantages of shellmoulding over conventional sand-casting are: •

it can be semi-automated, which reduces cost

•

finer sand can be used, which results in a smoother surface finish.

6.1.5 CENTRIFUGAL-CASTING

This technique involves the molten metal being poured into a rotating mould. The process is used for the manufacture of hollow cylinders (e.g. cylinder liners), bronze or white metal bearings etc. The rotation can result in acceleration forces of up to 60g and this produces high-quality, dense castings, since all of the slag migrates to the bore (due to it being of lower density than the metal) and it can then be machined out.


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6.1.6 DIE-CASTING

This process uses a permanent metal mould, which results in more accurate, and better finished, castings than those produced in sand. Die-casting, can be subdivided into ‘gravity’ or ‘pressure’ processes, depending on how the metal is fed into the mould. • Gravity Die-Casting - sometimes known as ‘Permanent-Mould Casting’. This casting process is virtually identical to sand-casting except that the mould (die) is metal. A wide range of metals can be cast and hollow castings are possible if a sand core is used. Fine grain structures are produced, due to the more rapid rate of cooling, compared to that achieved in sand-casting. • Pressure Die-Casting - as implied, molten metal is fed under high pressure (thousands of psi) and held during solidification. Most die-castings are in non-ferrous materials (aluminium, magnesium, zinc, copper and their alloys), because steels have too-high a melting temperature for the metal dies to accommodate. The dies are, usually, made from hard, tool-steels and are water cooled. This process can achieve excellent detail, super finish, low porosity, and thin sections. Expensive equipment is necessary, but very high production rates are possible. Automatic ejection occurs and, on small components, 100 units per minute is not uncommon. Hollow castings cannot be made by die-casting. 6.1.7 INVESTMENT-CASTING (LOST WAX)

This is a very old method of casting (which was used by the ancient Chinese), but it only became of great industrial importance in the 1950's, when gas turbine manufacturing began to increase. The process was ideally suited to the production of complex-shaped nozzle guide vanes and turbine blades which, often, contained tortuous inner passages, very thin sections and had to be cast in exotic materials. The basic process is as follows: •

A master die is made first from an easily worked metal such as brass.

•

Hot wax is then injected into the die, under pressure, to produce a wax pattern.

•

The wax pattern is then removed from the die and coated with a layer of investment material (a ceramic slurry or paste), usually by dipping a number of times.

•

When the investment coating is set, it is then heated to allow the wax to run out, and molten metal is then poured into the investment mould.

•

When cool, the investment coating is then broken away from the cast, metallic component.


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For obvious, reasons this investment-casting process is often referred to as the ‘Lost Wax’ process. It is a technique, which is capable of producing precision castings with a dimensional accuracy of less than 0.1 mm. Surface finish is also excellent, but the major advantage, that the process offers, is the ability to produce accurate, complex shapes which would be impossible by machining. 6.2 FORGING This is a squeezing/hammering technique, which is intended to achieve large deformation/shaping of the material. The process is usually carried out hot (i.e. above the re-crystallisation temperature), so that these large deformations can be attained without being accompanied by any massive, residual stresses. Sometimes a cold forging operation may be necessary but, in this instance, the material will be harder, stronger and pre-stressed (i.e. still containing unrelieved internal stresses). Forging ranges from the simplest form of the hand operations, conducted by the blacksmith, to the massive, mechanical, powered rams, used for very large forgings. The forging hammer will often have a relatively low strike rate, but sometimes high-speed, pneumatic hammers are used for High-Energy-Rate Forming. Forging not only shapes the metal, but also reduces grain size and produces a directional control of grain flow. Both of these are desirable features for many engineering applications, particularly for highly-stressed components, such as crankshafts and especially if they are subject to a mechanical fatigue environment.


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6.2.1 DROP-STAMPING

Drop-stamping, or drop-forging (refer to Fig. 16), involves the use of shaped dies and a heavy drop-hammer, which usually falls under gravity. The piece of material, to be forged, is placed between the top and bottom dies and the drophammer is allowed to fall the necessary number of times for the contact faces of the dies to come together. ‘Flash gutters’ are provided, to accommodate the excess metal (flash), which squeezes out between the top and bottom dies. Connecting rods are typical components made by the drop-forging process.

The Drop-Forging Process Fig. 16 6.2.2 HOT-PRESSING

Hot-pressing is similar, in principle, to drop-forging, but is actuated by one, long, steady, squeezing operation, as compared to a number of blows. This process tends to affect the whole structure of the component, whereas some forging processes, using multi- (but light) blows will, mainly, affect the material closest to the surface. 6.2.3 UPSETTING

Upsetting is, sometimes, called ‘Heading’ and usually involves locally heating of the end or ends of the material, immediately prior to forging. Poppet valves are formed in this way, as well as forged bolts. Sometimes this process is done cold (in which case it is referred to as ‘Cold Heading’), and some rivet heads are formed in this way.


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6.3 ROLLING Rolling can be carried out hot or cold. When done hot ,it is capable of achieving major re-forming/re-shaping, and slabs can be reduced to plate or sheet while bars of circular or rectangular cross section can also be produced. Hot rolling can also produce structural shapes such as ‘H’ or ‘I’ section beams. If the rolling is done cold, it is aimed at improved surface quality, better accuracy, and increased hardness/strength. Hot, dilute, sulphuric acid is used to remove the hot scale from steel prior to cold rolling. The rolling process would also be used to produce the clad (and unclad) sheets of aluminium alloys. 6.4 DRAWING Drawing is a purely, tensile operation, usually carried out hot. Wire, rod and tubing, can be produced by this process, where the material is pulled through a shaped, hardened die. A ductile material is essential. 6.5 DEEP DRAWING/PRESSING This process uses a ram, to deform a piece of sheet metal into a recessed die and is usually done hot. 6.6 PRESSING Pressing involves the use of male and female formers for shaping sheet material. The sheet is placed between the formers, which are then forced together by a powered ram. Pressing is usually done hot (except for the soft, ductile materials). 6.7 STRETCH-FORMING This is a technique used for shaping sheet metal over a stretch-block or former. The sheet metal is firmly gripped by clamps and the sheet is then stretched over the former (by moving the clamps or the former) and the material is stretched beyond its elastic limit so that permanent deformation occurs. This process is convenient for small batches of material (and is particularly financially attractive since only one former is needed) but, local changes of form (concave/convex or vice versa) cannot be produced by this process. 6.8 RUBBER-PAD FORMING In principle this process uses a flexible, rubber-pad, attached to a hydraulic ram, which forces a piece of sheet metal to conform to the shape of a forming block. Like stretch-forming, the process only uses one former, so it eliminates critical matching and alignment problems of conventional pressing, When used for small EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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batches (e.g. aircraft production), low-cost, easy to machine, materials can be used for the forming block. Rubber-Bag forming (Hydro-forming) uses the same principle, but incorporates a flexible diaphragm and hydraulic pressure in place of the rubber pad. 6.9 EXTRUDING The extrusion process, forces hot metal through a shaped die, to produce circular, rectangular, tubular, angular, half-round sections etc. In some respects, the process is similar to drawing, but extruding forces metal from a heated billet, through hardened dies by compression, whereas, in drawing, it is achieved by tension. Malleability is, therefore, an essential material property for the extrusion process. Extruding is normally restricted to aluminium alloys and copper alloys, where extrusion temperatures of 400ºC-500ºC and 650º-1000ºC respectively are used. Steel is extremely difficult to extrude, due to the excessive pressures required. 6.9.1 IMPACT-EXTRUSION

This process is, usually, a cold-forming operation, which is suitable to very soft and malleable materials (e.g. aluminium). The shaped component is formed, by forcing a punch onto a ‘blank’ of material within a shallow recess. The extruded shape results from the metal being forced to escape through the small gap, between the punch and the recess. 6.10 SINTERING Sintering; involves metal, in powder form, which is heated to approximately 70%80% of its melting temperature and then squeezed to shape in a die. The process is often used to form components made from materials with a very high melting temperature (e.g. tungsten). It also allows non-metallic materials, such as graphite and carbon, to be incorporated into the mixture. The operation is usually conducted in a controlled atmosphere (typically argon or nitrogen) to prevent oxidation. Under the high pressures used, a metallurgical bond occurs (diffusion bonding), between the particles of powder. The sintered end-product is, typically, around 10%-20% porous and can then be impregnated with graphite (or high melting-point grease), to provide excellent, self-lubricating properties for plain bearings, bushes etc. Sintering can be used where the combined properties of materials are required, as when copper and graphite are used for electrical brushes (i.e. copper to carry the current and graphite to act as a low-friction contact)


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Tungsten carbide cutting tools can also be produced in this way, by incorporating tungsten carbide particles within a cobalt matrix. Hot, Isostatic-Pressing, uses a similar technique to sintering, but uses higher temperature and very much higher pressures to produce zero porosity. The technique is sometimes used to heal micro-porosity in super-critical castings.) 6.11 SPINNING Spinning is an old process, in which a piece of sheet metal may be formed, to shape, around a rotating former, which is mounted on the spindle of a lathe. The necessary force to deform the sheet metal is generated by a long tool, which is levered about a suitably positioned fulcrum. For thin gauge, soft metals, the tool can be manipulated by hand, while, for thicker gauge materials, a hydraulic actuator is used on a purpose-built machine. Cones, flares, bowls and bell-mouth shapes, are produced by spinning. 6.12 CHEMICAL MILLING Chemical milling is, sometimes, referred to as chemical etching. It is a purely chemical process, not electro-chemical. Although simple in principle, chemical milling offers a method of producing complex patterns and lightweight parts and is used for incorporating integral ribs and stiffeners in sheet metal. Tapered sections can also be easily formed - the unwanted material being eaten away by a suitable chemical. The process is ideally suited to aluminium alloys. The chemical, in this instance, is a hot alkaline solution (usually caustic soda) and, while it is a relatively slow process, its unique advantages make it very attractive for airframe components. The areas, which must not be eaten away by the fluid, are simply protected by a thin layer of plastic, which can be brushed or sprayed on. Although the chemically etched surface is not very rough, a drop in fatigue strength does result and, in critical applications, restoration of fatigue strength is desirable. A light, peening operation, using glass beads or steel shot, achieves this. 6.13 ELECTRO-CHEMICAL MACHINING Using electrolysis and, by making the workpiece the anode of the dc electrical circuit, an electrolyte is pumped rapidly (under pressure) through the gap between the shaped cathode (also referred to as the tool) and the workpiece. The tool is moved slowly towards the workpiece, by a ram, so that metal is progressively removed from the workpiece, until the desired shape is achieved EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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The process is ideal for metals, which are difficult to machine by conventional methods, and the finish achieved is good. High electric current is required, and other, essential, requirements for the process are that the tool needs to be a good conductor (copper or brass) and it must resist corrosion, because the electrolyte is often a salt solution. 6.14 ELECTRO-DISCHARGE MACHINING E.D.M. This process is, sometimes, called spark machining (or spark erosion), because, rather than using electrolysis, the technique involves the removal of metal by the energy (and heat) of electrical sparks, which travel from the electrically negative tool electrode, through a dielectric fluid, and explosively strike the electrically positive workpiece. The intense heat of the strike, causes local particles of metal to instantaneously vaporise, without a molten metal phase (a process known as ‘sublimation’), though, away from the actual centre of the explosion, molten fragments of metal are washed away, with the vapour, by the dielectric fluid. A suitable fluid (usually kerosene) is fed, under pressure, between the electrode and the workpiece, to maintain a uniform electrical resistance. The spark rate is around 10,000 per second and the gap between the tool and the workpiece is critical and must be maintained, throughout the operation, at approximately 0.025 mm - 0.075 mm (0.001 in - 0.003 in). The real advantage of EDM is that, not only is it suitable on materials which are difficult to machine conventionally, but it also excels in its ability to produce highaspect ratio, very small holes of any cross-sectional, in very hard metals. Typical holes achievable, by this method, are in the regions of 0.025 mm diameter x 750 mm deep (0.010 in x 3 in). A novel variation of EDM is a technique sometimes referred to as ‘wire-cutting’, which uses a moving, fine piece of copper or nickel wire as the electrode. The wire, 0.05 mm - 0.25 mm in diameter (0.002 in - 0.010 in), is positioned by, and fed over, two pulleys and resembles a simple band-saw operation. The workpiece is mounted on a table, which can be moved in two axes and, when the table is computer controlled, the wire-cutting process can cut accurate, complex shapes in metals (e.g. dovetails, fir-trees etc.) which are difficult to machine with conventional tools.


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6.15 CONVENTIONAL MACHINING Conventional machining is done, using seven basic techniques, which are: • Drilling/reaming • Turning • Milling • Sawing • Shaping/planing/slotting • Broaching • Abrasive machining (i.e. grinding) These techniques have been well established for many years, and most of the advances, until relatively recently, have been confined to tooling improvements which have permitted higher material removal rates. The early, high-carbon steel tools, have been superseded by high-speed steels (tungsten/cobalt alloy steels), cemented carbides and ceramics. So-called ‘Machining Centres’ have also been developed, which are capable of automatic tool changes and of doing difficult types of machining without the need for transferring work to a different machine and re-setting up. In this way a much more versatile machine tool has evolved. However, the biggest single machining advance in modern times (especially with regard to aircraft manufacture) has been the introduction of Numerically Controlled (NC) machines. NC milling, in particular, has revolutionised airframe manufacture. NC machines are machines in which motion is controlled by a series of numbers, either via punched tape or magnetic tape. Instructions, on the tape, are based on the Binary System (or a variant) which is common to most electronic computing devices. The primary advantage of NC machining is the ability to accurately control the spindle, the tool or the workpiece movements in three directions (x, y and z axes) independently or simultaneously. NC machines are capable of producing compound shapes and contours, and are especially suited to the task of generating integral spars, ribs, and stiffeners in slabs or forgings. NC machines usually incorporate a feed-back system, which ‘tells’ the control unit how much actual movement is made, analysis is then done and final compensation eliminates any error (i.e. the motion ceases when the input and feed-back signals agree). Electrical control of the machine servo-motors, can control movements as small as 0.0005 mm (0.00002 in). CNC machines (i.e. Computer Numerically Control) differ from NC machines only in that the electronic control unit on the CNC machine is more sophisticated in EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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that it is adaptable to a wide variety of software and can accommodate a diverse range of programs. Although the capital cost of NC/CNC machines is high, the following advantages make such machines technically desirable and economically viable, where super-light, complex, high-tech, manufacture is concerned: • Complex shapes with integral features are possible • The number of jigs and fixtures is reduced • A reduction in manufacturing time • Adaptable to short runs • Greater accuracy and consistency • Program can be changed to accommodate modifications 6.16 SUPERPLASTIC FORMING Some Titanium alloys, when heated, become extremely ductile and can plastically deformed without necking occurring This superplasticity can be exploited in the forming process (refer to Fig. 17), when an inert gas is used to blow the material into the required shape

Superplastic Forming Process Fig. 17


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7 AIRCRAFT MATERIALS - COMPOSITE AND NON-METALLIC A composite is something, which is made up from many parts, and this term could be applied to a wide range of engineering materials. These would include not only the metallic alloys, but also the most earliest of all composite materials used by man, – wood (the tough, fibrous, xylem, or water-conducting tissue, of shrubs and trees, which contains lignin and cellulose). Brick, concrete, and glass are among the many other materials, which could be considered as composites. In the aerospace industry, the term ‘composite’ is used when referring to materials, which, in turn, are a combination of fibrous and synthetic resin materials that provide many advantages by their great strength-to-weight ratios. This topic covers a number of different materials, including plastics, resins, natural and synthetic rubbers, adhesives and sealants. Most of these materials will be found in use on modern aircraft. 7.1 PLASTICS The word plastic comes from the Greek plastikos – to mould, and plasticity (as was discussed in The Properties of Metals) is the ability to retain a deformation after the load, producing it, has been removed. Plastics are particularly useful for applications, which involve relatively low-stress levels, where lightness is important, and where low electrical or thermal conductivity is required. The earliest plastic materials (before the synthetics) were those made from the sap, or latex, of certain trees (gutta-percha), the secretions of tiny, scaly insects (shellac) and the softened, moulded parts of the horns of animals. The American inventor, John Wesley Hyatt (in 1869), produced the first synthetic plastic material (used as an inexpensive substitute for ivory), from the cellulose of plants (and called it Celluloid), while the chemist, L H Baekeland (in 1909) developed the first entirely synthetic plastic material (Bakelite), from phenolformaldehyde. Bakelite is hard and fairly brittle. It is often used with a suitable filler material (mica, or wood flour) and is widely used for various electrical mouldings and low-stressed handles. Plastics, however, is now the generic name, used to identify various materials (natural and synthetic), based on long-chain molecules (polymers) of carbon, that can be cast, extruded or moulded into various shapes or drawn out into filaments to be used as fibres. While the two major groups of plastics are the Thermoplastic and Thermosetting compositions, the manufacture of synthetic rubbers (called Elastomers) is also considered to be part of the plastics industry. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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7.1.1 THERMOPLASTIC MATERIALS

Thermoplastic materials, in their normal state, are hard but become soft and pliable when heated (the Greek word thermo – heat). When softened, thermoplastic materials can be moulded and shaped, and they retain their new shape when cooled. Unless their heat limit is exceeded, this process can be repeated many times without damaging the material. Two types of transparent thermoplastic materials are used for aircraft windshields and side windows, and are usually referred to as cellulose acetate and acrylic. Older aircraft used cellulose acetate plastic because of its transparency and lightweight. A disadvantage of cellulose acetate is its tendency to shrink and discolour with time, which has led to it being phased out almost completely. Cellulose acetate can be identified by its slight yellowish tint (especially when aged), and by the fact that a scrap of it will burn with a sputtering flame and give off black smoke. It will also react, and soften, upon contact with some materials, such as acetone. Acrylic plastics are identified by such trade names as Perspex (UK) and Plexiglass (USA). It is stiffer than cellulose acetate, more transparent and practically colourless. Acrylic burns with a clear flame and gives off a fairly pleasant odour. Acetone, if applied, will cause white marks but will leave the material as hard as it previously was. 7.1.1.1 Use of Thermoplastics

Thermoplastics are, normally, used where there are no unusual temperature changes and the majority of all plastics production is thermoplastics, which include: • Acetate - widely used for tool handles, and electrical goods. • Poly-Ethylene - commonly known as polythene. Its uses include flexible tubing, cable insulation and packaging. • Poly-Propylene - stronger, harder and more rigid than polythene. Used for such items as high-pressure air piping. • Poly-Vinyl-Chloride - commonly known as PVC. Varying degrees of rigidity/flexibility are achievable by varying the amount of plasticiser used. Rigid, moulded sections or piping can be produced and also flexible electric cable insulation • Polystyrene - can be produced in rigid form, but is more familiar in the expanded form, when it is useful for thermal insulation, buoyancy or shockresistant packaging.


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• Acrylics - these are particularly useful where light transmission is necessary. Perspex and Plexiglas belong to this family. They have excellent light transmission properties and are also resistant to splintering. There is a tendency for some fine craze-cracking to develop if exposed for long periods to ultra violet light. These transparent plastics may be solid or laminated. When laminated two or more layers are bonded together with a clear adhesive and, in this form, they are more shatter-resistant and are ideally suited to pressurised aircraft windows. An even stronger and more shatterproof transparent plastic can be achieved by stretching the acrylic in both directions before final shaping. These improved properties, result from the stretching operation causing a preferential alignment of the long-chain molecules. Extreme care should be taken when handling acrylics, as they are they are easily scratched. The acrylics are supplied with a paper or rubberised film, which should not be removed, until required for use. If dirty, they should be cleaned with cold water or soapy water. Care should also be taken when using solvents in the vicinity of acrylics. Some solvents, or their vapours, may cause crazing of the material. , Reference to the appropriate Manuals or manufacturers’ specification sheets are essential. • Poly-Carbonates - these have similar uses to the acrylics (Perspex etc) but are more temperature-resistant and also have superior impact strength. They are also more expensive. • Nylon - belongs to the polyamide family and is an extremely useful and versatile material. It is strong, tough and also has low friction properties. It can be used as a fibre or produced as a moulding. Popular uses include textiles, furnishings, ropes, tyre reinforcement, bushes, pulleys, gears, and lightweight mouldings such as brackets, handles etc. • Poly-Tetra-Fluoro-Ethylene - commonly known as ‘PTFE’, it is similar to nylon in appearance but is denser, whiter and much more expensive. It has a wax-like surface and this characteristic results in very low friction properties, which make it suitable for bushes and gears. It also has a high temperature capability (over 300ºC) and is extensively used as a non-stick coating e.g. Teflon. PTFE tape is often used as a thread sealant for oxygen pipe threads, and as backing rings for hydraulic seals 7.1.2 THERMOSETTING MATERIALS

Thermosetting materials (also called Thermosets) will, initially, soften when heated, but will remain soft for only a short time and will set (and harden) if the heat continues to be applied. The process of Thermosets becoming hard, when heated, is called ‘curing’ and curing can also be achieved by chemical (exothermic) reactions.


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During the curing process, the long-chain molecules of the material cross-link (link together between chains) and, once the cross-links are formed, the plastic becomes hard and cannot be re-softened by heating. Thermosets are, thus, chosen where a plastic component will be exposed to relatively high temperatures, as some of them can tolerate temperatures in excess of 250°C before beginning to char. Note: Thermosetting materials are generally stronger, have a lower ductility and lower impact properties than the Thermoplastics. 7.1.3 RESINS

Natural resins are obtained from the exudations from certain trees and other plants and as clear, translucent, yellow (amber), brown, solid, or semisolid agents, they are used in inks, lacquers, linoleum, varnishes and, of course, plastics. While the words plastics and resins are often used synonymously, they are, in fact, quite different, in that plastics refers to the material in the finished items while resins are the raw materials which may be found in the form of flakes, pellets, powder, or a syrup. Resins may be used alone to form plastics but, usually, additives are employed with them, to assist in the moulding characteristics, or to enhance the properties of the finished product. The resin may be thickened and given more ‘body’ by the addition of inert fillers, which may be used to fill gaps and voids in the structure. Typical fillers are microballoons, cotton and glass flock and aerosil (fumed silica). Reinforcing agents, plasticizers, stabilisers, colorants, flame-retardants, smoke suppressants and processing aids, such as lubricants and coupling agents, are among the other additives used with resins. Resins have little strength in themselves and are generally used to impregnate linen, paper, and ‘cloths’ made up from various synthetic fibres. For many years, aircraft control cable pulleys have been made from thermosetting resins, reinforced with layers of linen cloth. These pulleys are cured in a mould, at high temperature, and have high strength without causing wear to the control cables. When layers of paper are impregnated with a thermosetting resin such as phenolformaldehyde or urea-formaldehyde, they can be moulded into flat sheets or other shapes. Once hardened, the material makes an exceptional electrical insulator and can be found in use as terminal strips and printed circuit boards.


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Polyester Resin

Polyester resin can be extruded into fine filaments and woven into fabric (like nylon) or cast into shape and it is also useful as a heat-resistant lacquer. Glass fibres and mat, for example, have great strength for their weight, but lack rigidity so, to convert glass fibre into a useful structural material, it is impregnated with polyester resin and moulded into a desired form. Polyesters cure by chemical action, and, so, differ from materials, which cure by the evaporation of an oil or solvent. As polyester is thick and unmanageable, a styrene monomer is added to make it thinner and easier to work. If left alone, the mixture of polyester and styrene will, eventually, cure into a solid mass, so inhibitors are added to delay this curing process and to improve shelf life. A catalyst then has to be used, when the inhibitors are no longer wanted and the curing process is to be started and an accelerator will appreciably shorten the curing time of the resin, depending on the temperature and mass of the resin. The actual cure of polyester resin occurs when a chemical reaction between the catalyst and accelerator generates heat within the resin. This (exothermic reaction can be seen when a thick layer cures more rapidly than a thin layer. 7.1.3.2

Thixotropic Agents

The heat, generated by the chemical reaction, can make the material less viscous and cause it to ‘run’ (particularly if it is on a vertical surface). To overcome this problem, a thixotropic agent is added to the resin after mixing, to increase its viscosity. The increased viscosity allows the resin to remain in place no matter where it may be used. 7.1.3.3

Epoxy Resin

Another type of resin that can be used in place of polyester in laminated structures is epoxy resin. Epoxy resin has a low percentage of shrinkage, high strength for its weight and the ability to adhere to a wide range of materials Unlike polyester resins, that require a catalyst, epoxy resins require a hardener or curing agent without recourse to heating. There is also a difference in the mixing ratios between polyester and epoxy resins. For polyester resin, the ratio is 64:1, resin to catalyst whilst, for epoxy resin, the ratio is 4:1, resin to hardener.


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7.1.4 ELASTOMERS

From the Greek word elastos – elastic, elastomers may be natural or, synthetic materials (polymers) which have considerable elastic properties. Because they may also be moulded into shapes, which they retain, they qualify to be included in the category of plastics. Elastomers will tolerate repeated elongation and return to their original size and shape, in a similar way to natural rubber Some of the more common elastomers, to be found in the aerospace industry include: • Buna ‘N’ - also known as Nitrile. A synthetic rubber, made (initially in Germany) by the polymerisation of butadeine and sodium (hence BuNa), it has excellent resistance to fuels and oils, and is used for oil and fuel hoses, gaskets, and seals. This material also has low ‘stiction’ properties, when in contact with metal, and is, therefore, particularly suited to ‘moving-seal’ applications. • Buna - ‘S’ relatively cheap material, also with a performance similar to natural rubber. It is often used for tyres and tubes, but its poor resistance to fuels/oils/cleaning fluids makes it unsuitable for seals. • Fluoro-Elastomers - these have exceptional high-temperature properties and can be used at 250ºC. They are also solvent-resistant and are mainly used for high-temperature seals. A common name for these materials is Viton. These materials are expensive. • Neoprene - has very good tensile properties and excellent elastic recovery qualities. It is also solvent-resistant and, therefore, has a wide range of applications as fuel and hydraulic seals and gaskets. However, because of its special elastic recovery properties, it is also ideally suited to diaphragms and hydraulic seals. • Poly-Sulphide Rubber - although it possesses relatively poor physical properties, it has exceptionally high resistance to fuels and oils and is widely used for lining or sealing fuel tanks. It is also used for lightly stressed seals and hoses, which come into contact with fuels or oils. This compound is commonly known under the trade names of PRC or Thiokol. • Silicone Rubber - has very good high- and low-temperature properties (-80ºC to + 200ºC). It is often used for seals, but is also used for the potting of electrical circuits, because of its ability to retain its rubbery state, even at low temperatures.


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7.2 PRIMARY ADVANTAGES OF PLASTICS Plastics are being used on an ever-increasing scale and are frequently replacing some of the more conventional materials such as metals, wood and natural rubbers. Plastics have properties, which make them a popular choice over conventional aircraft materials. Some of the more important characteristics of plastics, which help to explain their popularity, are: • Lightness - most plastics have specific gravities of 1.1 to 1.6 whereas the more common engineering materials, such as aluminium and steel, have values of 2.7 and 7.8 respectively. • Corrosion Resistance - plastics will tolerate hostile corrosion environments and many of them resist acid attack. • Low Thermal Conductivity - this property makes many plastics ideal for thermal insulators. • Electrical Resistance - plastics are used in enormous quantities for electrical insulation applications. • Formability - many plastics are easily formed into the finished product, by casting moulding or extrusion, often in a single operation. • Surface Finish - excellent surface finishes can be achieved in the basic forming operation, so finishing operations are not necessary. • Relatively Low Cost – because, although some of the materials may not be particularly cheap, the lack of machining necessary and the high production rates possible, keeps the costs down. • Light Transmission - some plastics are naturally clear, whilst other are opaque. These characteristics, consequently, provide the possibility for a range of light-transmission properties. Optical properties can also be achieved with some plastics. • Vibration Damping - many plastics are naturally resistant to fatigue and, because of the high value of internal damping present, resonances will tend to be of relatively low amplitude. 7.3 PRIMARY DISADVANTAGES OF PLASTICS Although plastics are extremely useful materials, some shortcomings inevitably exist, particularly when compared to some metals. Plastics major deficiencies are: • Lack of Strength - most plastics are much weaker than metals and mild steel has approximately six times the strength of nylon. Mild steel, however, is six times the weight of nylon so, on a strength/weight ratio, they are comparable.


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• Low Stiffness - plastics have a very inferior value of Young’s Modulus compared with the common metals. • Low Impact Strength - many plastics have poor impact strength, but there are a few exceptions, such as with certain polycarbonates. • Poor Dimensional Stability - mainly due to high values of thermal coefficient of expansion. • Poor High-Temperature Capability - metals are generally capable of retaining reasonable strength at much higher temperatures than the plastics. The long-term maximum operating temperature, for the better plastics, is not usually above 250ºC. High-temperature metals can operate for long periods well in excess of 800ºC. • Moisture Absorption - many types of plastic absorb moisture, which can result in a significant loss of strength in a humid environment. • Ultra Violet Light - some plastics deteriorate when exposed to UV light for long periods. Increased brittleness and loss of strength can occur. 7.4 PLASTIC MANUFACTURING PROCESSES The most common manufacturing methods are: • Casting - the molten material is simply poured into a mould and allowed to set. • Moulding - powder, liquid or paste is forced into a set of shaped dies. • Extrusion - plastic is forced through a suitably shaped die. Rod, sheet, tube, angle sections etc. are produced this way. • Lay-up - load-carrying plastic fibres and an adhesive are layered in a mould or around a former. • Sandwich-Construction - plastic facings have, sandwiched between them, a honeycomb or foam core. Very stiff, but light, structures are achieved by this method. • Compression Moulding – the material is put into a heated, hardened, polished steel container (the die) and forced into shape, by a plunger. • Note: Vacuum Forming uses a similar tooling but, in this instance, the plastic is sucked into contact with the shaped die (a method often used to manufacture aircraft interior trim).


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7.5 COMPOSITE MATERIALS As previously stated with Plastics, the main reason for utilising composite materials, in aerospace structures, is to reduce weight, which has a direct benefit in lowering operating costs. Composites also provide further benefits in their ability to be easily formed, comparatively lower production costs, resistance to corrosion and reduced maintenance costs. The principal types of composite materials are those involving fibrous elements which may be used as strands, or be woven into fine ‘tapes’ and ‘cloths’ (or coarser ‘mats’), held in a suitable resin matrix and formed into the required shapes 7.5.1 GLASS FIBRE REINFORCED PLASTIC (GFRP)

The first man-made fibre, glass can be spun into cloth and used for fire-proof curtains or (when extremely pure glass is used), made into fibres which are able to transmit light over long distances. The ultimate tensile strength of undamaged, very small diameter glass fibres is extremely high, although the strength is reduced significantly if the fibres are slightly damaged. In its structural use it is often merely referred to as glass fibre or fibreglass, when glass fibres (in various forms) are bonded together by appropriate resins. When moulded with resin, the resulting composite is, also, of considerably lower strength but, nevertheless, good GFRP structures are stronger than mild steel and, on a simple strength-for-weight basis, can be comparable to high tensile steel if the fibre form and lay-up is near optimum. It is however, considerably less stiff than steel or even aluminium. A graphic example of GFRP flexibility is the enormous deflection, which takes place in the pole during a pole vault. As the glass fibres are about a hundred times stronger than the resin, it is obviously necessary to get as much fibre packed into the moulding as possible. Non-structural items may be made from, or include, a percentage of chopped strand mat, (i.e. glass fibres in a random, non- woven state) but, where considerable strength is required, uni-directional glass cloth is used. To provide all round strength, sheets of uni-directional cloth can be layed up at 90º to each other, in a similar manner to the grain in plywood. Sometimes such sheets are used as facings for an internal honeycomb of plastic-impregnated paper, to give a very efficient structure in terms of strength, stiffness and weight.


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The glass fibre sheet material can be supplied with cloth already impregnated with resin and partially cured (‘Pre-preg’), in which case it is necessary to keep the material in refrigerated storage. Resin curing is usually done at elevated temperatures (120°C - 170ºC), with the GRP component in its mould and, often, under pressure, in an autoclave. The main reasons for using GFRP are: • in instances where metal cannot be used (e.g. for radar domes or other nonelectrical conducting applications) • the ease and low cost of producing very complex shapes • to provide good strength/weight ratio • its ability to produce selected directional strength. The main disadvantage of glass fibre is that it lacks stiffness and, as such, is not suitable for applications subject to high structural loadings. 7.5.1.1

Ceramic Fibres

Made by firing clay or other non-metallic materials, ceramic fibres are a form of glass fibre, used in high-temperature applications. They can be used at temperatures up to 1650°C and are suited for use around engine and exhaust systems. Ceramic fibres are heavy (and expensive) and are only used where no other materials are suitable. 7.5.2 CARBON FIBRE REINFORCED PLASTIC (CFRP)

CFRP (also referred to as ‘Graphite’) is a composite material, which was primarily developed to retain (or improve upon) the high strength-to-weight ratio characteristics exhibited by GFRP, but with very much greater stiffness values. Carbon fibres are very stiff and, when formed into a composite, the Young's Modulus (‘E’) value can be higher than steel. CFRP is not only six times stiffer than GFRP but is also over 50% stronger. It also has twice the strength of highstrength aluminium alloy and three times the stiffness. Carbon fibres are typically less than 0.01 mm (0.0004 in) in diameter and are produced by subjecting a fine thread of a suitable nylon-type plastic to a very high temperature (to decompose the polymer), and driving off all of the elements with the exception of carbon. The carbon thread is then stretched, at white heat (2000°C-3000ºC), to develop strength. Unfortunately, the process is complex and very costly. Nevertheless, where the high cost can be justified, CFRP can offer considerable weight savings over conventional materials. CFRP components are generally made from ‘Pre-preg’ sheet (fibres impregnated with resin and a hardener, which only require heat and pressure to cure). Some specialist items are made by a EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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laborious, but ideal, process called ‘Filament Winding’, in which a carbon fibre string is wound over a former in the shape of the workpiece whilst bonded with resin. Because of CFRP's high stiffness modulus, it is also used extensively to stiffen GFRP or aluminium alloy structures. A material known as Carbon-Carbon (where the resin is also graphitised), is used for the rotors and stators on brake units. It offers a significant weight saving, as well as high efficiency, due to the fact that it dissipates the heat generated very quickly. Replacing 40% of an aluminium alloy structure by CFRP would result in a 40% saving in total structural weight and CFRP is used on such items as the wings, horizontal (and vertical) stabilisers, forward fuselages and spoilers of many aircraft. The use of composites, in the manufacture of helicopter rotor blades, has led to significant increases in their life and, in some cases, they may have an unlimited life span (subject to damage). The modern blade is highly complex and may be comprised of CFRP, GFRP, stainless steel, a honeycomb core and a foam filling. 7.5.3 ARAMID FIBRE REINFORCED PLASTIC (AFRP)

The aramid fibres are closely related to the nylon-type of synthetic fibres and are well known for their superior toughness, strength-to-weight characteristics and heat-resistance. Tyres, reinforced with aramid fibres are comparable to those reinforced with steel cords. Better known under its trade name – Kevlar –in cloth form, it is a soft, yellow, organic fibre that is extremely light, strong and tough. Its great impact-resistance makes it useful in areas, which are liable to be struck by debris, as experienced in areas around engine reverse-thrust buckets. Kevlar is used to manufacture bulletproof jackets and, also, as a reinforcement, in aircraft fuel tanks. 7.5.4 GENERAL INFORMATION

A sheet of fibre reinforced material is ‘anisotropic’, - which means its properties depend on the direction of the fibres. Random direction fibres would result in a much lower strength than uni-directional fibres, laying parallel to the applied load. However, the strength (and stiffness) of a uni-directional lay-up would be very low, with the applied load at 90º to the fibres, as this is primarily a test of the resin (hence the usual practice of placing alternate layers at 90º to each other). Due to small variations in the size of the individual fibres, and the final quality of the finished component (which can be affected by careless handling, variations in cleanliness or lay-up, voids, pressures, temperatures, etc), there will, inevitably, be a greater scatter on final strength than on a conventional, metallic component. Due allowance on stress reserve factors is, therefore, essential. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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It has already been stated that composites usually have good internal damping characteristics and are less prone to vibration resonances. Where high strength, combined with stiffness is required, then a CFRP is used but, when lesser levels of stiffness are necessary, then GFRP or AFRP are used. Composites have very low elongation properties and toughness. Aluminium alloy has a typical elongation-to-fracture value of 11%, whereas composites range from 3% for GFRP to 0.5% for CFRP. The maximum operating temperatures, for GFRP, CFRP and Kevlar composites, depend, to some extent, on the actual adhesives used, but are, generally, in the range 220°C-250ºC. Some composites, such as carbon fibre in a carbon matrix, have very high permissible operating temperatures (around 3000ºC), and are used for highenergy braking applications and as thermal barriers for space vehicles). Boron, Tungsten, Silicon Carbide and Quartz may also be used to provide fibres for high-temperature composites 7.5.5 LAMINATED, SANDWICH AND MONOLITHIC STRUCTURES

Laminated plastics consist of layers of synthetic resin-impregnated fibres (or other, coated, fillers), which are bonded together (usually heated and under pressure), to form a single laminate or sheet of composite material. Plastic laminates are used to ‘face’ other structural materials, in order to; • • • •

provide a more durable surface to a softer (less expensive) material enhance the surface appearance (colour, porosity, smoothness etc.) increase the strength and rigidity of many non-metallic structures produce other desirable surface characteristics such as when acid- or corrosion- resistance, non-conductivity, non-magnetisability or the ease of keeping a surface clean is required

To provide a light-weight structure, which possesses strength and rigidity, one of several structural materials, is sandwiched between two laminated composites. The sandwiched material (the core) may be made of a solid material, such as wood, or a series of thin corrugations of a material, which are joined and placed end-on (in the form of the cells of a honeycomb), within the laminates. Where wood is used, as the core material, it usually consists of low-density balsa wood, which has been cut across the grain and sandwiched between two layers of reinforced resin (or a metal). This construction makes an extremely light, yet strong material, which can be used as floor panels, wall panels and, occasionally, aircraft skins. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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The cellular core, used for laminated honeycomb material, may be made from resin-impregnated paper, or from one of the many fibre cloths. The core is formed or shaped and then bonded between two face sheets of resin-impregnated cloth. The finished sandwich structure is very rigid, has a high strength-to-weight ratio, and is transparent to electromagnetic (radar/radio) waves, making it ideal for radomes of all kinds. Metal honeycomb cores (made from light alloy or stainless steel), are also sandwiched between two face sheets of fibre-reinforced resins. On other occasions the metal honeycombs may be found sandwiched between sheets of light alloy, stainless steel or titanium. This type of core is referred to as ‘metalfaced honeycomb’ and is used where abrasion- and heat-resistance is important or when sound-absorption qualities are desired. In monolithic structures, angle sections (‘Top Hat’, ‘U’, ‘I’ and ‘Z’), frames ribs and stringers are fashioned from similar materials to the outer layers of the sandwich structure, then covered with the appropriate surface ‘skin’, before the stronger, metallic spars and hinges are attached, Such a structure can save many kilograms (or pounds) in the weight of the flying control surfaces (or the fin structure) of a large aircraft. 7.6 NON-METALLIC COMPONENTS In addition to the non-metallic materials, used in the aircraft structure, nonmetallic materials are used in many aircraft components and systems. Many of these materials require specialist knowledge and understanding, during aircraft maintenance. 7.6.1 SEALS

Seals or packing rings (refer to Fig. 18) serve to retain fluids and gases, within their respective systems, as well as to exclude air, moisture and contaminants. They also have to withstand a wide range of temperatures and pressures and, because of this, they have to be manufactured in a variety of shapes and materials. The most common materials, from which seals are manufactured, are natural rubber, synthetic rubber and Teflon (trade name for polytetrafluoroethane or PTFE). O-ring seals effectively seal in both directions of movement. They are used to prevent both internal and external leakage, and are the most commonly used seals in aviation


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Examples of Seals and Packing Rings Fig. 18

Where installations operate at pressures above 10.34 x 10³ kN/m² (1500 psi), additional back-up rings can be used to prevent the O-ring from being forced out or extruded. These back-up rings are usually made from Teflon, which does not deteriorate with age, is unaffected by system fluids and vapours and tolerates temperatures well in excess of those found in high-pressure hydraulic systems. O-rings are available in many different materials and sizes (both diameter and thickness). They are supplied in individual, hermetically-sealed, envelopes with all the necessary information marked on the packaging. This system has generally replaced the previously used, colour-coding of seals, which had severe limitations. For applications (such as in actuators) that subject a seal to pressure from two sides, two back-up rings can be used but, when the pressure is from one side only, a single back-up ring is adequate. Other seals, commonly found are V-ring and U-ring seals. The V-ring has an open ‘V’ facing the pressure and is located by the use of a male and female adapter. The U-ring seals will, usually, be found in brake unit assemblies and master cylinders, where pressures below 89 x 10³ kN/m² (1000 psi) are encountered. As they only seal in one direction, the concave surface must face towards the pressure.


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8 DETECTING DEFECTS IN COMPOSITE MATERIALS While composites do not suffer the corrosion and cracking problems, associated with metals and also have good fatigue characteristics, they do, however, require regular inspection for the defects to which they are particularly prone. The areas to be inspected are, usually well known and they will be detailed in the relevant chapter (51-57 for Airframe topics, 61-61 for Propellers) of the Aircraft Maintenance Manual (AMM). The inspection methods to be used will be found in the Non-destructive Testing Manual (NTM) and the approved repair procedures will be outlined in the Structural Repair Manual (SRM). Repairs in unexpected areas, or damage, which is not covered in the SRM, will necessitate the request of specific repair drawings from the aircraft manufacturer. 8.1 CAUSES OF DAMAGE If a sharp object strikes a thermosetting plastic, the plastic is liable to crack and shatter, like glass, with straight sharp edges. The reason for this is that, once a crack starts in the plastic, it travels very easily and quickly in a straight line. Damage of this kind would be disastrous in a load-bearing component. The damage appears as a ‘star’ in the composite, providing it has no surface finish applied to it. An important point about this type of damage is that there is little loss of strength in the overall material, in addition to the absence of the shattering that occurs without fibre reinforcement. The majority of damage to composite structures occurs during ground handling (such as from dropped tools), and damage from ground equipment. Bird-strike damage can also require extensive repairs. Damage to composite structures may result from a number of other causes such as: • Erosion caused by rain, hail, dust etc. • Fire • Overload caused by heavy landings, flight through turbulent air and excessive ‘g’ loading. • Lightning strikes and static discharge. • Chafing against internal fittings such as pipes and cables. 8.2 TYPES OF DAMAGE The types of damage, which may affect fibre-reinforced structures are: • Cracks which may simply affect the outer lamination or may penetrate through the skin. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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Fibre reinforced plastics however, apart from being much stronger than normal plastics, have different failure modes. Each strand of fibre acts as a trap, to stop cracks travelling through the plastic (refer to Fig. 19). A travelling crack quickly reaches a fibre, which is difficult to break so, instead, the crack travels along the fibre. Eventually the crack reaches another fibre and is deflected again. This process continues until the failure is divided into many small cracks, which will not have propagated far from the initial damage.

Fibre

(a) Crack travelling towards fibre

(b) Crack travelling along fibre

(c) Cracks around fibre

Crack Propagation within a Composite Fig. 19 • Delamination - which involves separation of the fibreglass layers and may affect single or multiple layers. • Debonding - when honeycomb sandwich structures are damaged, the effect usually entails separation of the honeycomb from the skin. The reason for this is that the bonding of the skin to the honeycomb walls is along very fine lines, and this bond is fairly easily broken. Once there is separation, the strength of the whole structure is reduced by a significant amount. Greater damage can be due to the crushing of the honeycomb core itself, which may require extensive repair or even replacement of the complete component. • Blisters - which usually indicate a breakdown in the bond within the outer laminations and may be caused by moisture penetration through a small hole, or by poor initial bonding • Holes - these may range from small pits, affecting one or two outer layers, to holes, which completely penetrate the component. Holes may be caused by lightning strikes or by static discharge. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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8.3 INSPECTION METHODS Areas on the aircraft that are likely to be damaged, should be inspected regularly, and complete removal of the component may be required at overhaul. 8.3.1 VISUAL INSPECTION

Visual inspection is used to detect cracks, surface irregularities (from an internal flaw) and surface defects such as delamination and blistering. A lamp and a magnifying glass are useful in detecting cracked or broken fibres. A small microscope or a x20 magnifier may be helpful in determining whether the fibres in a cracked surface are broken, or if the cracks affect only the resin. Delamination may sometimes be found by visual inspection. If the area is examined at an angle, with a bright light illuminating the surface, the delaminated area may appear to be a bubble, or an indentation in the surface. When viewed from the inside, a change of colour could indicate delamination because of a change in light reflection. A visual inspection can also find several manufacturing defects such as resin-rich or resin-starved areas, pinholes, blisters and air bubbles. 8.3.2 RING OR PERCUSSION TEST

To detect internal flaws, or areas suspected of delaminations, a ring or percussion, test can be used. In some instances a properly designed miniature hammer is used for the test while, in other procedures, a length of an appropriate hardwood, or a particular size of coin is employed to tap against the surface of the suspect area. Variations in the tapping sound will provide clues as to the quality of the bond. A sharp solid sound indicates a good bond, whilst a dull thud indicates bond separation. Care must be taken to make allowances for changes in material thickness, fasteners and earlier repairs, all of which can give false indications. Whenever damage is found visually, then a percussion test should be done around the area. In the majority of instances, if there is a hole, crack or other damage, there is, often, also delamination. 8.3.3 ULTRASONIC INSPECTION

To detect internal damage, an ultrasonic test may be done by authorised, specialist, personnel. This procedure involves the directing of a low-frequency ultrasonic beam through the structure and viewing the pattern of the resulting sound echo on an oscilloscope. 8.3.4 RADIOGRAPHY

Radiography can, sometimes, be used to detect cracks in the surface in addition to being able locate internal faults that cannot be visually detected. Radiographic procedures may also be employed to detect water ingress within honeycomb core cells. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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8.4 ASSESSMENT OF DAMAGE One of the greatest problems, caused by replacing aluminium alloy with composite structures (especially honeycomb sandwich), is the inspection for damage. It is unfortunate that when a composite of any kind is struck, the majority of the damage occurs internally and, often, there is little or no visible damage showing at the surface It is vital that ANY damage to a composite structure be thoroughly inspected, not only for damage to its surfaces, but also (in a sandwich structure) for possible damage to its core, which is usually softer than the skins. Damage that gives little clue to its depth or significance is often referred to as Barely Visible Damage (BVD). As with metal structures, the damage occurring to GFRP or CFRP structures may be classified as negligible (or allowable), repairable by cover patch, repairable by insertion or repairable by replacement. These classifications may only be determined by reference to the appropriate aircraft SRM. Signs of secondary damage (i.e. damage occurring remote from the primary damage) must not be overlooked. This is particularly important in the case of impact damage where the shock may be transmitted through the structure, to cause damage away from the point of impact. In some instances secondary damage may be more serious than the primary damage. Sometimes damage may be difficult to detect, due to the natural flexibility of the material which may cause it to spring back into shape. Any evidence of cracking, straining, crazing or scuffing of the gel coat should be regarded with suspicion, as it may indicate the presence of damage. Where delamination is known, or suspected to exist, the area surrounding the visible damage should be checked to determine the extent of the damage and integrity of the laminations.


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9 BASIC COMPOSITE REPAIRS WARNING: THE CHEMICALS, USED IN LAMINATING RESINS AND CLEANING AGENTS, ARE HAZARDOUS SUBSTANCES AND EXTREME CARE IS CALLED FOR WHEN HANDLING THEM. MOST RESINS ARE AN IRRITANT TO THE SKIN. MANY PEOPLE ARE ALLERGIC TO THE RESIN, AND REPEATED SKIN CONTACT CAN CAUSE SERIOUS DAMAGE. IF SYMPTOMS OF AN ALLERGY APPEAR WHEN THE RESIN IS USED, FURTHER CONTACT SHOULD BE AVOIDED AND THE SYMPTOMS SHOULD SLOWLY FADE AWAY. DIRECT SKIN CONTACT WITH THE RESIN SHOULD BE AVOIDED, AND RUBBER OR PLASTIC GLOVES WORN WHEN THERE IS A POSSIBILITY OF THE HANDS BECOMING CONTAMINATED. THE RESINS AND SOLVENTS, USED WITH SYNTHETIC FIBRES, ARE ALL POISONOUS. EVERY PRECAUTION SHOULD BE TAKEN TO KEEP THEM AWAY FROM FOOD. THE FACE, AND ESPECIALLY THE EYES, SHOULD ALSO BE PROTECTED FROM RESIN AND ITS SOLVENTS. IF A ROTARY GRINDER IS USED ON A GLASS FIBRE LAMINATE, MUCH GLASS AND RESIN DUST WILL BE PRODUCED AND A RESPIRATORY MASK SHOULD BE WORN FOR PROTECTION. THE SAME DUST IS LIKELY TO CAUSE AN IRRITABLE SKIN RASH TO DEVELOP ON THE FOREARMS, ESPECIALLY WHEN GLASS FIBRE IS BEING HANDSANDED. BEFORE WASHING HANDS AND ARMS, AFTER WORKING WITH GFRP, IT IS ADVISABLE TO RINSE THEM IN COLD WATER. THE ARMS SHOULD BE WASHED IN SOAPY WATER AND THE OPERATOR SHOULD AVOID SCRATCHING, ESPECIALLY WHILE DUST IS LYING ON THE SKIN. Before commencing repairs on any composite material (whether it be a simple ‘fibreglass’ skin or a complex honeycomb sandwich), the complete area of the damage must be carefully surveyed. This must be done in accordance with the AMM, to ensure that ALL damage is discovered and assessed. Any subsequent repair will depend on the type of damage, the extent of that damage and the importance (significance), to the safety of the aircraft, of the material being repaired. The AMM will provide either a repair scheme or component replacement information.


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The strength of a glass fibre repair is dependent on the strength of the bond to the original structure. Since the repair receives its working loads through this bond, it is imperative that every effort is made to ensure a sound connection. Some of the important considerations are: • Correct Surface Preparation • Correct Bond Strength - this requires correct procedures to be used during the repair process • Uniform Stress: - once again correct procedures during repair will ensure that local stress concentrations are minimised. 9.1 REPAIR OF A SIMPLE COMPOSITE PANEL If a heavy object has been dropped onto a glass-fibre-reinforced epoxy panel, the damage could consist of a small hole surrounded by damaged composite (refer to Fig. 20). Point of Impact Damaged Area Front of Panel

Rear of Panel Damage to Composite Fig. 20 The damaged material is removed first (refer to Fig. 21), bearing in mind that the damage may be small on the front, but may extend some distance at the rear side of the panel.

Undamaged Panel Damaged Area Removed Fig. 21


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Once the hole has been cleaned out, degreased and its surfaces roughened up, a piece of glass cloth is laid over the hole, followed by several other pieces, each on top of the previous piece. In this way, the hole is filled with successive layers and completed with several large layers over the final surface (refer to Fig. 22). The SRM will give the exact procedures for each repair.

Repair to Composite Fig. 22 9.2 REPAIR OF A SANDWICH PANEL These repairs are considered to be more difficult than composite panels, due to their complexity, and require skilled personnel. In this example, the assumption is that the dropped tool has broken the skin and damaged the core. As previously stated, the first task is to remove the damaged material, usually with a router (refer to Fig. 23).

Plug of Damaged Material Removed

Damage Removed (with a Router) Fig. 23


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A ‘plug’ of honeycomb is cut to the correct dimensions, without gaps, and bonded into the hole (refer to Fig. 24).

Honeycomb Plug

Core

Damage Plugged Fig. 24

Once the plug is bonded in place, the upper skin can be repaired in much the same way as with the composite panel. Several layers of mat are then bonded carefully onto both the original surface and the plug (refer to Fig. 25).

Completed Repair of Damaged Area Fig. 25 Note: The above examples are only an outline of the full repairs that may be done, during aircraft maintenance.


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9.3 GLASS FIBRE REINFORCED COMPOSITE REPAIRS These course notes deal exclusively with repairs to glass fibre reinforced composites, including honeycomb-cored structures. To summarise, glass fibre composites have two basic constituents, namely the glass fibre and the surrounding plastic matrix. The glass fibres reinforce the plastic matrix and carry most of the load. The matrix gives the composite its rigidity and protects the fibres from attack by moisture or chemicals. Glass fibres are generally woven into a fabric, which gives a regular orientation to the fibres and allows them to be handled more easily. To produce a glass fibre laminate, successive layers of the fabric are placed into position and impregnated with resin. The liquid resin solidifies within a few hours and after post curing at elevated temperatures, forms a strong matrix around the fibres. Using this technique, intricate shapes can easily be formed with the load carrying filaments orientated in the best possible manner. It is also possible to reinforce the laminate locally and to mould in load bearing fittings etc. into the laminate. 9.4 TYPES OF GLASS REINFORCEMENT After production of the basic glass fibres, they are collected together to form a collection of continuous, parallel fibres known as a roving. Glass fibre cloth is made by weaving rovings together. Depending on the closeness of the weave, and the number of rovings in each weave of the fabric, different weights of cloth may be produced. There are two main types of glass cloth, uni-directional and bi-directional. 9.4.1 UNI-DIRECTIONAL CLOTH

A uni-directional glass cloth has the majority of the glass fibres lying parallel and in one direction, with only enough transverse fibres to hold the fabric together. Roving may also be used either individually, or grouped together, to give a fully, uni-directional composite.


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9.4.2 BI-DIRECTIONAL CLOTH

A bi-directional cloth has the same number of roving in both warp and weft directions and, as such, can take stresses in both directions. There are two main types of bi-directional cloth: • Plain weave is woven with an ‘over one and under one’ configuration and is used for most flat surfaces. • Twill weave, has a weave with an ‘over one and under two’ configuration. This gives drapeability and is used where curved component shapes are required. 9.4.3

CHOPPED STRAND MAT

Chopped strand mat has random short fibres, lightly held together with a binder. A laminate of this material is heavy and of low strength, compared with one which is made of woven fabric. As a result, it is of little use in aircraft construction. 9.4.4 RESIN

The choice of resin for a particular application, is most important, because resins are produced with the necessary properties to suit only certain requirements and are, therefore, not suitable for universal application. Some resins are supplied as a three-part mix, consisting of resin (adhesive), accelerator and catalyst. It is vitally important, when mixing this type of resin, that the accelerator is never mixed with a free catalyst, otherwise an explosion may occur. The correct mixing procedure must be followed so that the resin and catalyst must be mixed together before adding the accelerator Most laminating resin comes in two-liquid parts, namely a resin and a hardener. Once hardener is mixed with the basic resin a chemical reaction begins and the mixture begins to solidify (cure). .Resin Mixing In any resin mix, the proportions are absolutely critical, since the cured strength depends on it. The proportions are normally specified by weight of the quantity of resin required. An excess of hardener in the mixed resin is as damaging as a deficit. In both cases the cured resin will have an incomplete molecular structure and result in poor physical properties. Scrupulous cleanliness is essential in the mixing process, which should be performed in a warm, dry atmosphere in a well-ventilated and dust-free room. The materials should be measured in clean glass, or non-absorbent cardboard, containers.


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9.5 POT LIFE The temperature of the resin mix affects the rate at which the curing reaction occurs. If the temperature is too low the resin will be too thick to work, whereas if the temperature is too high, the resin will be comparatively thin and will drain out of the laminate before solidification occurs. Ambient temperature and humidity requirements are specified by the resin manufacturer. The length of time before a mix of activated resin begins to solidify is called ‘pot life’ and is dependent on the temperature and quantity of resin. Once the resin becomes thick and stringy, the curing process is well on its way. Resin in this state should not be used ,since the cured strength properties will be seriously degraded. To prevent waste, only sufficient resin should be mixed for the task in hand. 9.6 CURING Most resins used in aircraft structures will cure at standard room temperature (20ºC) but may take several days to reach a fully cured state. Once the resin has hardened, post-curing, at elevated temperature, is required for the resin to gain its full strength. For repair purposes the heat is usually applied by means of an infra-red lamp or electric heater. For components, which have been removed from the aircraft, an oven of suitable size may be used, to allow accurate control of temperature. If a large enough oven is not available, then a hot-air ‘tent’ should be constructed around the repair, and a thermometer used, to measure the average temperature inside the tent. Temperature may also be controlled by use of a temperature-indicating lacquer or pencil. These, when applied adjacent to a repair, will melt or change colour when a pre-determined temperature has been reached. The times and temperatures, required to effect a cure, are specified in the relevant SRM. The maximum curing temperature must not be exceeded. A typical time and temperature would be 8 hours at 60ºC. The use of pressure is normally specified for a repair whilst it is being cured. This assists in maintaining the correct profile of the repair and improves the bond. Pressure may be applied by clamps, weights or by a vacuum bag. Once the resin has cured, it is absolutely neutral. It will not swell or shrink with changes in climate and is only attacked by a few chemicals.


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9.7 GEL COAT The durability and appearance of a glass fibre moulding is dependent on its exposed surface. The purpose of the gel coat is to provide a resin-rich covering of the exposed surface of the laminate. This prevents the outermost glass fibres of the laminate from becoming exposed to attack by moisture and sunlight. If the gel coat is pigmented, then a solid coloured surface is also given to the laminate. Generally, the gel coat surface is incorporated in the moulding process, but it may also be used as paint and, after curing, it can polished to give a smooth, glossy surface. 9.8 STORAGE OF GFRP MATERIALS GFRP materials are expensive and, to ensure maximum shelf life, they should be stored in proper conditions. 9.8.1 STORING RESIN

Most laminating resins have a limited shelf life, which is specified by the manufacturer. In general, they should be stored in airtight tins at a cool temperature (usually below 10ºC). The resin should be removed from storage at least 24 hours before use, to allow it to assume workshop temperature. Depending on the type of resin, the shelf life may be up to 12 months, after which time it must be discarded. Resins, which have absorbed moisture, and become cloudy, should normally be discarded, but they can sometimes be recovered by heating them to 120ºC, to evaporate the moisture. If the resin clears on cooling, it may be used but, if it remains cloudy, it must be rejected. 9.8.2 STORING HARDENER

Hardeners generally react with oxygen in the air and must be stored in airtight containers. Some hardeners may crystallise if they become cold. To liquefy the hardener it should be gently warmed and then allowed to cool at room temperature. Note: The catalyst and accelerator, of a three-part laminating resin, should be stored separately to avoid inadvertent contact. 9.8.3 STORING FABRICS

Glass fabric should be stored in a warm, dry atmosphere, free from dust, oil or other contaminants. In order to preserve the fibre surface treatment it must no get damp. Before use, it is recommended that the fabric is heated to 45ºC in an oven, to drive off any moisture that may be in the fabric. Pre-preg fabrics should be stored in refrigerated conditions and all fabrics should be stored in their original wrappings. . EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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9.9 PREPARATION FOR REPAIR When the damage has been assessed as repairable, preparatory steps may be taken which are common to most types of repair. • The gel coat should be removed by grinding, or by gently chiselling and peeling it away, to determine whether the glass fibres are damaged. Signs of overstraining of the structure will show up as white cracks in the laminations. If the rear of the structure is accessible, a strong light, shone through the laminates, will show up any damage (delamination or cracks) as a dark area. The affected area should be cut out and the damage treated as a hole. • The damaged area should be cleaned and then cut back until sound material is reached. No evidence of whitening or cracking must be allowed to remain. Note: Before cutting out the damage, the area should be marked in some way, to determine its orientation for future reference (refer to Fig. 26). Any control linkages, bearings etc. should be covered to keep out glass dust and surplus resin. \\\\\\\\\\\\\\\\\\\ \\\\\\\\\\\\\\\\\\\ \ \ Sound Material \ \ \ Repair Area \ \ \\\\\\\\\\\\\\\\\\\ \

**

\ \ \ \ \ \ \ \

*

Orientation Mark

Orientation Marks on Repair Fig. 26

•

The type and number of glass cloth layers, used in the damaged area must now be determined. This may require the manufacturer to be consulted.

•

It is possible to analyse a sample of material, removed from the damaged area, by igniting one corner of the sample with a match or cigarette lighter. This burns off of the resin and allows individual fabric layers to be separated. The weight and direction of the fibre may now be determined and related to the parent laminate by reference to the previously applied orientation.


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• Notes should be made to ensure that the repair will be to the same specification as the original laminate (i.e. number, weight and direction of each layer). If the structure used a core material, the type and thickness should be noted. If the core is wood, the grain direction should be noted. •

The patch edges may now be prepared according to the particular repair being followed (scarf or stepped)and any surface that will have fibre bonded to it must have a thorough preparation (see the following paragraph, entitled ‘Surface Preparation’) When preparing a chamfered (scarfed) edge, the sanding direction should be towards the tip (refer to Fig. 27). The prepared edges should be examined for any sign of delamination, which must be removed by further sanding.

Sanding Direction

Direction of Sanding Fig. 27

Note: Some manufacturers specify that cut-outs should have radiused corners, while others permit square corners. • The inside of the structure should now be cleaned out and any loose pieces of glass fibre and accumulations of dust removed.


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9.9.1 SURFACE PREPARATION

The area, which is to be repaired, must be thoroughly degreased (using acetone or another, approved fluid). Once cleaned, the area should not be touched with bare hands. All paint, gel coat etc. must be removed from the repair area. The following is a typical example of the procedure that should then be adopted: • The repair area should be thoroughly abraded, using glass or garnet paper. The object of this abrasion is to remove the top film of resin from the glass and slightly roughen the glass fabric so that it becomes whiskery. Note: Care must be taken to ensure that not too much of the glass fabric is abraded. • Any dust must be removed with a clean cloth. • The newly exposed surface is thoroughly degreased, using a clean cloth saturated with the appropriate fluid (acetone). • The acetone must be allowed to evaporate from the surface. Careful use of a hot air blower is recommended to drive off any traces of acetone that may be trapped in the surface fibres. • Having cleaned the surface, the repair should commence as soon as possible. 9.10 TECHNIQUES OF LAMINATING GLASS FIBRE While the actual procedure will be detailed in the SRM, a typical list of the techniques to be adopted would include: • Pre-shaped templates are used to cut out the required pieces of cloth for the repair. • The workshop temperature must be between 15ºC and 23ºC with a relative humidity of not more than 65%. • The quantity of resin required should be estimated and mixed in the correct proportions of resin and hardener according to the manufacturer’s instructions. The container in which the resin is mixed must be clean and there must be no possibility of the container contaminating the contents (for this reason ‘unwaxed paper cartons’ are recommended). Note: If the resin is for structural repair work, a small sample (about 1cc) of mixed resin is cast in a container made from aluminium foil. The sample should be labelled and placed aside to cure for later inspection. • A coat of resin is brushed onto the prepared surface and the first layer of cloth is placed on the resin. The cloth is stippled into the resin, ensuring that the cloth weave pattern is not disturbed and that all the air bubbles are worked out.


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• The brush used for stippling should be slightly wet with resin which will allow the cloth to ‘wet out’ more quickly and help to prevent the cloth sticking on the brush. Note. Beware of using too much resin as this will result in a resin-rich and heavy repair. Ideally there should be just enough resin in a laminate to wet out the cloth. The fibres, when correctly wetted out, are almost invisible. • The edges of the cloth are trimmed, to ensure that the repair only covers the correct area. This is done, by lifting the edge of the patch and removing the excess with a sharp pair of scissors. • Each subsequent layer of cloth is then positioned and stippled into the preceding layers (trimming as necessary) until the laminate is complete. • When laminating is complete, the repair must be allowed to cure without any further disturbance. 9.11 PRE-WETTING GLASS FIBRE There are a few occasions (during aircraft structure repairs) when the use of prewetted cloth is expedient. The cloth is laminated on flat cellophane or plastic film and as many as four layers may be laminated at once. The pre-wetted cloth is then transferred to the job and stippled into place before the plastic film is then peeled off. During these occasions the following points must be noted: • Care must be taken to ensure that the pre-wetted cloth produces a good bond to the parent material. • The plastic backing film should be peeled off as the cloth is being laid because, with it in place, the laminations cannot assume a double curvature or irregular shape. • It is important to ensure that no bubbles are trapped, though it is quite difficult to detect bubbles in a multi-layer lamination. • The edges of each cloth layer must be staggered so that there is not an abrupt end to a number of layers.


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10 ADHESIVES AND SEALANTS WARNING: CONTROLLED VENTILATION, PROTECTIVE CLOTHING, AND ANTI-FIRE/EXPLOSION PRACTICES, ARE ABSOLUTELY ESSENTIAL WHEN WORKING WITH ADHESIVES AND SEALANTS. ALTHOUGH MANY OF THE ADHESIVES IN CURRENT USE ARE SUPPLIED IN FILM FORM, SOME ARE LIQUIDS OR PASTES, FROM WHICH, TOXIC/FLAMMABLE VAPOURS ARE EMITTED, PRIOR TO CURING. MANY OF THE NECESSARY SURFACE PREPARATION SOLVENTS ALSO GIVE OFF TOXIC/FLAMMABLE VAPOURS. Adhesive bonding has been used on an ever-increasing scale and particularly in the aerospace industry. Adhesives are used for constructional tasks varying from aircraft fuselages, flight control surfaces, to propellers and helicopter rotor blades. 10.1 THE MECHANICS OF BONDING The actual adhesive bond may be achieved in two ways: • Mechanical: - here the adhesive penetrates into the surface and forms a mechanical lock, by keying into the surface. It also forms re-entrants, where the adhesive penetrates behind parts of the structure, and becomes an integral part of the component to be joined. • Chemical (Specific): - in this method of bonding, the adhesive is spread over the surfaces to be joined and forms a chemical bond with the surface In practice, most adhesives use both ways of bonding to form a joint. 10.1.1 STRESSES ON A BONDED JOINT

Adhesive joints are liable to experience four main types of stress Joint stress is at a maximum when the adhesive is in shear (refer to Fig. 28). Adhesives should not be used if significant stresses will be carried in tension or peel. Lap joints are the types more, generally favoured, as the strength of the adhesive bond is proportional to the area bonded,):


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Tensile. Where the two surfaces are pulled directly apart.

Joint in Tension

Shear. Where the two surfaces tend to slide across each other. Joint in Shear

Cleavage. Where two edges are pulled apart.

Joint in Cleavage

Peel. Where one surface is stripped back from the other Joint in Peel

Stresses on Bonded Joints Fig. 28


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10.1.2 ADVANTAGES OF ADHESIVES

The major reasons for the widespread use of adhesives are as follows: • No weakening of the component due to the presence of holes. Also providing a smooth finish due to lack of rivet heads. • No local stress raisers, which are present with widely-pitched conventional fasteners (Bolts, rivets etc.). • Can be used to join dissimilar materials and materials of awkward shapes and of different thickness, as rivetting and welding are not always possible on very thin (or very thick) materials. • Although the strength per unit area, may be inferior to a mechanical or welded joint, adhesive bonding takes place over a greater continuous area and, therefore, gives comparable or increased strength, coupled with improved stiffness. • Adhesive and sealants provide electrical insulation and prevent dissimilarmetal corrosion between different materials. • Leak-proof (fuel and gas) joints can be achieved. • The elastic properties of some adhesives, gives flexibility to the joint and may help to damp out vibrations. • Heat-sensitive materials can be joined. 10.1.3 DISADVANTAGES OF ADHESIVES

The major disadvantages associated with adhesive bonding are: • Limited heat resistance. This restricts the process to applications where environmental temperatures will not, generally, be above 200ºC. • Poor electrical and thermal conductivity. • High thermal expansion. • Limited resistance to certain chemicals (i.e. some paint strippers). • Integrity difficult to check with non-destructive testing procedures.


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10.1.4 STRENGTH OF ADHESIVES

The three most important considerations are: • Fail Stress: - fail load within the glued area • Creep behaviour • Durability: - its long-life capability without serious deterioration. 10.2 GROUPS AND FORMS OF ADHESIVES There exists, an enormous range of adhesives, and the correct type, for a particular application will be specified in the relevant repair procedure. Great care must be taken that only the correct type is used as, otherwise, a catastrophic failure may well occur, should an unsuitable adhesive be used on a critical structure. The two major groups of adhesives are: • Flexible • Structural 10.2.1 FLEXIBLE ADHESIVES

Flexible adhesives are used when some flexing, or slight relative movement of the joint, is essential and where high load-carrying properties are not paramount. In general, flexible adhesives are based on flexible plastics or elastomers, whereas structural adhesives are based on resins, (the most common ones being epoxy or polyester) 10.2.2 STRUCTURAL ADHESIVES

Structural adhesives are primarily aimed at applications where high loads must be carried without excessive creep. They are, therefore, relatively rigid, but without being excessively hard or brittle Note: Another group of adhesives is the two-polymer type, which has a reasonably even balance of resin and elastomer, which results in a flexible, yet fairly strong, adhesive 10.2.3 ADHESIVE FORMS

Adhesives can be obtained in a variety of forms, the most common being liquid, paste or film. Others, available, are those such as the special foaming types, which are used to splice honeycomb sections together. Some require heat for curing, whilst others can be cured by the addition of a catalyst or hardener.


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10.3 ADHESIVES IN USE To achieve optimum bonding, performance, and life in service, from adhesives and sealants, it is absolutely crucial to follow carefully planned processes and procedures and to pay the utmost attention to quality at every stage. In fact, the major criticisms, levelled against the use of adhesives, are: • Absolute cleanliness at all stages is essential. Surface preparation of the component is also crucial. To ensure consistent results on structural components, a purpose-built ‘clean room’ is required, in order to reduce contamination to a minimum. • Pressure and heat may be required. Sophisticated equipment is required to produce pressure over the components in areas where adhesives are applied. This will often entail vacuum bags, purpose-built ovens, or pressurised curing ovens (autoclaves). • Inspection of the bonded joint is difficult. Special inspection techniques and test pieces are necessary to check the integrity of the bond. Prior to preparing the mating surfaces for ‘gluing’, it is necessary to carry out a ‘dry’ lay-up i.e. a trial assembly of all related parts to check and adjust the fit if necessary. This procedure is essential, to enable the final assembly ‘wet’ lay-up to proceed without delay, and without the risk of generating swarf or of contaminating specially prepared surfaces. 10.3.1 SURFACE PREPARATION

• Grease, oil, or other contaminants, must be removed by suitable solvents. • An optimum surface roughness must be produced. • Once pre-treated, a surface must be protected from harmful contamination until the bonding process is complete. • Surfaces to be bonded are normally thoroughly cleaned/degreased in a suitable solvent. This may be followed by a chemical etch or light blasting treatment, followed by a water wash and subsequent drying. 10.3.2 FINAL ASSEMBLY

The adhesive is applied (usually within a specified time, otherwise re-processing may be necessary), and the assembly suitably clamped, or put in a nylon vacuum bag, and heated in an autoclave. The curing process then takes place under carefully controlled temperature and pressure conditions. When cool, the component is inspected, visually for positioning and for a satisfactory spew line. The glue-line thickness is also checked, with a calibrated electronic probe, and specimen test pieces are tested for shear and peel properties.


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Following a satisfactory inspection, the component is finally given appropriate corrosion protection (usually over-painting). Note: After commencing the final (wet) lay-up, curing of the adhesive must be carried out within a specified time (usually 12 hours). If this period is exceeded by a few hours it is necessary to increase the temperature and pressure levels during curing (and to obtain an official ‘concession’ cover for this discrepancy). If the permissible time between wet lay-up and curing is greatly exceeded (e.g. a full shift or day), it will be necessary to dismantle and not only re-commence the wet lay-up, but also to, possibly, repeat some of the preliminary surface preparation treatments (such as etching). 10.3.3 TYPICAL (ABBREVIATED) PROCESS

• Dry lay-up (i.e. ‘dummy run’) • Prepare faces to be bonded (alumina blast, etch (pickle) anodise, etc). • Water wash and dry. • Apply adhesive in clean room and clamp or apply vacuum bag. • Cure in press/oven or autoclave (typically 120ºC - 170ºC) • Release autoclave pressure when cool. • Inspect: • Positioning, uniform, continuous glue-line etc. • Glue-line thickness (electronic probe). • Specimen test-piece results (shear and peel). • Carry out final post-cure surface treatments. (e.g. over-painting of primer, sealant or top coat of solvent-resistant paint) 10.4 SEALING COMPOUNDS Certain areas of all aircraft are sealed to withstand pressurisation, prevent fuel or fume leakage and to delay the onset of corrosion, by sealing against the weather. Most sealant compounds, consist of two or more ingredients, that are compounded to produce a desired combination of strength, flexibility and adherence. Some materials are ready-for-use, straight from their packaging, whilst others require mixing before application.


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10.4.1 ONE-PART SEALANTS

One-part sealants are prepared by the manufacturer and are ready for application straight from their packaging. The consistency of some of these compounds can be altered to satisfy a particular application method. If, for example, thinning is required, then a thinner (recommended by the sealant manufacturer), is mixed with the sealant. 10.4.2 TWO-PART SEALANTS

Two-part sealants are compounds requiring separate packaging, to prevent curing prior to application. The two parts are identified as the base sealing compound and the accelerator. Two-part sealants are generally mixed, by combining equal portions (by weight), of the base and accelerator compounds and any deviation from the prescribed ratios will result in inferior sealing or adhesion. Many common sealants/adhesives are produced in pre-measured kits, that simply require the mixing together of the whole quantities of the materials supplied. These eliminate the need for balances and other weighing equipment. The instructions must be followed but, in general, require the addition of the accelerator to the base compound, followed by thorough mixing before application. A working life is usually quoted, which applies after mixing, so the work must be thoroughly prepared prior to mixing. Some materials may be kept, after mixing, for a limited time, by the use of refrigeration. The instructions will give details if this is possible. 10.4.3 SEALANT CURING

The curing rate, of mixed sealants, varies with temperature and humidity. For example, at temperatures below 15°C, curing is extremely slow. At temperatures above 21°C, curing times are usually faster. For best results, a temperature of around 25°C, with a relative humidity of 50%, is ideal for curing most sealants. If the temperature of curing is increased to accelerate the curing time, it must not exceed 50°C at any time during the curing cycle. The heat can be applied, by using infrared lamps, or heated air, providing the air is dry and filtered. A practical test, to see if curing has been completed, can be done by laying a sheet of cellophane on the work, and checking whether the sheet adheres to it (lack of adhesion indicates full curing).


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INTENTIONALLY BLANK


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11 CORROSION Corrosion costs the civil aircraft industry many millions of pounds (sterling) each year and, with care and good husbandry, this figure can be reduced. The more that aircraft can be manufactured, operated and maintained with the short- and long-term considerations of the effects of corrosion in mind, then the more those maintenance costs will be reduced. Metallic elements are usually compounded with other elements, in the ground, before they are mined and (compared to the actual metals into which they are subsequently formed) they are relatively stable. Corrosion is the tendency of metals to revert to the thermodynamically more stable, oxidised, state. This occurs when they react with dry air to form metal oxides, or with acids and alkalis to form metallic salts. Some metals, such as gold and platinum, strongly resist corrosion. Reactions, between metals and their environments, can occur in either of two (often simultaneous) ways: • •

chemical (oxidation) electrochemical (galvanic)

In both cases, the metal is converted into metal compounds such as carbonates, hydroxides, oxides or sulphates. The corrosion process involves two concurrent changes. The metal that is attacked, suffers an Anodic change while the corrosive agent undergoes a Cathodic change. The result is that material is lost from the Anode and gained by the Cathode, forming an ionic bond. 11.1 CHEMICAL (OXIDATION) CORROSION In a strict chemical sense, oxidation occurs whenever a metal is converted to its ions. An ion is a neutral atom that has gained or lost one or more of its electrons. The term oxidation is, however, normally used to describe the direct combination of a metal with the oxygen of the atmosphere. The phenomenon is essentially a ‘dry’ one, although water vapour, in the air, does play a part in the oxidation of some metals. With the exception of gold and platinum, all metals, in contact with air, form a very thin, visible oxide film. Chemical corrosion can be caused by direct exposure, of the metal surface, to caustic liquids or gaseous agents such as: • Spilled battery acids or battery fumes. Spilled acids are less of a problem now that Nickel Cadmium batteries are in common use.


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• Flux deposits from inadequately cleaned joints. Flux residues are hygroscopic (readily absorb moisture). • Entrapped caustic cleaning compounds. Caustic cleaning solutions should be kept capped when not in use. Many corrosion-removal solutions are, in fact, corrosive agents and should be carefully removed after use. 11.1.1 EFFECT OF OXIDE THICKNESS

The oxide film, that forms on metals, generally tends to protect them from further corrosive attack. The oxidation rate normally falls sharply as the film thickness increases (refer to Fig. 29), so that, at some time, there is virtually no further increase in film thickness.

Temperature Constant Oxide Thickness

Time Oxide Thickness over Time Fig. 29 The graph shows the normal situation with no temperature increase but, occasionally, there is a continuation of oxidation, due to the fact that oxides may react chemically, or combine with, water to produce a film that is not impervious to the passage of further oxygen through it. The oxide skin may also crack or flake and expose the metal surface to further oxidation.


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11.1.2 EFFECT OF TEMPERATURE

The effect of an increase in temperature usually results in an increase in the rate of oxidation of a metal (refer to Fig. 30). The actual curves are not as smooth as those shown.

550°C 525°C 500°C 450°C Oxide Thickness

Time Effect of Temperature on Oxidation Fig. 30


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11.1.3 EFFECT OF ALLOYING

Alloying a metal with another metal sometimes improves the oxidation resistance of the original metal (refer to Fig. 31). The graph shows the effect of adding varying amounts of aluminium (Al) to iron. It can be seen that larger amounts of aluminium result in a slower oxidation rate.

0% Al

Oxide Thickness

+ 3% Al

+ 7% Al

Time Effect of Alloying on Oxidation Fig. 31 The reason for this effect is that the oxide film, which forms, is rich in aluminium oxide, and provides more protection than iron oxide. This process is also involved when chromium is added to nickel to produce ‘stainless steel’, on which, the reaction with air on the chromium produces a protective film of chromium oxide.


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11.2 ELECTROCHEMICAL (GALVANIC) CORROSION A more complicated form of corrosion, which can occur not only on the surface of a metal, but also within the granular structure of the metal (especially in alloys). 11.2.1 THE GALVANIC CELL

The mechanism of electrochemical corrosion (on single metal and at bimetallic surfaces) is similar to that of a primary cell, which produces a low-voltage direct current. In its basic form, it consists of two dissimilar metals in the presence of an electrolyte, An electrolyte is a chemical (or its solution in water), which is able to conduct an electric current, due to the process of ionisation. This forms a simple electric cell in which the less ‘noble’ metal (the anode) is eaten away. When, for example, zinc and copper plates, are partially immersed in an electrolyte, of dilute sulphuric acid, and are connected to an ammeter and voltmeter, the potential difference, between the plates, causes a current to flow (refer to Fig. 32). V e

A

Zn Zn ZNSO4

++

e

°° ° ° ° ° ° ° ° 2H+2e H2 Cu

H2 SO4

2H+ + SO4 --

A Galvanic Cell Fig. 32 The zinc forms the anode of the cell, and is oxidised into ions that dissolve into the acid. At the surface of the copper plate (the cathode), a balancing reaction occurs. The electrons, formed in the anode, are conducted around the circuit and meet with positively charged hydrogen ions at the cathode, to give off hydrogen gas. The thermodynamic driving force of this cell is the difference in galvanic potential between the two metals (zinc and copper). The metal of lower potential (the anode) in such a cell is oxidised or corroded. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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Similar electrochemical corrosion processes, with balancing anodic and cathodic reactions, occur in neutral (non-acidic) electrolytes such as water. The anodic reaction will involve oxidation (corrosion) of the metal with the lower galvanic potential, but the cathodic reaction will, usually, be the reduction of oxygen dissolved in the electrolyte. 11.2.2 FACTORS AFFECTING THE RATE OF CORROSION IN A GALVANIC CELL.

The onset of corrosion (and its severity) will depend upon several factors: • Conductivity of the Solution: - Should the resistance of the solution increase, then the rate of current flow will decrease. This explains why little corrosion occurs in pure water (which has a high resistance), whilst quite severe corrosion occurs in salt water which conducts electricity quite well. Adding various chemicals to the electrolyte can change the resistance and, therefore, the reaction of the galvanic cell. Adding sodium chloride (salt) to the solution, lowers the resistance of the circuit and, hence, increases the current. An acid, such as hydrochloric acid, added to the solution, will remove the oxide film from the plate, which will also lower the resistance, and increase the current flow. • Potential Difference between the Metals: - The galvanic potentials of metals and alloys, can be measured and typical values found in solutions of seawater, or water with 3.5% salt dissolved in it. Table 6 shows, in any combination of two metals, that one will be the anode, and one the cathode. It will NOT, however, predict the severity of the corrosion, as this depends on the type of electrolyte present. Table 6 EXTRACT FROM THE GALVANIC SERIES Extract from the Galvanic Series (Based on Hydrogen at 25°C (298 K)) Potential in Volts

Material

-2.71 -2.38 -1.66 -1.63 -0.76 -0.74 -0.44 -0.40 -0.25 -0.14 -0.13 0 +0.34 +0.80 +1.2 +1.43

Sodium Magnesium Aluminium Titanium Zinc Chromium Iron Cadmium Nickel Tin Lead Hydrogen Copper Silver Platinum Gold


Anodic/Cathodic Anodic

Cathodic

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• Electrical Resistance: - As the corrosion products build up between two metals in contact, and with an electrolyte present, the products can, in some instances, increase the resistance of the action. This will result in slowing or even halting the reaction. Alternatively, the products can bridge any insulation, which has been placed between the metals, and start an electrolytic action. • Ratios of Areas: - If the ratio of the anode to cathode area is not unity, then the rate of corrosion can be much faster (or slower), than would be obtained if they were of equal areas. If the cathode area is small, relative to the anode area, then the rate of corrosion is slow. If the cathode area is much larger than the anode area, then the corrosion can be quite severe (refer to Fig. 33).

Aluminium Rivet

Steel Sheet

Steel Rivet

Aluminium Sheet

Effect of Anode and Cathode Areas on the Rate of Corrosion Fig. 33 • Single Metal Cells: - Corrosion can happen within alloys or metallic mixtures and can occur between metal grains and their grain boundaries, as well as in several other places. It can also occur if small metallic impurities are present within a pure metal, even if the amount of impurity is merely a fraction of one percent. The removal of impurities from metals, at the manufacturing stage, can greatly improve their corrosion resistance. • Oxygen Concentration (Differential Aeration): - Corrosion can occur when the composition of the electrolyte varies at different parts of the contact area. For example, if the electrolyte is in contact with the air, the oxygen can be absorbed, giving a high ‘dissolved oxygen’ level, whilst the electrolyte elsewhere (in a crevice perhaps), will be low in dissolved oxygen. The effect of this is to make the metal, close to the highly oxygenated part, a cathode and that in contact with lower oxygenated part, an anode and so corrosion will begin and, consequently, the crevice (pitting) increases in depth. • Non-Uniform Temperature: - Differences in temperature at varying points will also have the effect of producing different potentials at these points. This can result in severe corrosion in components such as radiators and heat-exchangers.


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11.3 TYPES OF CORROSION There are many forms of corrosion. The form may depend on the metals involved, their function, atmospheric conditions and corrosive agents present. The following are the more common found on aircraft structures. Surface Dissimilar Metal Intergranular Exfoliation Stress Fretting Crevice Filiform Pitting Corrosion Fatigue Microbiological Hydrogen Embrittlement 11.3.1 SURFACE CORROSION

General roughening, etching or pitting of the metal surface, frequently accompanied by a powdery deposit of corrosion products, may be caused by direct chemical or electrochemical attack. Corrosion can spread under the surface coating unnoticed, until the paint or plating is lifted off the surface by the corrosion products or forms blisters. Surface corrosion is a fairly uniform corrosion attack, which slowly reduces the cross-section of the metal. It is, possibly, the least damaging form of corrosion. A mild attack may result in only general etching of an area, whilst a heavier attack may produce deposits which depend on the type of metal that is being attacked. ‘Pure’ aluminium, stainless steel and copper have more resistance to surface corrosion than aluminium alloy, magnesium alloy and non-stainless steels. This type of corrosion only becomes serious over a period of time and gives a warning of worse corrosion to follow. 11.3.2 DISSIMILAR METAL CORROSION

Galvanic action leads to one of the more common forms of corrosion, which occurs between two dissimilar metals in contact with each other and where there is moisture present. It is caused by the difference in galvanic potential of the two metals where plating or jointing compound has been removed or omitted. This type of corrosion can occur, for example, where steel bolts, nuts, or studs are in contact with magnesium-rich alloys such as aircraft wheels. This may be taking place out of sight and may result in extensive pitting. It may or may not be accompanied by surface corrosion.


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11.3.3 INTERGRANULAR CORROSION

This corrosion is also known as intercrystalline corrosion, and results from microgalvanic cells at the grain boundaries in the metal (refer to Fig.34). Corrosion progresses from the metal surface, in narrow pathways, along grain boundaries, often penetrating quite deeply and having a serious, mechanical weakening effect. The amount of metal corroded is small, relative to the volume of metal affected. Indications of the damage may NOT be visible to the naked eye. Intergranular corrosion may often be detected by ultrasonic, eddy current or radiographic inspection procedures.

Intergranular Corrosion Fig. 34


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11.3.4 EXFOLIATION CORROSION

Exfoliation (or layer) corrosion, of certain wrought aluminium alloys, is a form of intergranular corrosion in which the attack occurs in layers parallel to the surface. The wedging action, of the corrosion products, occupies a larger volume than the alloy, and will cause lifting of the metal surface, causing it to ‘exfoliate’. This occurs at an early stage, when the corrosion is on, or just below, the surface. Exfoliation corrosion often attacks 7000 series alloys (those with an appreciable amount of Zinc). When the corrosion occurs well below the surface, extensive damage can occur before the surface deformation is apparent. Spars, stringers and other high-strength parts, which are extruded or hot rolled, are often (because the grains tend to form in layers) susceptible to this kind of corrosion if they have been poorly heat-treated. 11.3.5 STRESS CORROSION

Stress corrosion cracking is a cracking process, caused by the combined action of a sustained tensile stress and a corrosive environment. Only certain combinations of alloys and environments result in stress corrosion cracking, although this type of failure may occur at stresses well below the yield strength of the alloys. Many of the high-strength structural alloys, used in aircraft, are prone to stress corrosion cracking in a wide range of environments and they are particularly susceptible in marine environments. In aircraft alloys, the principal stresses, causing this stress corrosion cracking, are not the applied service loads, but the stresses developed within the metal during manufacture and during assembly. For example, internal stresses can arise from quenching after heat-treatment, from ‘force fits’, from badly mating parts, or from welding procedures. Service stresses are only significant when they act in the same direction as internal or assembly stresses. Stress corrosion cracking has three distinct phases in that there is an initial ‘Incubation’ period, (when a stress corrosion crack starts from pitting or film breakdown). The incubation is followed by a period of ‘Slow Growth’ of the stress concentrations and culminates in a short, ‘Rapid Crack-Growth’ rate. In highly stressed parts (e.g. landing gear components), cracks may originate from a stress raiser such as a scratch or surface corrosion. This problem is characteristic of aluminium, copper, stainless steels and high-strength alloy steels and may occur along lines of cold working. Signs of stress corrosion are given by minute cracks radiating from areas of the greatest stress concentration. Likely areas for this type of corrosion are U/C jacks, shock absorbers, bellcranks with pressed-in bushes, or other areas where parts are a force fit, highly stressed or have residual stresses induced during the forming process.


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11.3.6 FRETTING CORROSION

Fretting corrosion is the result of rubbing movement between two heavily loaded surfaces, one, or both, of which are metallic. The rubbing action destroys any natural protective film and also removes particles of metal from the surface. In its early stages, the debris of this corrosion forms a black powder. These particles form an abrasive compound, which aggravates the effect of the rubbing action and the surface is continually removed to expose fresh metal to the corrosive attack. This form of attack can eventually cause cracking and fatigue failure. The most likely areas affected are gears, screw jacks, loose panels, splined hydraulic pump drives and rivets (when they become loose). , It may be serious enough to cause cracking and fatigue failure. 11.3.7 CREVICE CORROSION

Crevices are liable to preferential attack, usually by a differential aeration form of corrosion, intensified by the high ratio of cathode to anode area involved. The attack is more severe where crevices collect dust and moisture (Fig. 35).

Low Oxygen Concentration (becomes anodic) High Oxygen Concentration (becomes cathodic)

Crevice Crevice Corrosion Fig. 35 Severe localised corrosion occurs at narrow openings or gaps between metal components, often due to flexing. Corrosive agents are able to penetrate into the joint. 11.3.8 FILIFORM CORROSION

Filiform corrosion occurs beneath thin, protective coatings, on aluminium and steel alloys, with the paint or coating often bulging or blistering. On aircraft structures, the attack often starts at fasteners and extends as thread-like lines of corrosion under the paint. It may not be readily visible until it has become quite severe. The damage tends to be very shallow and is not, usually, structurally dangerous.


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11.3.9 PITTING CORROSION

Pitting corrosion can occur on aircraft materials when the protective film, whether applied or natural, breaks down locally and this may also lead to intergranular corrosion. The corrosion often stems from the screening effect of silt, scale or corrosion deposits that reduce the oxygen concentration at local points on the metal surface, which establishes differential concentration cells. Local rough spots, inclusions, contaminations and lack of homogeneity in the alloy or metal are also possible causes of pitting. In size and depth, the pits are widely variable and a large number of pits can give a surface a ‘blotchy’ appearance. Aluminium and magnesium alloys, chromium-plated and stainless steels (including nitrided surfaces), are all particularly susceptible to this form of corrosion. Pitting corrosion of an aluminium alloy component can be detected by the appearance of white powder on the surface of the metal (refer to Fig. 36).

Pitting Corrosion of an Aluminium Alloy Component Fig. 36 EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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CORROSION FATIGUE

This is similar to stress corrosion cracking, except that the applied loads are cyclic instead of static. Crack propagation is aided by the corrosion that occurs, at the root of the crack, during the tensile part of the loading cycle. 11.3.11

MICROBIOLOGICAL CONTAMINATION

This is caused, directly or indirectly (and in one or more ways), by microorganisms which are not only able to produce corrosive substances (such as hydrogen sulphide, ammonia and inorganic acids), but can also act as depolarisers or catalysts in corrosion reactions. Local depletion of oxygen and water, held in contact with a metal surface, by matted fungi and micro-organisms, all contribute towards establishing corrosive environments. The commonest form of microbiological corrosion in aircraft, is that, which is caused by contamination of fuel tanks (unless the fuel has an additive to protect against it). The growth of the fungi depends on several conditions, but a high ambient temperature can drastically increase the rate of growth, and especially so when the temperature is above 30°C with a high relative humidity. This microbiological growth is sometimes called Cladosporium Resinæ. Where fungal growth has formed, there is a probability that corrosion of the tank will occur. The organisms, resembling a mucous, can cause problems with filters and with the fuel contents gauge units. The roots of the fungus, penetrating the internal sealing and protective coatings of fuel tanks can cause further problems. In well-developed contaminations, a dense mat of fungus forms on the floor of the tank, retaining water and preventing free flow to the water drain-valve. In integral fuel tanks, this can result in serious corrosion of the aircraft structure such that penetration of the bottom wing skin has been known to occur. Spillage, of organic materials, from around galley and toilet areas, provides a further source of microbial contamination. There is evidence that such spillage can be more corrosive than its chemical composition (acidity and chloride content) possibly due to fermentation by yeast and bacteria. 11.3.12

HYDROGEN EMBRITTLEMENT OF STEELS

Many of the standard surface protection treatments, including cleaning and electroplating, are liable to introduce hydrogen into steel. To avoid embrittlement, the steels must be ‘baked’, at a temperature of around 200°C, following the treatments. The duration of the baking is dependent on the strength of the steel. High-tensile steels are much more susceptible to hydrogen embrittlement than are other metals.


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Hydrogen embrittlement shows itself in slow strain-rate mechanical tests and not by fast rate tests such as in impact testing. These steels can show a sudden failure after many weeks of loading at well below their normal yield strength. 11.4 FACTORS AFFECTING CORROSION Many factors will affect the cause, type, speed of attack, and seriousness of metal corrosion. Some are beyond the control of the aircraft designer or maintenance engineer while some of them can be controlled. 11.4.1 CLIMATIC

The environmental conditions under which the aircraft is operated and maintained cannot normally be controlled. The following factors will effect the rate at which corrosion will occur. • Marine environments (exposure to salt water) will increase rate of corrosion. • Moisture laden atmosphere as against a dry atmosphere. The USA store hundreds of aircraft in a desert (dry) atmosphere for emergency war use. • Temperature considerations i.e. Hot climate against cold climate. High temperatures will increase the rate of corrosion (all chemical reactions occur faster at higher temperatures). The worst conditions would exist in a hot, wet, maritime environment. 11.4.2 SIZE AND TYPE OF METAL

Some metals corrode more easily than others. Magnesium corrodes readily, whilst Titanium is extremely corrosion-resistant because it oxidises readily. Thick structural sections are also more susceptible than thin sections, because variations in physical characteristics are greater. Such sections are also likely to have been cold worked and are, therefore, more susceptible to stress corrosion. 11.4.3 CORROSIVE AGENTS

Foreign materials, that may adhere to metal surfaces, and, consequently result in corrosion, can include: • Soil and atmospheric dust • Oil, grease and engine exhaust residues • Salt water and salt moisture condensation • Spilled battery acids and caustic cleaning solutions • Welding and brazing flux residues


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11.5 COMMON METALS AND CORROSION PRODUCTS One of the problems involved in corrosion control, is the recognition of corrosion products whenever they occur. The following brief descriptions are of typical corrosion products, common to materials used in aircraft construction. 11.5.1 IRON AND STEEL

The most common, and easily-recognisable, form of corrosion is red rust. The initial oxide film, formed on freshly exposed steel, is very thin and invisible. In the presence of water, or in a damp atmosphere, especially if sulphur dioxide (industrial atmosphere) or salt (marine environment) is present, thick layers of hydrated oxide develop. These layers vary in colour from brown to black. Rust promotes further corrosion by retaining salts and water. Mill scale (a type of oxide formed at high temperatures), also promotes rusting, by forming an electrolytic cell with the underlying steel. Heavy deposits of rust can be removed only by abrasive blasting or by immersion in rust-removing solutions. Surface rust can develop on steel nuts, bolts and other fasteners and may not adversely affect the operational integrity of the equipment. Its appearance is an indication that adequate maintenance procedures have not been followed. 11.5.2 ALUMINIUM ALLOYS

The corrosion of aluminium and its alloys, takes a number of different forms. It may vary from general etching of the surface, to the localised, intergranularattack, characteristics of some strong alloys in certain states of heat-treatment. The corrosion products are white to grey and are powdery when dry. Superficial corrosion can be removed by scouring, light abrasive blasting, or by chemical methods. In general, pure aluminium sheet and ‘alclad’ surfaces have good corrosion resistance, except in marine environments. In these areas, aluminium and its alloys need protection and high-strength aluminium alloys are always given a substantial protective treatment.


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11.5.3 MAGNESIUM ALLOYS

Magnesium corrosion products are white and voluminous, compared to the base metal. When the failure of protective coatings on magnesium alloys occurs, the corrosive attack tends to be severe in the exposed areas, and may penetrate totally through a magnesium structure in a very short time. Any corrosion, on magnesium alloys, therefore requires prompt attention. In contrast to highstrength aluminium alloys, the strong magnesium alloys, used in aircraft, do not suffer intergranular attack. Corrosion is readily visible on the surfaces of Magnesium Alloys. 11.5.4 TITANIUM

Titanium is highly corrosion-resistant, but should be insulated from other metals to avoid dissimilar metal corrosion of the adjacent material. Titanium alloys can suffer stress corrosion at temperatures above 300°C when in the presence of salt and fatigue cracks can develop more quickly in a saline atmosphere. Cadmium can penetrate the surface of titanium alloys and embrittle them at all temperatures above ambient (as can Lead, Tin and Zinc at temperatures higher than approximately 120°C)). Embrittlement can occur if the cadmium is plated onto the titanium or if cadmium-plated steel parts (and cadmium-contaminated spanners) are used with titanium. Great care must be taken to ensure that these conditions never occur if at all possible. 11.5.5 COPPER ALLOYS

Copper and its alloys are relatively resistant to corrosion. Tarnishing has no serious consequences in most applications. Long-term exposure to industrial or marine atmospheres gives rise to the formation of the blue-green patina (aerugo or verdigris) on copper surfaces, while brasses can suffer selective removal of zinc (de-zincification). In aircraft construction, copper-based alloys are frequently cadmium-plated, to prevent dissimilar metal corrosion. 11.5.6 CADMIUM AND ZINC

Cadmium and zinc are used as coatings, to protect the parts to which they are applied. Both confer sacrificial protection on the underlying metal. Cadmium is normally chosen for use in the aircraft industry, as it is more durable under severe corrosive conditions such as in marine and tropical environments. Both metals produce white corrosion products.


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11.5.7 NICKEL AND CHROMIUM

Electroplated nickel is used as a heat-resistant coating, while chromium is used for its wear-resistance. Both metals protect steel only by excluding the corrosive atmosphere. The degree of protection is proportional to the thickness of the coating. Once the underlying steel is exposed (through loss of the coating, due to abrasion or other damage), then the coatings actually accelerate the rusting, due to the fact that the steel is more anodic than the protective coating. Chromium is also highly resistant to corrosion, whilst Nickel corrodes slowly in industrial and marine atmospheres, to give a blue-green corrosion product. 11.6 CORROSION REMOVAL General treatments for corrosion removal include: • Cleaning and stripping of the protective coating in the corroded area. • Removal of as much of the corrosion products as possible. • Neutralisation of the remaining residue. • Checking if damage is within limits • Restoration of protective surface films • Application of temporary or permanent coatings or paint finishes. 11.6.1 CLEANING AND PAINT REMOVAL.

It is essential that the complete suspect area be cleaned of all grease, dirt or preservatives. This will aid in determining the extent of corrosive spread. The selection of cleaning materials will depend on the type of matter to be removed. Solvents such as trichloroethane (trade name ‘Genklene’) may be used for oil, grease or soft compounds, while heavy-duty removal of thick or dried compounds may need solvent/emulsion-type cleaners. General purpose, water-removable stripper is recommended for most paint stripping. Adequate ventilation should be provided and synthetic rubber surfaces such as tyres, fabric and acrylics should be protected (remover will also soften sealants). Rubber gloves, acid-repellent aprons and goggles, should be worn by personnel involved with paint removal operations. The following is the general paint stripping procedure: • Brush the area with stripper, to a depth of approximately 0.8 mm – 1.6 mm (0.03 in – 0.06 in). Ensure that the brush is only used for paint stripping. • Allow stripper to remain on the surface long enough for the paint to wrinkle. This may take from 10 minutes to several hours.


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• Re-apply the stripper to those areas which have not stripped. Non-metallic scrapers may be used. • Remove the loosened paint and residual stripper by washing and scrubbing the surface with water and a broom or brush. Water spray may assist, or the use of steam cleaning equipment may be necessary. Note. Strippers can damage composite resins and plastics, so every effort should be made to 'mask' these vulnerable areas. 11.6.2 CORROSION OF FERROUS METALS

Atmospheric oxidation of iron or steel surfaces causes ferrous oxide rust to be deposited. Some metal oxides protect the underlying base metal, but rust promotes additional attack by attracting moisture and must be removed. Rust shows on bolt heads, nuts or any un-protected hardware. It’s presence is not immediately dangerous, but it will indicate a need for maintenance and will suggest possible further corrosive attack on more critical areas. The most practical means of controlling the corrosion of steel is the complete removal of corrosion products by mechanical means. Abrasive papers, power buffers, wire brushes and steel wool are all acceptable methods of removing rust on lightly stressed areas. Residual rust usually remains in pits and crevices. Some (dilute) phosphoric acid solutions may be used to neutralise oxidation and to convert active rust to phosphates, but they are not particularly effective on installed components. 11.6.3 HIGH-STRESSED STEEL COMPONENTS

Corrosion on these components may be dangerous and should be removed carefully with mild abrasive papers or fine buffing compounds. Care should be taken not to overheat parts during corrosion removal. Protective finishes should be re-applied immediately. 11.6.4 ALUMINIUM AND ALUMINIUM ALLOYS

Corrosion attack, on aluminium surfaces, gives obvious indications, since the products are white and voluminous. Even in its early stages, aluminium corrosion is evident as general etching, pitting or roughness. Aluminium alloys form a smooth surface oxidation, which provides a hard shell, that, in turn, may form a barrier to corrosive elements. This must not be confused with the more serious forms of corrosion. General surface attack penetrates slowly, but is speeded up in the presence of dissolved salts. Considerable attack can take place before serious loss of strength occurs. Three forms of attack, which are particularly serious, are: • Penetrating pit-type corrosion through the walls of tubing.


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• Stress corrosion cracking under sustained stress. • Intergranular attack characteristic of certain improperly heat treated alloys. Treatment involves mechanical or chemical removal of as much of the corrosion products as possible and the inhibition of residual materials by chemical means. This, again, should be followed by restoration of permanent surface coatings. 11.6.5 ALCLAD

WARNING: USE ONLY APPROVED PAINT STRIPPERS IN THE VICINITY OF REDUX BONDED JOINTS. CERTAIN PAINT STRIPPERS WILL ATTACK AND DEGRADE RESINS. USE ADEQUATE PERSONAL PROTECTIVE EQUIPMENT WHEN WORKING WITH CHEMICALS. USE ONLY THE APPROVED FLUIDS FOR REMOVING CORROSION PRODUCTS. INCORRECT COMPOUNDS WILL CAUSE SERIOUS DAMAGE TO METALS. Obviously great care must be taken, not to remove too much of the protective aluminium layer by mechanical methods, as the core alloy metal may be exposed, therefore, where heavy corrosion is found, on clad aluminium alloys, it must be removed by chemical methods wherever possible. Corrosion-free areas must be masked off and the appropriate remover (usually a phosphoric-acid based fluid) applied, normally with the use of a stiff bristled brush, to the corroded surface, until all corrosion products have been removed. Copious amounts of clean water should, next, be used to flood the area and remove all traces of the acid, then the surface should be dried thoroughly. Note: A method of checking that the protective aluminium coating remains intact is by the application of one drop of diluted caustic soda to the cleaned area. If the alclad has been removed, the alumium alloy core will show as a black stain, whereas, if the cladding is intact, the caustic soda will cause a white stain. The acid must be neutralised and the area thoroughly washed and dried before a protective coating (usually Alocrom 1200 or similar) is applied to the surface. Further surface protection may be given by a coat of suitable primer, followed by the approved top coat of paint. 11.6.6 MAGNESIUM ALLOYS

The corrosion products are removed from magnesium alloys by the use of chromic/sulphuric acid solutions (not the phosphoric acid types), brushed well into the affected areas. Clean, cold water is employed to flush the solution away and the dried area can, again, be protected, by the use of Alocrom 1200 or a similar, approved, compound.


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11.6.7 ACID SPILLAGE

An acid spillage, on aircraft components, can cause severe damage. Acids will corrode most metals used in the construction of aircraft. They will also destroy wood and most other fabrics. Correct Health and Safety procedures must be followed when working with such spillages. Aircraft batteries, of the lead/acid type, give off acidic fumes and battery bays should be well ventilated, while surfaces in the area should be treated with antiacid paint. Vigilance is required of everyone working in the vicinity of batteries, to detect (as early as possible) the signs of acid spillage. The correct procedure to be taken, in the event of an acid spillage, is as follows: • Mop up as much of the spilled acid using wet rags or paper wipes. Try not to spread the acid. • If possible, flood the area with large quantities of clean water, taking care that electrical equipment is suitably protected from the water. • If flooding is not practical, neutralise the area with a 10% (by weight) solution of bicarbonate of soda (sodium bicarbonate) with water. • Wash the area using this mixture and rinse with cold water. • Test the area, using universal indicating paper (or litmus paper),to check if acid has been cleaned up. • Dry the area completely and examine the area for signs of damaged paint or plated finish and signs of corrosion, especially where the paint may have been damaged. • Remove corrosion, repair damage and restore surface protection as appropriate. 11.6.8 ALKALI SPILLAGE

This is most likely to occur from the alternative Nickel-Cadmium (Ni-Cd) or NickelIron (Ni-Fe) type of batteries, containing an electrolyte of Potassium Hydroxide (or Potassium Hydrate). The compartments of these batteries should also be painted with anti-corrosive paint and adequate ventilation is as important as with the lead/acid type of batteries. Proper Health and Safety procedures are, again, imperative. Removal of the alkali spillage, and subsequent protective treatment, follows the same basic steps as outlined in acid spillage, with the exception that the alkali is neutralised with a solution of 5% (by weight) of chromic acid crystals in water.


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11.6.9 MERCURY SPILLAGE

WARNING: MERCURY (AND ITS VAPOUR) IS EXTREMELY TOXIC. INSTANCES OF MERCURY POISONING MUST, BY LAW, BE REPORTED TO THE HEALTH AND SAFETY EXECUTIVE. ALL SAFETY PRECAUTIONS RELATING TO THE SAFE HANDLING OF MERCURY MUST BE STRICTLY FOLLOWED. Mercury contamination is far more serious than any of the battery spillages and prompt action is required to ensure the integrity of the aircraft structure. While contamination from mercury is extremely rare on passenger aircraft, sources of mercury spillage result from the breakage of (or leakage from) containers, instruments, switches and certain test equipment. The spilled mercury can, quickly, separate into small globules, which have the capability of flowing (hence its name ‘Quick Silver’) into the tiniest of crevices, to create damage. Mercury can rapidly attack bare light alloys (it forms an amalgam with metals), causing intergranular penetration and embrittlement which can start cracks and accelerate powder propagation, resulting in a potentially catastrophic weakening of the aircraft structure. Signs of mercury attack on aluminium alloys are greyish powder, whiskery growths, or fuzzy deposits. If mercury corrosion is found, or suspected, then it must be assumed that intergranular penetration has occurred and the structural strength is impaired. The metal in that area should be removed and the area repaired in accordance with manufacturer’s instructions. Ensure that toxic vapour precautions are observed at all times during the following operation: • Do not move aircraft after finding spillage. This may prevent spreading. • Remove spillage carefully by one of the following mechanical methods: • Capillary brush method (using nickel-plated carbon fibre brushes). • Heavy-duty vacuum with collector trap. • Adhesive tape, pressed (carefully) onto globules may pick them up • Foam collector pads (also pressed, carefully, onto globules). • Alternative, chemical methods, of mercury recovery entail the use of: • Calcium polysulphide paste. • Brushes, made from bare strands of fine copper wire • Neutralise the spillage area, using ‘Flowers of Sulphur’.


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• Try to remove evidence of corrosion. • The area should be further checked, using radiography, to establish that all globules have been removed and to check extent of corrosion damage. • Examine area for corrosion using a magnifier. Any parts found contaminated should be removed and replaced. Note: Twist drills (which may be used to separate riveted panels, in an attempt to clean contaminated surfaces) must be discarded after use. • Further, periodic checks, using radiography, will be necessary on any airframe that has suffered mercury contamination. 11.7 PERMANENT ANTI-CORROSION TREATMENTS These are intended to remain intact throughout the life of the component, as distinct from coatings, which may be renewed as a routine servicing operation. They give better adhesion for paint and most resist corrosive attack better than the metal to which they are applied. 11.7.1 ELECTRO-PLATING

There are two categories of electro-plating, which consist of: • Coatings less noble than the basic metal. Here the coating is anodic and so, if base metal is exposed, the coating will corrode in preference to the base metal. Commonly called sacrificial protection, an example is found in the cadmium (or zinc) plating of steel. • Coatings more noble (e.g. nickel or chromium on steel) than the base metal. The nobler metals do not corrode easily in air or water and are resistant to acid attack. If, however, the basic metal is exposed, it will corrode locally through electrolytic action. The attack may result in pitting corrosion of the base metal or the corrosion may spread beneath the coating. 11.7.2 SPRAYED METAL COATINGS

Most metal coatings can be applied by spraying, but only aluminium and zinc are used on aircraft. Aluminium, sprayed on steel, is frequently used for hightemperature areas. The process (aluminising), produces a film about 0.1 mm (0.004 in) thick, which prevents oxidation of the underlying metal. 11.7.3 CLADDING

The hot rolling of pure aluminium onto aluminium alloy (Alclad) has already been discussed, as has the problem associated with the cladding becoming damaged, exposing the core, and the resulting corrosion of the core alloy


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11.7.4 SURFACE CONVERSION COATINGS

These are produced by chemical action. The treatment changes the immediate surface layer into a film of metal oxide, which has better corrosion resistance than the metal. Among those widely used on aircraft are: • Anodising of aluminium alloys, by an electrolytic process, which thickens the natural, oxide film on the aluminium. The film is hard and inert. • Chromating of magnesium alloys, to produce a brown to black surface film of chromates, which form a protective layer. • Passivation of zinc and cadmium by immersion in a chromate solution. Other surface conversion coatings are produced for special purposes, notably the phosphating of steel. There are numerous proprietary processes, each known by its trade name (e.g. Bonderising, Parkerising, or Walterising). 11.8 LOCATIONS OF CORROSION IN AIRCRAFT Certain locations in aircraft are more prone to corrosion than others. The rate of deterioration varies widely with aircraft design, build, operational use and environment. External surfaces are open to inspection and are usually protected by paint. Magnesium and aluminium alloy surfaces are particularly susceptible to corrosion along rivet lines, lap joints, fasteners, faying surfaces and where protective coatings have been damaged or neglected. 11.8.1 EXHAUST AREAS

Fairings, located in the path of the exhaust gases of gas turbine and piston engines, are subject to highly corrosive influences. This is particularly so where exhaust deposits may be trapped in fissures, crevices, seams or hinges. Such deposits are difficult to remove by ordinary cleaning methods. During maintenance, the fairings in critical areas should be removed for cleaning and examination. All fairings, in other exhaust areas, should also be thoroughly cleaned and inspected. In some situations, a chemical barrier can be applied to critical areas, to facilitate easier removal of deposits at a later date, and to reduce the corrosive effects of these deposits. 11.8.2 ENGINE INTAKES AND COOLING AIR VENTS

The protective finish, on engine frontal areas, is abraded by dust and eroded by rain. Heat-exchanger cores and cooling fins may also be vulnerable to corrosion. Special attention should be given, particularly in a corrosive environment, to obstructions and crevices in the path of cooling air. These must be treated as soon as is practical.


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11.8.3 LANDING GEAR

Landing gear bays are exposed to flying debris, such as water and gravel, and require frequent cleaning and touching-up. Careful inspection should be given to crevices, ribs and lower-skin surfaces, where debris can lodge. Landing gear assemblies should be examined, paying particular attention to magnesium alloy wheels, paintwork, bearings, exposed switches and electrical equipment. Frequent cleaning, water-dispersing treatment and re-lubrication will be required, whilst ensuring that bearings are not contaminated, either with the cleaning water or with the water-dispersing fluids, used when re-lubricating. 11.8.4 BILGE AND WATER ENTRAPMENT AREAS

Although specifications call for drains wherever water is likely to collect, these drains can become blocked by debris, such as sealant or grease. Inspection of these drains must be frequent. Any areas beneath galleys and toilet/wash-rooms must be very carefully inspected for corrosion, as these are usually the worst places in the whole airframe for severe corrosion. The protection in these areas must also be carefully inspected and renewed if necessary. 11.8.5 RECESSES IN FLAPS AND HINGES

Potential corrosion areas are found at flap and speed-brake recesses, where water and dirt may collect and go unnoticed, because the moveable parts are normally in the ‘closed’ position. If these items are left ‘open’, when the aircraft is parked, they may collect salt, from the atmosphere, or debris, which may be blowing about on the airfield. Thorough inspection of the components and their associated stowage bays, is required at regular intervals. The hinges, in these areas, are also vulnerable to dissimilar metal corrosion, between the steel pins and the aluminium tangs. Seizure can also occur, at the hinges of access doors and panels that are seldom used. 11.8.6 MAGNESIUM ALLOY SKINS

These give little trouble, providing the protective surface finishes are undamaged and well maintained. Following maintenance work, such as riveting and drilling, it is impossible to completely protect the skin to the original specification. All magnesium alloy skin areas must be thoroughly and regularly inspected, with special emphasis on edge locations, fasteners and paint finishes. 11.8.7 ALUMINIUM ALLOY SKINS

The most vulnerable skins are those which have been integrally machined, usually in main-plane structures. Due to the alloys and to the manufacturing processes used, they can be susceptible to intergranular and exfoliation corrosion. Small bumps or raised areas under the paint sometimes indicate exfoliation of the actual metal. Treatment requires removal of all exfoliated metal followed by blending and restoration of the finish.


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11.8.8 SPOT-WELDED SKINS AND SANDWICH CONSTRUCTIONS

Corrosive agents may become trapped between the metal layers of spot-welded skins and moisture, entering the seams, may set up electrolytic corrosion that eventually corrodes the spot-welds, or causes the skin to bulge. Generally, spotwelding is not considered good practice on aircraft structures. Cavities, gaps, punctures or damaged places in honeycomb sandwich panels should be sealed to exclude water or dirt. Water should not be permitted to accumulate in the structure adjacent to sandwich panels. Inspection of honeycomb sandwich panels and box structures is difficult and generally requires that the structure be dismantled. 11.8.9 ELECTRICAL EQUIPMENT

Sealing, venting and protective paint cannot wholly obviate the corrosion in battery compartments. Spray, from electrolyte, spreads to adjacent cavities and causes rapid attack on unprotected surfaces. Inspection should also be extended to all vent systems associated with battery bays. Circuit-breakers, contacts and switches are extremely sensitive to the effects of corrosion and need close inspection. 11.8.10

MISCELLANEOUS ITEMS

Loss of protective coatings, on carbon steel control cables can, over a period of time, lead to mechanical problems and system failure. Corrosion-resistant cables, can also be affected by corrosive, marine environments. Any corrosion found on the outside of a control cable should result in a thorough inspection of the internal strands and, if any damage is found, the cable should be rejected. Cables should be carefully inspected, in the vicinity of bell-cranks, sheaves and in other places where the cables flex, as there is more chance of corrosion getting inside the cables when the strands are moving around (or being moved by) these items.


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12 AIRCRAFT FASTENERS WARNING: ONLY THE APPROVED METHODS OF FASTENING DEVICES MUST BE USED ON AIRCRAFT. SUBSTITUTION WITH INCORRECT PARTS CAN CAUSE FATAL FAILURES. Fasteners, or fastening devices, are used to create secure joints between two or more components. Types of fastening devices, used on aircraft, vary in accordance with the materials, which require joining, and the importance of the joined components, or structures, to the safety of the aircraft. The environment in which the joint must operate and the frequency (and ease) with which the joint may need to be disassembled, for inspection, replacement or repair, will also influence the choice of fasteners to be employed. Fasteners may be metallic or non-metallic (or composites of both types). They may be flexible or rigid (or a combination of both) and may be used to form the three basic categories of joints. 12.1 TEMPORARY JOINTS Temporary joints are used where the joint can be disassembled without damage and where, usually, the same fastener can be used to reassemble the joint. Bolts and nuts, circlips and quick-release fasteners are, typically, used in temporary joints. 12.2 PERMANENT JOINTS Permanent joints are those which are not intended to be disassembled on a frequent basis (if at all), and are joints where either the fastening medium or the joined components will suffer damage in their separation. Adhesives, rivets and welds are examples of uses of permanent joints. 12.3 FLEXIBLE JOINTS Flexible joints allow movement of the joined components relative to each other. Anti-vibration mounts, universal couplings and hinges are devices which may be employed in flexible joints. Whatever fasteners are used, to make a particular joint, it must be ensured that only the approved materials are utilised and that their legality is confirmed. This can be done by reference to published Part Numbers, which are to be found in Aircraft Maintenance Manuals, Wiring Diagrams, Structural Repair Manuals, Illustrated Parts Catalogues (also called Illustrated Parts Lists) and other, approved, publications. The use of non-approved fasteners can lead to expensive and, possibly, fatal failures in aircraft and their associated structures. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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12.4 SCREW THREADS Threaded fasteners allow parts to be fastened together with all of the strength that unthreaded fasteners provide. However, unlike rivets and pins, threaded fasteners may be disassembled and reassembled an almost infinite number of times. Due to the large range of different available fasteners, great care must be always be taken to select the correct fastener for each particular installation. Aircraft, bolts, nuts, screws and studs are manufactured to the many, different, International Standards and in a variety of different thread forms, as can be seen in Table 7. Most aircraft now use unified or metric threads but, however, some older aircraft use obsolete British Association (BA), British Standard Fine (BSF) or Whitworth (BSW) thread forms. None of these are compatible with the unified (or metric) thread forms. Table 7 COMMON INTERNATIONAL THREAD STANDARDS International Standard American National Coarse American National Fine Unified Coarse Unified Fine British Association British Standard Fine ISO Metric

Common Abbreviation ANC ANF UNC UNF BA BSF M

12.4.1 THE INCLINED PLANE AND THE HELIX

The value of the wedge, as a means of transmitting motion, is well known. For a constant effort applied in driving a wedge, a smaller angle of inclination between the planes will cause a greater force to be exerted through a shorter distance. Conversely, a larger angle will cause less force to be exerted through a greater distance (refer to Fig. 37).


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Transmission of Motion with an Inclined Plane Fig. 37 Whilst the wedge is, generally, used as a means of transmitting motion, it must be remembered that the action may be reversed and the wedge can be caused to move when a force is applied to the inclined surfaces. This is readily appreciated when the angle is large (and the larger the angle of inclination becomes, then, the more readily is the motion reversed), but, no matter how small the angle may be, the resultant of forces applied will still tend to produce movement. Friction, between the surfaces, may, however, prevent movement from actually occurring. When a continuous, inclined plane is cut around the outside (or the inside) of a cylinder, then a spiral (also known as a ‘helix’) is produced (refer to Fig. 38). The helix angle is important in screw threads, because it dictates the number of threads, which can be cut, per axial linear increment (millimetres or inches) on, or in, the cylinder.

Helix Angle

Helix Angle of a Screw Thread Fig. 38 EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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In a similar manner to the previously mentioned wedges, a thread with a small helix angle (a fine thread), will exert a greater force than one with a larger helix angle (a coarse thread) for a given cylinder diameter. Fine threads are, normally, associated with small and delicate instruments or in equipment, where secure holding power is often required of miniature-sized fasteners. The greater ‘wedging action’ of fine threads also makes them much more dependable in situations where vibration (or a change of temperature) has the tendency to loosen threaded joints. Most aircraft components are assembled using fine threads on the various bolts, nuts, screws and studs, which are then, often, further secured by some other, mechanical, process, to reinforce their resistance to the effects of temperature changes and vibration. 12.5 SCREW THREAD TERMINOLOGY It is often disputed as to the difference between a bolt and a screw, but, generally, it is accepted that a bolt is considered to be a threaded fastener, which has a definite plain portion on the shank, between its head and the beginning of the thread, and is used in conjunction with a nut, whereas a screw is threaded all the way to the head. Because there are so many variations in terminology, with the numerous manufactures, the only safe way of describing a threaded (or any other) fastener is to use the correct terminology, found in the relevant IPC, when ordering replacement items. When defining the length of bolts, reference is usually made to the length of the plain portion of the shank, of hexagonal-headed bolts (refer to Fig. 39), while screw lengths are designated differently, according to their type.

Designation of Fastener Lengths Fig. 39


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Screw threads are usually formed with a ‘clockwise’ turning groove and are referred to as ‘right-hand’ threads, but there are occasions where the thread is formed with the groove spiralling in an ‘anti-clockwise’ direction and, in this instance, they are designated as ‘left-hand’ threads. While a traditional thread shape can be used to illustrate the terminologies, associated with screw threads (refer to Fig. 40), the actual profile, of any thread, will be determined by the Standard or specification to which it is manufactured. This of course, will also be influenced by the use to which the threaded item is to be put.

Screw Thread Terminology Fig. 40 The following terms are used to define the characteristics of a threaded item: •

Major Diameter: The largest diameter of the thread, measured at right angles to the axis.

•

Minor Diameter: The smallest diameter of the thread, measured at right angles to the axis.


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•

Pitch: The distance from the centre of one crest to the centre of the next, measured parallel to the axis.

•

Depth of Thread: The distance between the root and crest, measured at right angles to the axis.

•

Lead: The distance a screw moves axially in one complete turn. In the case of multi-start threads, the lead is equal to the pitch multiplied by the number of starts.

•

Single Start Thread: Term used when there is only one screw thread cut in the material.

• Multi-Start Thread: Consists of two or more separate, parallel threads cut into the material carrying the thread. This method is used in order to achieve a quick-acting motion between two threaded items. • Runout: The part of the thread where the minor diameter increases until it equals the major diameter and merges with the plain portion of the shank. The runout cannot be used and any nut, rotated onto the runout, would become ‘thread-bound’. 12.5.1 SCREW THREAD FORMS

The form of a screw thread will depend upon the function for which it is to be used (refer to Fig. 41). Where the thread is used to join components together (nuts, bolts, screws and studs) then the conventional, truncated ‘V’-shaped threads, similar to the ISO Metric thread, will be found. Turnbuckles and similar devices, (which are employed as adjusters of either the tension or of the distance between components), may also use ‘V’-shaped threads, while the Acme, Buttress and Square threads are utilised to transmit movement or power (as may be seen in lathes, vices and Flap Jacks).


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Pitch 60° Square Thread

ISO Metric Thread Pitch

Pitch 45° 29°

ButtressThread

Acme Thread Screw Thread Forms Fig. 41

Thread forms have developed over the years, from the early standardisation on the BSW thread (with its rather coarse thread, which was prone to slackening when subjected to vibration), to the modern, finer threads which are more suitable for use on aerospace components and structures. In an attempt to provide a common standard, Canada, the United States of America, and the United Kingdom adopted the Unified system of threads. The International Standard Organisation (ISO), later, recommended that the Unified system be used internationally, in parallel with a system using Metric units of measurement, but with a similar form of thread profile and standards of tolerances Unified Coarse (UNC) and Unified Fine (UNF) threads may be found wherever their use is appropriate, but special threads, such as UNS (for high-temperature applications) and UNJ (increased fatigue strength) have become more common. Screw threads may be formed, by such processes as tapping, dieing, and machine cutting or (where maximum fatigue resistance is required of a bolt), by rolling.


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12.5.2 OTHER THREAD FORMS

In the United States, a parallel but differing group of thread systems is used. The four main systems are ANC, ANF (also referred to as AF), UNC and UNF, with the NC and NF having a finer thread than the UNC and UNF. 12.5.3 CLASSES OF FIT

In addition to being identified as either coarse or fine, the threads are also classified by their class of fit, as can be seen in Table 8. Table 8 CLASSES OF THREAD FITS Class of Fit Type of Fit 1 Loose 2 Free 3 Medium 4 Close 5 Tight A Class 1 fit can be tightened, all the way down, by hand (such as with a wingnut), whilst a Class 4 or 5 fit requires a spanner throughout the tightening operation. The Class 3 fit is the type mostly employed on aircraft, and would be typical of a thread which is designed for use in a high-temperature environment and may require the application of an anti-seize compound before installation. By comparison, a fastener which is going to be subjected to the high tension or shear loads, associated with the securing of aircraft engine parts, would need to be a Close tolerance type of fit.


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12.5.4 MEASURING SCREW THREADS

It is not considered a normal operation to measure a screw thread, as its identification can be found in the IPC and supplied under a manufacturer’s part number. Whilst this is true and the manuals should always be used, there are other ways of identifying screw threads. One method is to identify the screw by means of various marks, normally found on the head of the screw. These marks may give a clue as to which type of thread the screw has (AF, BSF, or Metric etc.). A measurement across the thread crests, using a micrometer, would give the diameter of the screw in question. Finally, the identifying head markings would also give the material from which the screw is made. Two useful tools (refer to Fig. 42) may be used for different stages of thread measurement. The profile gauge can be used to ensure that the tool, which is cutting the thread, is of the correct type. The pitch gauge can be used to find the thread size by simply fitting the various blades of the gauge against the screw thread until a match is achieved Profile Gauge

55°

47½°

Pitch

Gauge 60° 0.25 – 2.5 mm 60°

Thread Profile Gauge and Screw Pitch Gauge Fig. 42


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12.6 BOLTS The bolts, used in the construction of aerospace components and structures, have evolved into a bewildering range of materials, shapes and sizes, all of which are dictated by the applications for which the items have been designed Standards and systems have been established, to provide identification of the many different forms of threaded devices, in order to ensure that only the correct items are installed in the relevant locations. It is stressed here, that only the approved design materials may be used for aerospace components and, while a selection of some of the bolts are presented in these course notes, by way of introduction, the relevant AMM, SRM and IPC will be the sole authority for deciding the correct type of bolt that is to be used in a particular application. 12.6.1 BRITISH BOLTS

An extensive range of bolts and screws is provided for, in the specifications drawn up by the Society of British Aerospace Companies (SBAC). The following abbreviations (some of which have, already, been discussed are in common use: • • • • • • • • • • •

AGS AS Al. Al. BA BSF. HTS. HTSS. LTS. SS UNC. UNF.

Aircraft General Standard Aircraft Standards Aluminium Alloy British Association British Standard Fine High Tensile steel High Tensile Stainless Steel Low Tensile Steel Stainless Steel Unified National Coarse Unified National Fine.

12.6.2 IDENTIFICATION OF BS UNIFIED BOLTS

British Standard Unified (BS Unified) bolts are identified by the use of an alphanumeric code, which provides information relating to the type, material, surface finish, length, diameter and any other important characteristics of the threaded device Table 9 shows a (very small) selection of aircraft standard bolts and screws with a (shortened) description of the type of device and the materials from which it is made. Reference to the table shows that the code A102 signifies a hexagonal-headed bolt which is made of high-tensile steel, while the code A175 represents a 100° countersunk-headed bolt, made from an aluminium alloy. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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Table 9 Examples of Code Numbers for Unified Threads Standard No. Description A102 Hex. Headed Bolt A104 Hex. Headed Bolt A111 Hex. Close Tolerance. Bolt A112 Shear Bolt A174 100º Countersunk. Head. Bolt A175 100º Countersunk. Head. Bolt A204 100º Countersunk. Head. Screw A205 Pan Head. Screw

Material HTS. SS HTS HTS SS Al Al HTS HTS

Other methods of indicating that an item has a Unified thread are: • Three contiguous (touching) circles marked in a convenient position (machine items). Note: Due to the difficulty in applying the identifying marks to individual items, it is planned to merely mark the packets in which the threaded devices are marketed, so that some, or all, of the identification marks will not be seen on the items (particularly screws). Great care must, therefore, be taken to ensure that the items being used are correctly identified and to the approved standard. • A shallow recess in the head of a bolt, equal to the nominal diameter of the thread (cold forged items). • A ‘dog point’ (small protrusion) on the threaded shank end (usually applies to screws). Further numbers and letters are added to the identifying code, to provide information relating to the length (usually of the plain shank or gripping portion) and to the diameter of the items. The length is given by a number, which signifies increments of tenths of an inch, so that a 5 would represent a bolt with a plain shank of 0.5 in, while the number 12 would signify the plain shank as being 1.2 in long Reference to Table 10, will show how the diameter of an item is designated by the addition of another letter to the system, so that a bolt, with the code marking of A102 9 E, would signify a Unified-threaded, hexagon-headed bolt, made from high-tensile steel, with a plain shank length of 0.9 in, and a diameter of ¼ in.


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Table 10 EXAMPLES OF BS UNIFIED BOLT CODES Diameter Code Diameter 0-80 UNF J 3/8" UNF (UNJF) 2-64 UNF L 7/16" UNF (UNJF) 4-40 UNC N 1/2" UNF (UNJF) 6-32 UNC P 9/16" UNF (UNJF) 8-32 UNC Q 5/8" UNF (UNJF) 10-32 UNF UNJF) S 3/4" UNF (UNJF) 1/4" UNF (UNJF) U 7/8" UNF (UNJF) 5/16" UNF (UNJF) W 1" UNF (UNJF)

Note: In the earlier UK system (which may be encountered on older, or homeconstructed, light aircraft), bolts more than ¼ inch diameter are normally BSF, whilst bolts less than ¼ inch diameter (and most screws) are BA. Both of these items also use a number to represent their nominal length and a letter code (as can be seen in Table 11) to identify their diameter. Other bolts of this era may have nicks at the corners of the head (High Tensile Steel) or a raised ring on the bolt head (Cold Rolled) to assist differentiation of their particular designations. Table 11 EXAMPLES OF BA AND BSF BOLT AND SCREW CODES Code

Size

Code

Size

A B C E G J L N

6 BA

P Q S U W X Y Z

9/16" BSF 5/8” BSF 3/4" BSF 7/8" BSF 1" BSF 12 BA 10 BA 8 BA

4 BA 2 BA 1/4” BSF

5/16" BSF 3/ 8" BSF

7/16" BSF 1/2" BSF


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12.6.3 AMERICAN BOLTS

American aircraft bolts and nuts are threaded in the NC (American National Coarse), the NF (American National Fine), the UNC (Unified National Coarse), and the UNF (Unified National Fine) thread series. The item is often coded to give the diameter of the threaded portion and the number of threads per inch (tpi). Aircraft bolts may be made from HTS, Corrosion-Resistant Steel or Aluminium Alloy. Head types may be hexagonal, clevis, eyebolt, internal wrenching and countersunk (refer to Fig. 43) and head markings may be used to indicate other features such as close tolerance, aluminium alloy, CRS or other types of steel.

Examples of Aircraft Bolts Fig 43


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12.6.4 IDENTIFICATION OF AN STANDARD BOLTS

While there are several different US Standards, there is only need to discuss one type for the purpose of these course notes, as the others are very similar. AN bolts come in three head styles, Hexagon Head, Clevis and Eyebolts and Table 12 provides an indication of the various code numbers in use. Table 12 EXAMPLES OF AN STANDARD BOLTS (EARLY SERIES) AN No. Type Material Process Thread Size 3 – 20

Bolt, hex. Head

Steel CRS Al. Al.

Thread Type

Cadmium Plated Nil Anodised

No. 10 to 1¼”

UNF

21 – 36

Bolt, Clevis

Steel

Cadmium Plated

No. 6 to 1”

UNF

42 – 36

Bolt, Eye

Steel

Cadmium Plated

No. 10 to 9/16”

UNF

73 – 81

Bolt, hex. Drilled head

Steel

Cadmium Plated

No. 10 to ¾”

UNF or UNC

173 – 186

Bolt, close - tolerance

Steel

Cadmium Plated thread & head

No. 10 to 1”

UNF

Note: The later series uses a different number system For identification purposes the AN number is used to indicate the type of bolt and its diameter. In addition a code is used to indicate the material, length and presence of a split pin or locking wire hole as follows: •

Diameter: The last figure, or last two figures, of the AN number indicates thread diameter, 1 = No. 6, 2 = No.8, 3 = No.10, and 4 = ¼” with subsequent numbers indicating the diameter in 1/16” increments. Thus an AN4 is a hexagon headed bolt of ¼” diameter and an AN14 is a hexagon headed bolt of 7/8” (14/16”) diameter.


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Lengths: The length of a bolt, in the case of a hexagonal headed bolt, is measured from under the head of the first full thread (refer to Fig. 44) and is quoted in 1/8” increments as a dash number. The last figure of the dash number represents eighths and the first figure inches, so that an AN4 – 12 is a ¼” diameter hexagon headed bolt, 1¼” long.

Drilled Shank CRS

Steel

Steel, Close Tolerance

Length ‘L’

CRS, Close Tolerance

Diameter

Aluminium Alloy

Drilled Head, AN 73 -81

Drilled Head, (Except AN 73 –81)

Grip

Aluminium Alloy, Close Tolerance

Head Markings for AN Bolts Fig. 44

•

Position of Drilled Hole: Bolts are normally supplied with a hole drilled in the threaded part of the shank, but different arrangements may be obtained: Drilled shank = normal coding e.g. AN24 – 15 Un-drilled shank

= A added after dash No.

e.g. AN24 – 15A

Drilled head only

= H added before dash No. (replacing dash) A added after dash No.

e.g. AN25H15A

Drilled head and shank

= H added before dash No.

e.g. AN25H15

• Material: The standard coding applies to a non-corrosion-resistant, cadmium-plated steel bolt. Where the bolt is supplied in other materials, letters are placed after the AN number as follows: • C

= Corrosion Resistance Steel C.R.S. e.g. AN25C15

• DD = Aluminium Alloy

e.g. AN25DD15


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Thread: Where the bolt is supplied as either UNF or UNC threads, a UNC thread is indicated by placing an A in place of the dash, e.g. AN24A15

12.6.5 SPECIAL-TO-TYPE BOLTS

The hexagon headed aircraft bolt AN3 – AN20 (refer to Fig.45), is an all purpose structural bolt used for applications involving tension or shear loads where a light drive fit is permissible.

Eye Bolt

Clevis Bolt

Special-to-Type Bolts Fig. 45 Alloy steel bolts, smaller than 3/16” diameter, and aluminium alloy bolts smaller than ¼” are not used on primary structure. Other bolts may be used as follows: •

Close Tolerance Bolts: These bolts are machined more accurately than the standard bolt. They may be hexagon headed (AN173 – AN186) or have a 100º countersunk head (NAS80 – NAS86). They are used in applications where a tight drive fit is required (the bolt requires the use of a 340g - 400g (12oz – 14 oz) hammer to drive it into position.

•

Internal Wrenching Bolts: (MS 20024 or NAS 495) these are fabricated from high-strength steel and are suitable for tensile or shear applications. The head is recessed to allow the insertion of a hexagonal key used for installing or removing the bolt. In Dural-type material, a heat-treated washer must be used to provide an adequate bearing surface for the head.


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•

Clevis Bolts: The head of a clevis bolt is round and either slotted, for a standard screwdriver, or recessed, for a cross-pointed screwdriver. This type of bolt is used only for shear loads and never in tension. It is often inserted as a mechanical pin in a control system.

•

Eyebolt: The eye is designed for the attachment of cable shackles or turnbuckles and the bolt is used for tensile loads. The threaded end may be drilled for ‘safetying’.

12.6.6 METRIC BOLTS

The identification of a Metric bolt is by the use of the diameter in millimetres, immediately after the capital letter ‘M’. In this way, M6 represents a 6 mmdiameter bolt. The length is also shown in millimetres, so the bolt M6 -15 will be a 6 mmdiameter bolt, which is 15 mm long. The basic terminology, for identifying bolts of the Metric system, involves the nominal length, the grip length and diameter (refer to Fig. 46). Length

Grip

Diameter

Metric Bolt Terminology Fig. 46


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12.7 NUTS Aerospace standard nuts are made in a variety of shapes and sizes. They can be made of cadmium-plated carbon steel, stainless steel or anodised 2024–T aluminium alloy and can have right- or left-hand threads (refer to Fig. 47). As a general rule, nuts are manufactured from the same material as the bolt or screw to which they are attached, with the exception of high-tensile steel bolts, with which, mild steel nuts are used.

Selection of Typical Nuts Fig. 47 As they do not have any identifying marks or lettering, they are usually identified by their colour and their constructional features. Familiar types of nuts include: • Castle Nuts: which are used with drilled shank hexagon-headed bolts or studs, eye-bolts and clevis bolts. They are fairly rugged and can withstand large tensile loads. The slots (castellations) are designed to accommodate a split (cotter) pin. • Slotted Nuts: are similar in construction to the castle nuts and are used in similar applications, except that they are normally used for engine use only. • Plain Hexagon Nuts: are of rugged construction and suitable for large tensile loads. Since they require an auxiliary locking device, their use on aircraft is limited. • Light Hexagon Nuts: are a much lighter nut, used for miscellaneous light tensile requirements. • Plain Check (or Lock) Nuts: are employed as locking devices for plain nuts, for threaded rod ends and for other devices. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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• Wing Nuts: are used where the desired tightness can be obtained merely with using the fingers and where the assembly is frequently removed. 12.7.1 STIFFNUTS AND ANCHOR NUTS

An ordinary standard nut will depend upon friction between the engaging threads to ensure its tightness. The enemy of this system is vibration, which can cause the nut to slacken off, and in extreme cases, unwind itself completely from the bolt or screw. In areas where this might occur, locking devices are used. These either increase the frictional resistance between the threads, or take the form of positive securities that prevent any movement of the nut once they have been applied. Stiffnuts and anchor nuts (refer to Fig. 48) employ various means of increasing the friction forces between the threaded devices and common types include: • Nyloc: This looks like a standard hexagonal nut, but has a plastic insert in the counter-bored end. This insert is initially unthreaded and has an internal diameter slightly smaller than the nut thread, so that, as the nut is screwed on the bolt, the plastic insert is displaced and a high degree of friction is created. Another type of plastic ‘stop’ nut is named the ‘Capnut’. This type is completely sealed and is used in pressurised compartments and fuel and oil tanks etc. Note: As the insert is nylon, this type of stiffnut should not be used in high or low temperature areas. A typical maximum temperature would be 120ºC. A similar type of stiffnut has a fibre insert instead of nylon, and is called a ‘fibrelock nut’. Neither nylon nor fibrelock stiffnuts should be re-used. • Oddie: The top of this nut has a slotted end, consisting of six tongues, which form a circle slightly smaller than the bolt or stud diameter. As the nut is turned, a friction load is imparted onto the threaded device. • Philidas: This nut has a circular crown which is slotted horizontally in two places The thread on the slotted part is slightly ‘out of phase’ with the rest of the thread, so that increased friction is achieved when the nut is turned. • Aerotight: Similar to the Philidas in appearance, except that the slots are vertical. Its locking method is also similar. • Lightweight: The locking section of this stiffnut is slightly oval in shape and so causes increased friction when the thread passes through it. Note: Metal hexagonal type stiffnuts may be re-used, provided they are not being used in vital areas such as flying controls and they retain their friction effect. A recognised rule for serviceability is that they are discarded when they can be screwed all the way down, on a new bolt, using only the fingers. • Anchor nuts and Stripnuts: Anchor nuts are supplied with single or double attachment points and may be either fixed or floating in a cage.


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The anchor nut may be a single unit stiffnut, integral with the base plate, or it may be an assembly, comprising stiffnut, cage and base plate. Single attachment types are used in corners or where space is limited and have two adjacent fixing points. Double anchor nuts have a hole either side of the stiffnut. They are fitted to the structure by riveting. Where a number of anchor nuts are required, to secure panels etc. a number of stiffnuts may be fitted into metal strips for ease of securing. Stripnuts are usually of the floating variety.


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Stiffnuts and Anchor Nuts Fig. 48


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12.8 SCREWS Screws are, probably, the most commonly used threaded fastener in aircraft construction. They differ from bolts in that they are generally made from lowerstrength materials. They can be fastened by a variety of tools, including screwdrivers, spanners and Allen keys. Most screws are threaded along their complete length, whilst some have a plain portion for part of their length. There are a number of different types of screw, which, can be used for a wide range of tasks. It is common sense that great care must be taken to replace screws with the correct items, by using the markings on the screw, the IPC and any other systems in current use within the supply department, to protect against incorrect screws being installed. Another point, requiring care, is the difference in terminology between the British and American names for screw heads. What the British refer to as a ‘countersunk -headed’ screw, the Americans call a ‘flat-head’ or ‘flush’ screw. Similarly, ‘mushroom-headed’ screws are known as ‘truss-heads’ in the USA. 12.8.1 MACHINE SCREWS

Machine screws (refer to Fig. 49) are used extensively for attaching fairings, inspection plates, fluid line clamps and other light structural parts. The main difference between aircraft bolts and machine screws, is that the threads of a machine screw usually run the length of the shank, whereas bolts usually have an unthreaded grip length. The most common machine screw used in aviation is the fillister-head screw, which can be wire-locked using the drilled hole in the head. The flat-head (countersunk-head) screw is available with single or cross-point slotted heads. The round-head screw and the truss-head (mushroom-head) screw, provide good holding properties on thin metal sheets.


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A Selection of Machine Screws Fig. 49


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12.8.2 STRUCTURAL SCREWS

Structural screws (refer to Fig. 50) are made of alloy steel, are heat-treated and can be used in many structural situations. They have a definite grip and the same shear strength as a bolt of the same size. They are available with fillister, flat or washer heads. The washer head screw has a washer formed into its head to increase its holding ability with thin materials, much like the truss or mushroom head. 100°

Grip

Grip

Length

Diameter

Diameter

Structural Screw Terminology Fig. 50 12.8.3 SELF-TAPPING SCREWS

Self-tapping screws (refer to Fig. 51) have coarse threads and are used to hold thin sheets of metal, plastic and plywood together. The type A screw has a gimlet (sharp) point, and the type B has a blunt point with threads that are slightly finer than the type A. There are four types of head in normal use: • round head • countersunk oval-head • truss or mushroom-head • flat countersunk-head.


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Four Types of Self-Tapping Screw Heads Fig. 51 12.9 STUDS Studs are metal rods that are threaded at both ends (refer to Fig. 52). In general they are used where it is not possible, or desirable for a bolt to be used. Like many screw types of fastener, most studs are produced in a standard form, with variants used for special purposes. For example, where a standard type is unsuitable, such as when being used in a soft metal, then a stepped stud (which has a greater holding power) would be used. A stepped stud would also be used where a damaged thread had been removed, the hole drilled out and re-tapped. It will be appreciated that the security of a stud depends upon the friction between its thread and that of the tapped hole (the ‘metal’ thread) into which it is inserted. If this friction fails to hold the stud, it will work loose and all precautions to prevent the nut from slackening will be negated.


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Typical Studs Fig. 52 12.9.1 STANDARD STUDS

By far the most widely used stud is the standard (plain, or parallel) type, in which the diameter of the whole stud, along its length, is constant. Standard studs are classified by the thread type, diameter and overall length. The ‘metal’ thread is, usually, finished very slightly oversize to give a tight fit into the tapped hole. Other variants of the standard stud are available for use in circumstances that require special consideration. To meet special requirements, the various types of standard studs may also be supplied with non-standard lengths of plain portion and ‘metal’ end. A simple method of fitting and removing a stud is by running two plain nuts down the ‘nut’ end of the stud and cinching (locking) them together using two spanners. The stud can then be screwed into or removed from the material. Breaking the cinch then separating and removing the nuts completes the operation. 12.9.2 WAISTED STUDS

Waisted studs are used where reduction of weight, without the loss of strength, is of paramount importance. The diameter of the plain portion of the stud is reduced to the minor diameter of the end threads, thus lightening the stud without impairing its effective strength.


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12.9.3 STEPPED STUDS

This type affords a stronger anchorage than the standard type, if the ‘metal’ end of the stud has to be housed in soft metal. The thread of the ‘metal’ end is one size larger than that of the ‘nut’ end. For example, a ¼ inch BSF stepped stud has a plain portion of ¼ inch thread on the ‘nut’ end and a 3/16-inch thread on the ‘metal’ end. Stepped studs are also used as replacements for standard studs when the tapped stud-hole has to be re-drilled and tapped with a larger thread, due to damage. 12.9.4 SHOULDERED STUDS

This type is used where maximum rigidity of assembly is of prime importance. The stud is machined from oversize bar and a projecting shoulder is left between the ‘metal’ end of the thread and the normal diameter plain portion. This shoulder seats firmly on the surface of the ‘metal’ and gives additional resistance to sideways stresses. The clearance hole in the second component, through which the ‘nut’ end and plain portion of the stud passes, must be machined at the inner end to give clearance to the stud shoulder. 12.10 THREAD INSERTS Thread inserts are a means of providing a stronger anchorage, for bolts, screws or studs, in the comparatively softer metal alloys (aluminium, magnesium, bronze), wood, plastics or composite materials. They may also be used when it is necessary to do a repair to a threaded hole that has suffered damage. There are two basic types of thread insert (Wire and Thin Wall), but the designs of each type will vary according to the many manufacturers or to the environment in which the fastener must operate. 12.10.1

WIRE THREAD INSERTS

Wire thread inserts consist of a very accurately formed helical coil of wire, which has a diamond (rather than a round) cross-section and is usually made from corrosion-resistant steel or heat-resistant nickel alloy. Specifically sized drills, taps and thread gauges (provided by the insert manufacturer) are required to form the tapped holes for the inserts and another special tool is necessary to insert the wire coils correctly into their prepared holes.


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THIN WALL INSERTS

Thin wall inserts appear in a variety of designs, materials and surface finishes and consist of a thin tube, which is threaded internally and may, or may not, be threaded externally. Similarly, special tools are required from the manufacturer to prepare the holes for the inserts and various methods are adopted to secure each particular type of thin wall insert into its hole. Thin Wall inserts include: • Key-Locked Inserts: Key-Iocked inserts are threaded both internally and externally and, after being screwed into the prepared hole, are (as their name implies), locked into their holes by tiny wedges or keys. The keys are then pressed (or hammered) into place between the insert and the wall of the hole. • Swaged Inserts: Swaged inserts are also threaded internally and externally and are, again, screwed into the hole before a tool is used to deform (swage) the insert so that it is locked into the hole. • Ring-Locked Inserts: Ring-Iocked inserts, with internal and external threads, are screwed into holes which are counter bored, to allow a special lock-ring to be installed, (after the insert) and yet another special tool is used to complete the locking action of the lock-ring. • Bonded Inserts: Bonded inserts are, usually, only internally threaded (to hold the bolt, screw, stud etc) and are secured in the prepared hole by the use of adhesives. Obviously, from this information, it can be seen that great care must be taken to ensure that only the approved types of inserts are used in aerospace components and that the procedures for their installation and removal (laid out in the relevant Manuals) are carefully followed.


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12.11 DOWELS AND PINS Dowels and pins used in aircraft can include the Roll Pin, Clevis Pin, Split (Cotter) Pin, and Taper Pin. 12.11.1

DOWELS

While not usually used as fasteners, dowels are rods or pins of the appropriate material which are fixed (often permanently) in one of the components of a joint such that the protruding shank of the dowel locates with a corresponding hole in the item being attached, thus ensuring accurate assembly. Two examples of the use of dowels may be found where a Propeller Control Unit is attached to an engine casing and there is a requirement for absolute accuracy in the alignment of the oil tubes and, again, where the segments of an engine compressor need to be joined with precision so that the rotating members do not foul the stationary parts. 12.11.2

ROLL PINS

Roll pins (refer to Fig. 53) are often used to secure a pulley to a shaft or to provide a pivot for a joint where the pin is unlikely to be removed. A roll pin is normally made from flat spring steel that is rolled into an incomplete cylindrical shape that allows the pin to compress when it is pressed into the hole, and creates a spring action that holds the pin tight within the bore of the hole. To remove a roll pin it must be driven from the hole with a correct-sized punch.

Roll Pin Fig. 53


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CLEVIS PINS

Clevis or flat-head pins (refer to Fig. 54) are used for hinge pins in some aircraft control systems. They are made of cadmium-plated steel and have grip lengths in 1/16-inch increments. When a clevis pin is installed, a plain washer is usually placed over the end of the shank and a cotter (split) pin is inserted, through the pre-drilled hole in the clevis pin, to lock it in place.

Diameter Length Clevis Pin Fig. 54

12.11.4

TAPER PINS

Both the plain and threaded taper pins (refer to Fig. 55) have a taper of 1 in 48 and are used in various locations during aircraft construction. They are designed to carry shear loads and are manufactured from high-tensile steel. The pins do not allow any loose motion or play and are used for joining tubes and attaching collars to shafts. The plain taper pin is forced into the hole, which is reamed to the specified size with a Taper Pin Reamer, and is held in place by friction alone. To ensure security, it can also be wire locked in place, by passing the lock wire through the pre-drilled hole in the pin then securing the wire around the shaft. Plain taper pins, which have no lock wire holes, may have their smaller ends peened, after being installed, to secure them in their holes. The Threaded Pin is similar to the plain pin except that its small end is threaded to accept either a self-locking shear nut or a shear castle nut with split pin. Some taper pins can be found with a split small end, which can be spread much like a split pin, to prevent it loosening. These pins are sometimes referred to as bifurcated taper pins.


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All taper pins are measured by the diameter of their small end and their overall length.

Plain Taper Pin

Threaded Taper Pin Taper Pins Fig. 55 12.12 LOCKING DEVICES The problems associated with threaded devices, and the effects of vibration on their security, were discussed previously, when the use of stiffnuts and anchor nuts was considered. In addition to using methods which increase the friction between threads, there are several other ways in which the integrity of a threaded joint can be assured. 12.12.1

SPRING WASHERS

These washers are available in a variety of forms (refer to Fig. 56). In some instances (particularly with light alloy assemblies), spring washers are assembled with plain facing washers between the spring washer and the component. This is done to prevent damage to the surface finish when the spring washer is compressed although, with steel assemblies, the plain washer is usually omitted. It is good practice to renew spring washers during overhaul or repair. This procedure is most essential in engines and engine components as well as where units have reciprocating parts; such as in compressors or pumps. In normal circumstances, however, spring washers can be re-used if they have retained their ‘springiness’ and ‘sharpness’. Types of spring washers include: • Single and Double Coil Washers: Manufactured from rectangular-sectioned steel sheet and formed into a portion of a helix, the single and double coil are the most common types of spring washer to be found on aircraft components


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Single Coil


Double Coil

Crinkle

Cup

Spring Washers Fig. 56 • Crinkle Washers: Crinkle washers are usually manufactured from either copper alloy or corrosion- resistant steel. They are often used in lightly loaded applications such as instruments and electrical installations. • Cup Washers: Cup (or Belleville) washers are manufactured from spring steel and are ‘dished’ to form a spring of high rating. The flattening of the washer, during tightening, exerts an axial load to the nut, which will resist any tendency of the nut to lose torque. Assembly should always be in accordance with the manufacturer’s instructions. 12.12.2

SHAKE-PROOF WASHERS

Flat washers of this type (refer to Fig. 57), are manufactured from steel or phosphor bronze and are used in place of spring washers. In some circumstances conical shake-proof washers are used for locking countersunk screws. Either the internal or the external diameters can be serrated, the serration being designed to bite into the component and nut to prevent rotation. All shake-proof washers should be used only ONCE. It is rare for these washers to be specified in assemblies where an anti-corrosion treatment of the components has been specified, as this could damage the treatment.


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Shake-proof Washers Fig. 57

12.12.3

TAB WASHERS

Tab washers (refer to Fig. 58), are normally used on plain nuts. The washers are manufactured from thin metallic sheet material and have two or more tabs projecting from the external diameter. They can also be designed for locking two or more nuts. When the washer is installed, one tab is bent against the component or inserted into a hole provided, whilst a second tab is bent against the flat (or flats), of the nut, after it has been torqued down correctly. Note: Multi-tab washers can be re-used until all tabs have been used once.

Tab Washers Fig. 58


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LOCK PLATES

In certain circumstances, the torque applied, the thread, or the type of nut, being used may not guarantee that the nut would not unwind in use (such as during vibration). Lock plates (refer to Fig. 59) are used where positive retention of a nut is required.

Lock Plate Fig. 59 The nut is torque loaded and then (only if necessary) turned a small amount, (< 1/12 revolution) until its flats align with the hole in the lock plate. The plate usually has 12 facets to allow for this adjustment. The plate is then placed over the nut and the small setscrew fastened into the tapped hole adjacent to the nut. Removal of the nut simply involves removing the setscrew, lifting off the plate and unwinding the nut. Note: A Tab washer could be used to do the same task. The lock plate is used where the nut is frequently removed – the plate can be used indefinitely providing it retains a good fit with the nut. 12.12.5

SPLIT (COTTER) PINS

These pins (refer to Fig. 60) are usually manufactured from either cadmiumplated carbon steel or from corrosion-resistant steel. Their primary purpose is to lock slotted and castellated nuts as well as for securing clevis pins. The nuts are locked onto their bolts by passing the pin through the hole in the bolt and the nut castellations. The legs of the pin are spread in one of two methods. Whilst either of these methods will secure the nut to the bolt, different airworthiness authorities prefer one method to the other. The pins are measured by diameter and length. It must be noted that the nuts must never be over-torqued to get the holes into line. The nut must either be backed-off, if this is permitted, or washers added under the nut. Often a stated torque value will be over a small range rather than a set figure. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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This allows very small movement of the nut to facilitate alignment of the locking pins. Details of the correct method for each task will be in the AMM.

Two Methods of Securing Split Pins Fig. 60 12.13 LOCKING WIRE Wire-locking (or ‘Safetying’ as it is known in the USA), is the commonest form of locking in use throughout the aircraft industry. The wire is usually made of corrosion-resistant steel or heat-resistant nickel alloy. Fine copper wire is also used for some special locking operations. The wire is normally classified by its diameter in increments of ‘Standard Wire Gauge’ (SWG) or ‘American Wire Gauge’ (AWG). The most usual gauge used is 22 SWG (or its American equivalent), although great care must be taken to check the correct wire gauge for each particular application. Wire-locking is a positive method of securing items such as bolts, pipe unions, turnbuckles and nuts. Components designed to be wire-locked have holes in the appropriate positions to enable the lock wire to pass through. When installing the wire it should not span a distance of more than 75 mm (3 in) without being supported. The wire is also positioned so that the item being locked will be restrained from turning in a loosening direction. There should be approximately eight turns to every 25.4 mm (1 in) length of wire and no length of more than 9.5 mm (3/8 in) should be left untwisted. The angle of pull, or approach (refer to Fig. 61), should be not less than 45° to the rotational axis. When the wire has been passed through the last hole, the wire must be pulled tight and the twisting continued for at least 12 mm – 13 mm (½ in). The wire is then cut and the end doubled under, to prevent personnel getting ‘snagged’ or badly cut.


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Wire-locking Angles Fig. 61 Some forms of wire-locking are done with a single strand of the specified wire, especially in cases of where a complete ring or similar formation of nuts is found (refer to Fig. 62). The wire is passed in sequence, through the holes in their respective nuts and bolts (or screws), until the wire ends meet. Again the wire must be threaded so that any tendency, of a nut or bolt, to attempt to slacken off, will add tension to the wire.

Single Wire ‘Safetying’ for Closely-Spaced, Drilled-Head Screws Fig. 62


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USE OF LOCKING WIRE WITH TURNBUCKLES

As with any threaded fastener, turnbuckles must be locked to prevent them from coming loose and jeopardising the control runs they are connecting. There are a number of different types of wire-locking used on turnbuckles and the AMM must be consulted to find which method is specified. Methods used include the single wrap and single wrap spiral as well as the double wrap and double wrap spiral. The single wrap and single wrap spiral use a single strand of the appropriate wire that passes through the hole in the centre of the turnbuckle, finishing up wrapped around each end. The single wrap spiral also uses a single piece of wire that is spiralled around the turnbuckle barrel and passed through the centre hole twice. Two pieces of wire are used in the double wrap method, which are basically two single wraps, one in each direction. A double wrap spiral consists of two single wrap spirals, again one in each direction. 12.13.2

USE OF LOCKING WIRE WITH LOCKING TABS.

When locking tabs are used, they should be installed in such a way that the tabs and the wire are in complete alignment (refer to Fig. 63). Whenever possible, the closed end of the wire should be in the tab and the twisted end at the component to be locked, although the exact method may be found in the AMM.

Locking Wire and Locking Tabs Fig. 63


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THIN COPPER WIRE

Thin copper wire is used to hold some switches and levers in a ‘set’ position and, thus, prevents the accidental operation of those switches which control certain critical systems such as emergency circuits. When the switch is required to be operated, then a deliberate movement is made, which will break the copper wire and permit movement of the switch. A secondary purpose of copper wire is as an indicator or ‘witness’, where a broken wire indicates that the switch or control has been operated. This method is employed on systems where it is necessary to know when a system has been operated (such as in a Fire Protection system). 12.14 QUICK-RELEASE FASTENERS Special fasteners have been designed to hold fairings, cowlings and inspection panels in position and to allow their rapid removal and replacement during servicing. 12.14.1

DZUS FASTENERS

Cowling and other inspection access doors will usually be found with Dzus fasteners, that can be locked and unlocked by a quarter turn of the stud (refer to Fig. 64). These fasteners consist of a hard spring-steel wire, which is riveted across an opening on a fixed part of the airframe. The stud is mounted onto the panel (or removable part), using a metal grommet. When the panel is closed, a quarter turn of the stud pulls the wire into the curved slot of the stud, securing the panel to the airframe. Panels (and cowlings) usually have a number of fasteners installed to ensure full security and, to indicate that all fasteners are correctly secured, the cowling will have a series of lines marked (painted) on the surface. When the studs are correctly fastened, then their screwdriver slots will be in-line with the lines marked on the surface of the panels. Some Dzus fasteners have a built-in receptacle, which guides the legs of the stud onto the wire, to facilitate correct engagement.


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Dzus Fastener Construction Fig. 64 12.14.2

ODDIE FASTENERS

Oddie fasteners (refer to Fig. 65) have a central stud, which is held in position in the panel with a rubber washer or a coiled spring. A two-legged clip is fastened to the fixed component (usually with rivets). The stud is bullet-shaped and has two recesses opposite each other at the joint end. The fastener is locked by positioning the recess in line with the legs of the spring, and then pressing the stud home. This is achieved by ensuring the screwdriver slot is in line with marks on the panel. There should be a definite click as the fastener engages. A quarter turn of the stud will release it from the spring, and free the panel.


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Oddie Fastener Construction Fig. 65 12.14.3

CAMLOC FASTENERS

Camloc fasteners (refer to Fig. 66) consist of a spring-loaded stud assembly and a receptacle. The stud assembly is fastened to the removable panel whilst the receptacle is fastened to the airframe. To lock the fastener, the stud is pushed against its spring with a screwdriver and given a quarter of a turn clock-wise. As a result, the cross-pin, on the stud, rides up a cam in the receptacle and draws the two components together. Finally the stud spring pulls the cross pin into a locking groove at the end of the cam. The fastener is unlocked by a quarter turn anti-clockwise when the stud spring causes the stud to snap outwards.

Camloc Fastener Fig. 66


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AIRLOC FASTENERS

Airloc fasteners (refer to Fig. 67) consist of a stud with a cross-pin in the removable cowling or door, and a sheet spring-steel receptacle in the structure. The fastener is locked by turning the stud through a quarter turn. The pin drops into an indentation in the receptacle and holds the fastener locked.

Cross Pin Installed Pin

Studs Stud Receptica l

Airloc Fastener Fig. 67

12.14.5

PIP-PINS

Quick-release ‘Pip-pins’ are used in assemblies where it is necessary to rapidly remove or reposition components. They usually take the place of more permanent bolts. The ‘pip-pin’ quick-release fastener (refer to Fig. 68) operates on a push-pull principle. It consists of a hollow body containing a spring-loaded plunger. When the pin is pushed into a hole, two steel locking balls, held in the shank of the pin, EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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move into a recess in the plunger. When the pin is fully home, and the pushing pressure is released, the balls are forced to protrude from the shank, as the spring around the plunger expands, and so lock the pin in position. A pip-pin is removed by a simple pull on the ring. This action aligns the groove in the plunger with the two locking balls that retract to allow the pin to be withdrawn. Pip-pins will be found in many places where two components have to be separated at regular intervals and also require a hinging action. An example of where pip-pins would be required is on engine cowlings. These have to be opened daily to allow for engine inspection, and are removed completely for engine changes.

Pip Pin Locking Balls

Pin Release Ring Typical Pip-Pin Fig. 68 12.14.6

CIRCLIPS AND LOCKING RINGS

Circlips and locking rings (refer to Fig. 69) are manufactured from spring sheet metal or spring steel wire, They may also be specially designed for a particular purpose. Hardened and tempered to give either and ‘inward’ or ‘outward’ spring, they can be used for locking several parts together, locating components within bores or for locating components onto shafts. Spring sheet circlips have holes in the ends to allow circlip pliers to be inserted, enabling the circlip to be removed or installed as required. Spring wire rings usually have one bent end that is inserted into a radial hole, drilled through the component, which matches an inner or outer ring. All circlips are subject to some damage at times and it will usually be a requirement, after they have been removed, to inspect them thoroughly. Any that show damage or corrosion should be discarded, although it is usual practice to discard the wire type circlips whenever they are removed


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External

Internal

External

Spring Sheet Plate Type Internal

External

Spring Wire Type Squeeze legs together, or expand, to remove the rings Circlips and Locking Rings Fig. 69

12.14.7

KEYS AND KEYWAYS

These items can be found where chain-wheels or pulleys are located on shafts. A key, with its associated keyways (the name given to the channel, which is cut into the respective components, to receive the key), is used to transmit the driving force from one part to the other. There are different types of keys and keyways, and these will be covered in greater depth in the section on transmissions.


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PEENING

Peening (refer to Fig. 70) is a method of preventing a threaded device (bolt, nut or screw), becoming loose by distorting the end of the thread, after installing the device. The distortion is normally achieved (using a centre punch) by striking the thread of the bolt or screw where it emerges from the threaded device, thus jamming and effectively locking the threaded device and preventing it from loosening. When using a nut and bolt combination, then one and a half threads of the bolt must protrude from the nut in order to create an effective peening. The disadvantage of peening (and the distortion of the thread) means that, once the joint is dismantled, then the threaded device is useless and can only be discarded.

Peened (Burred)

Metal Peened (Burred) into Slot of Fastener Fastener

Peened (Burred) Fasteners Fig 70


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12.15 GLUE/ADHESIVE BONDED JOINTS As was previously discussed (in the section on Composite and Non-Metallic Materials), these are permanent joints in which an adhesive is used to join two, or more, materials together. The materials can be any of the large variety of fabrics found in the aerospace industry (metal, paper, plastic, rubber or wood). Some advantages of using adhesives, to make joints, are that the materials being joined may or may not be similar and the joints can be made proof against the leakage of gases and liquids. Adhesives are normally good electrical insulators, which can greatly reduce dissimilar corrosion on metal joints, and are not, normally, affected by temperature changes. Joining with adhesives not only saves the weight (and costs) associated with threaded fasteners (and rivets), but also eliminates the need to make holes in the structure, for those fasteners, which avoids the possibility of potential stress raisers. The absence of fasteners in an aircraft’s skin results in a smoother airflow around the aircraft, and thus contributes to its aerodynamic efficiency. Adhesive bonded joints also provide greater stiffening to the structure, compared to that achieved with mechanical fastenings. There are, however, some disadvantages in that the surfaces, of the items to be stuck together (the adherends), must be free from grease, oil or dust, and the type of adhesive must be suitable for the conditions or environment in which it is intended to be placed. Fumes from adhesives can be narcotic, toxic and extremely flammable, so that great care must be taken when applying adhesives. This entails working in wellventilated conditions, wearing the appropriate personal protective equipment and observing the relevant safety precautions to prevent (and, if necessary, fight) the outbreak of fire. 12.15.1

LOCKING BY ADHESIVES

Applying Shellac, Araldite etc to DTD 900 specification, may be used to lock many small components, particularly those in instruments, valves, switches etc. Adhesive is applied to the outside of the nut face and the protruding screw thread, or to the component and screw head, after tightening, and prevents movement between relevant parts. It is good practice, when using Araldite, to mix a separate sample under similar conditions, to check that it hardens within the specified time period. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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LOCTITE

Loctite is the trade name for a liquid sealant, used approved, proprietary material, which hardens assembly. Loctite is supplied in various grades to strength in a variety of applications from stud housings.

to lock metal threads. It is an in the screw threads after give a predetermined locking locking to retaining bearing

When using Loctite, it is advisable to degrease the parts to achieve maximum strength. If the threads are not degreased, about 15% of the locking strength is normally lost. Loctite should only be used when specified by the approved drawings or instructions, and applied in accordance with the manufacturer’s directions 12.15.3

SYNTHETIC RESIN ADHESIVES

Synthetic resin adhesives are used extensively for joints in wooden structures, to avoid the localised stresses and strains, which may be set up, following the use of mechanical methods of attachment. Synthetic resin adhesives, used for gluing aircraft structural assemblies, must comply with the requirement prescribed in an acceptable specification Synthetic resin adhesives usually consist of two separate parts, namely the resin and the hardener. The resin develops its adhesive properties only as a result of a chemical reaction between it and the hardener. 12.15.4

TESTING OF ADHESIVE JOINING TECHNIQUES

Frequent tests would be made to ensure that joining techniques are satisfactory. Whenever possible, tests should be done, using off-cuts of actual components from each batch. Where off-cuts are not available, tests should be done on representative test pieces. 12.16 METAL-TO-METAL BONDED JOINTS Metal-to-metal joints involve the use of heat, to raise the temperature of the metals to a point where, either by the use of hammering, by the application of pressure, or by a chemical reaction between the metals being joined, the metals fuse together and thus create the required bond. 12.16.1

WELDING

Welding is the fusing together, by heating the point or edge of contact of two or more pieces of metal (and applying a filler rod if required), making one continuous piece. Welded joints are normally considered to be part of an aircraft’s permanent structure and they would not be dismantled during routine maintenance. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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Only a trained welder, authorised by the CAA, may weld component parts of a British-registered civil aircraft and that person is required to submit, to the CAA, a series of test welds, for examination, every twelve months. It is, therefore, beyond the scope of these course notes to consider the various forms of electric arc, gas, resistance, seam or spot welding techniques. 12.16.2

SOFT SOLDERING

Soft soldering is the permanent joining of metals, using a filler metal that melts at a temperature considerably lower than the metals being joined. The filler metal is an alloy consisting, mainly, of lead and tin (with, possibly, antimony and bismuth), mixed in varying proportions, depending on the use for which it is intended. To ensure a satisfactory joint, the solder must form a metallic bond ('key') with the surfaces, being joined and, to allow this to happen, the joint surfaces must be free of oil, grease, dust, and corrosion. It is also necessary to use of an approved substance (a ‘flux’), which is applied to the metals, to prevent the formation of potentially corrosive oxide films while the metals are being heated (usually by conduction of the heat from a soldering ‘iron’) and joined. 12.16.3

HARD SOLDERING

Hard soldering includes Silver Soldering and Brazing. In these processes, the fillers melt at higher temperatures than soft solder and provide a much stronger joint, which is also capable of operating at higher temperatures. Silver Solder consists of an alloy of copper and silver (with a melting point almost twice that of the soft solders) while Brazing uses a copper-zinc alloy with a melting point higher than that of Silver Solder. The source of heat used for hard soldering is, usually, a direct flame and a different flux is also necessary to prevent oxidation of the joint. Hard soldered joints have their fillers drawn into them by capillary action, therefore the gap between components must be kept uniform and closely controlled. As with all soldered joints, the surfaces being joined must be clean and free of oil, grease, corrosion, scale etc. Mechanical methods of cleaning can include emery cloth, wire brush or filing.


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13 AIRCRAFT RIVETS An aircraft, even though made of the best materials and strongest parts, would be of doubtful value unless those parts were firmly held together. Several methods are used to hold parts together; welding or soldering, threaded fasteners and riveting being three of the main methods. The use of threaded fasteners, and soldering, has been mentioned previously. Rivets are an alternative method of fastening structure, a rivet being a metal pin on which a head is formed, during manufacture. The rivet is inserted into a predrilled hole and the plain end of its shank is deformed (‘set’ or ‘closed’) by the use of a hand- or power-tool. Rivets create a joint at least as strong as the material that is being joined. Rivets are normally strong in shear, but they should not be subjected to excessive tensile loads. There are two main categories of rivet: • Solid rivets: which are ‘set’ using a riveting gun on the manufactured head and a reaction (bucking) bar on the remote side • Blind rivets: which may be installed where access is restricted to the shank end of the rivet. 13.1 SOLID RIVETS There are a number of different types of rivet head, the most common being the mushroom and round heads (refer to Fig. 71). Both of these rivets project above the surface of the metal that is being riveted. The countersunk head, however, provides a flush and smooth surface, when closed and the flat (or pan) head can be used internally, when a flat head will make closing the rivet easier.

Length

Diameter

Mushroom

Pan

Round

Types of Solid Rivets Fig. 71


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The majority of aircraft rivets are manufactured from aluminium alloys. Rivets can also be made from other materials such as steel, Monel metal, titanium or copper. Material specifications for British and American rivets are not identical. The manufacturer’s publications (AMM or CMM) will give details on which rivets can be used if the specified ones are unavailable. The dimensions that identify the size of a rivet are simply its length and diameter. Other identifying features are the shape of the head, (including the countersink angle, if applicable) and the material from which the rivet is made. This latter requirement involves many different identifying marks and letters. 13.2 RIVET IDENTIFICATION The identification of solid rivets covers a multitude of marks and letters that indicate not only the material, but also the heat treatment, (if any), that the rivet has gone through. The American rivets are, usually, ‘natural’ (gold) or grey in colour and have head markings, whilst British rivets, generally, use a combination of colour and alpha/numeric codes. 13.2.1 SOLID RIVETS (BRITISH)

Standards for British Solid rivets are issued by the Society of British Aerospace SBAC (AS series) or the British Standards Institute (SP series). The standards overlap to a certain extent, with obsolete rivets, in the AS range, being replaced by SP rivets.


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13.2.2 RIVET IDENTIFICATION (BRITISH)

A standard number and a part number are used to identify rivets. The standard number identifies the head shape, material and finish. This is followed by a three or four figure code, the first one or two figures indicating the shank diameter in thirty-seconds of an inch and the last two, the length in sixteenths of an inch. Example: A British rivet, with the identifying code AS 162-408, would be a 90° countersunk, aluminium alloy (5% magnesium) rivet, of 1/8 inch diameter and 1/2 inch long. The AS 162 indicates the head type and material, while the ‘-4’ indicates that it has a 4/32 inch (1/8 inch) diameter and ‘08’ indicates it has a length of 8/16 inch (1/2 inch). 13.2.3 RIVET MATERIAL IDENTIFICATION (BRITISH)

Tables 13 and 14 give details on materials and identification marks for the various types of AS rivets. Many of these rivets are obsolescent and have been superseded by rivets conforming to SP standards. Table 15 gives details of material and identification information for SP rivets with the standard numbers shown in Table 16. SP rivets are also available in metric sizes. Note: The colour coding (of both British systems) of solid rivets is generally the same as that used for the similar material in the other system. For example (in both systems) pure aluminium rivets are black, Hidiminium rivets are violet, Monel rivets are natural and 5% magnesium rivets are green. This way of coding allows material types to be more easily identified.


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Table 13 MATERIAL IDENTIFICATION OF AS RIVETS Material Type Identification Marks Finish Dural ‘D’ on shank end Natural Al. Alloy ‘X’ on shank end Dyed or Anodised (5% Mg.) Green Hidiminium ‘S’ on shank end Dyed Violet Monel ‘M’ on shank end Natural or Cadmium Plated

Table 14 TYPICAL SPECIFICATION NUMBERS OF AS RIVETS Material Snap Mush 90º Csk 100º 120º 90º Close Spec. Csk Csk Tol. L37 AS156 AS158 AS161 AS164 AS2918 L58 AS157 AS159 AS162 AS4716 AS165 L86 AS2227 AS2228 AS229 AS2230 AS3362 DTD204 AS5462 AS465 -

Material. Spec. L36 L37 L58 L86 BS1109 DTD204

Table 15 MATERIAL IDENTIFICATION OF SP RIVETS Material Type Identification Marks Finish (On shank end) Aluminium ‘I’ Black Anodic Dural ‘7’ Natural Al. Alloy ‘8’ Green Anodic (5% Mg.) Hidiminium ‘0’ Violet Steel Cadmium Monel ‘M’ Natural or Cadmium

Table 16 TYPICAL SPECIFICATION NUMBERS OF SP RIVETS Material Spec. Snap Head Mushroom Head 100º Csk Head L36 SP77 SP68 L37 SP78 SP83 SP69 L58 SP79 SP84 SP70 L86 SP80 SP85 SP71 BS1109 SP76 SP86 DTD204 SP81 SP87 EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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13.2.4 SOLID RIVETS (AMERICAN)

These are generally used in normal construction and repair work. They are identified by the kind of material from which they are made, their head type, shank size and temper condition. Typical head types (refer to Fig. 72) are Roundhead, Brazier head, 100º Countersunk head, Flat head and Universal head.

AN Rivet Head Types

Plain

A

Dimple

AD

Raised Dot

D

2 Raised Dashes

DD

Cross

B

AN Material Identification and Code Letters Fig. 72


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13.2.5 RIVET IDENTIFICATION (AMERICAN)

The rivets, shown in Fig. 72, are of the AN (Air Force-Navy) designation and are merely used to illustrate a typical coding system. The other most common standard for American rivets is the MS (Military Standards) system which, whilst having slight differences from the AN system, uses similar terminology to describe the many forms of rivets. A part number (using the standard letters AN or MS) identifies each type of rivet, so that the user can select the correct rivet for the task. After the standard letters, there follows a number, which indicates the particular type of rivet head, Next comes a letter (or letters), denoting the material composition, which is followed by another figure expressing the diameter of the rivet shank in 32nds of an inch. The last number(s), separated by a dash from the diameter number, express the length of the rivet shank in 16ths of an inch. Example: An American AN system rivet with the identifying code AN470 AD 3-5, would be a Universal head, aluminium alloy (2117-T) rivet, of 3/32 inch diameter with a shank length of 5/16 inch. Note: With countersunk rivets, the length is the overall length. Head markings, using dimples and raised dots (or dashes and rings) are also used as an aid to indicate the material content of the rivets. Protective surface coatings, used by the manufacturers, are shown by colours, where zinc chromate is usually yellow, an anodised rivet is usually pearl grey and a metal sprayed rivet has a silvery grey colour. 13.2.6 RIVET MATERIAL IDENTIFICATION (AMERICAN)

As previously stated, the material used for the majority of aircraft solid rivets is aluminium alloy. Digits and letters identify the degree of temper condition, of aluminium alloy rivets, in a similar manner to that used for sheet aluminium alloy. The normal material grades are 1100, 2017-T, 2024-T, 2117-T and 5056. The 1100 (A) rivet is 99.45% pure aluminium and, as such, is very soft. It would be used for riveting lightweight, soft, aluminium structures, where strength is not a factor. The 2117-T (AD) rivet is made from aluminium alloy and (as has previously been mentioned) is known as the ‘field’ rivet. It is the most commonly used rivet, mainly because it is ready to use as received and needs no further heat-treatment. It also has a high resistance to corrosion.


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The 2017-T (D) and 2024-T (DD) rivets are made from high strength heattreatable aluminium alloys. They are used where more strength is required than that obtained from the ‘field’ rivet. The rivets need to be heat-treated and, if not required immediately, they should be refrigerated until needed. The 2017-T rivet should be driven within 1 hour of removal from refrigeration (or following heattreatment) and the 2024-T must be driven within 10-20 minutes. The 5056 (B) rivet is used for riveting magnesium alloy structures, because of its galvanic compatibility with magnesium (to reduce the risk of corrosion). Mild Steel rivets are used for riveting steel parts while Corrosion Resistant Steel rivets are used for riveting CRS components in fire-walls and exhaust areas etc. Note: The absence of a letter following the AN standard number indicates a rivet manufactured from mild steel. Monel (M) rivets are used for riveting nickel-steel alloys. They may also be used as a substitute for CRS rivets when specified. Copper (C) rivets are also available, but their use is limited on aircraft. They may only be used on copper alloys or non-metallic materials, such as leather. Note: Most metals, including aircraft rivets, are subject to corrosion. This may be the result of local climatic conditions or the fabrication process used. It can be reduced to a minimum by using the correct materials and by the use of protective coatings on the structure and the rivets. The use of dissimilar metals should be avoided where possible and, as previously stated, the rivet manufacturers usually apply a protective coating on the rivets, which may be either of a zinc chromate, a metal spray or an anodic film finish. 13.3 HEAT-TREATMENT/REFRIGERATION OF SOLID RIVETS The action of closing a rivet, and the strength required on completion, dictates whether any heat-treatment will be required prior to closing. As previously discussed, some rivets, for non-structural applications, can be manufactured from pure aluminium. These are given no heat-treatment and are soft, both before and after closing. Among the most common rivets in use (and which are made of aluminium alloy) are those already identified, in the American AN specification system, as ‘AD’ rivets. AD rivets are heat-treated during manufacture and remain easy to close whilst possessing adequate strength. Where rivets of a stronger material are required, then ‘D’ and ‘DD’ rivets can be used. These are also made from aluminium alloys, but to different (AN) specifications. They are heat-treated, just prior to use, and either formed within a short time period of time (in which they ‘age-harden’), or they are stored, in a EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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refrigerator, at temperatures well below zero degrees Celsius (which retards the age-hardening process), until required for use. They are known as ‘icebox’ rivets in the USA. 13.3.1 HEAT-TREATMENT.

Metal temper is important in the riveting process, especially with aluminium alloy rivets. These generally have the same heat-treating characteristics as sheet alloys and can be annealed and hardened in much the same manner. The rivet must be soft, or comparatively soft before a good head can be formed. The 2017-T and 2024-T rivets must be solution-treated before being driven and then they harden with age. The process of heat-treatment of rivets (normalising) may be achieved in either an electric, air furnace or in a salt bath. The temperature range, depending on the alloy, is in the region of 495ºC - 505ºC. For convenient handling, the rivets are heated on a tray or in a wire basket and, after heating for the required period, they are finally quenched in cold water 13.3.2 REFRIGERATION.

The heat-treated rivet will begin to age harden immediately after treatment and, if the rivets are not to be set immediately, they may be refrigerated to delay the age-hardening process. The solution-treated rivets are stored at low temperature (below freezing) and, under these conditions, will remain soft enough for driving for up to 2 weeks. Any rivets not used in that period should be removed and reheat treated. It should be noted that refrigeration only delays age-hardening and that agehardening will continue at a rapid rate as soon as the rivets are removed from the refrigerator. 2017-T rivets must be driven within 1 hour of refrigeration and 2024-T rivets, within 10 minutes 13.3.3 USE OF DIFFERENT TYPES OF RIVET HEAD

The many forms of rivet heads have evolved due to the specific requirements of an application and, whether they are of the British or American (or any other) standards, their designs and uses are fairly similar. A selection, considered here, gives typical used for the more common types of rivets: • Brazier head: has a head of larger diameter, making it suitable for riveting thin sheet. It offers only a slight resistance to airflow and is often used on exterior skins, especially on aft sections of fuselage and empennage. A modified brazier head rivet is also produced which has a reduced head diameter.


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• Countersunk head: this rivet is flat topped and bevelled towards the shank so that it can be installed into a countersunk or dimpled hole and so be flush with the material’s surface. The countersunk angle may vary from 78º to 120º (the 100º rivet being the most common type). Countersunk rivets are used to fasten metal sheets which overlap others. They are also used on exterior surfaces of the aircraft, because they offer only a slight resistance to airflow and therefore minimise turbulence. • Flathead: used on interior structures, where there is insufficient clearance to use a roundhead rivet. • Roundhead: used in the interior of the aircraft and has a deep rounded top section. The head is large enough to strengthen the sheet around the hole and to offer resistance to tension. • Universal head: this rivet is a combination of brazier, flathead and roundhead. It is used in aircraft construction and repair in both interior and exterior locations. It may be used as a replacement for all protruding head types of rivet. 13.4 BLIND AND HOLLOW RIVETS There are many places in an aircraft where access to both sides of the structure is impossible, or where limited space will not permit the use of a reaction (bucking) bar. Also, in the attachment of many non-structural parts, such as aircraft interior furnishings, flooring material, de-icer boots etc, the full strength of solid shank rivets may not be necessary. For use in such places, special rivets have been designed which can be set from one side only. These rivets are often lighter than solid rivets, yet amply strong enough for their intended use. The rivets are produced by several manufacturers, and have unique characteristics requiring special installation tools and procedures. The same, general, basic information, relating to their fabrication, composition, uses, selection, installation, inspection and removal procedures applies to most of them. Hollow rivets that can be closed by pulling a mandrel through them are often known as ‘blind’ rivets and these in turn can be described as Mechanically Expanded Rivets. They can fall into one of three main types: •

Self-Plugging (friction lock) rivets

•

Self-Plugging (mechanical lock) rivets

•

Pull-Through rivets

Where blind or hollow rivets are installed in place of solid rivets, (due, perhaps, to the lack of access to the both sides of the joint), they must, in the absence of specific instructions, be of the same material as the original solid rivet, and be of equivalent shear strength. The shear strength, of the rivet, may be increased, by using a form of ‘plug’ to fill the hollow shank of the rivet. EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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13.4.1 FRICTION LOCK RIVETS

These are generally fabricated in two parts, consisting of a rivet head with a hollow shank and a stem that extends through the hollow shank. They may, typically, be of the ‘friction lock’ protruding head or countersunk head styles of rivet (refer to Fig. 73). Several events occur in sequence when a pulling force is applied to the stem of the rivet: •

The stem is pulled into the rivet shank

•

The mandrel part of the stem forces the rivet shank to expand

• When friction (pulling action) becomes great enough, it caused the stem to fracture at the weakest point. The bottom end of the stem is retained in the shank, giving much greater shear strength than could be obtained from a hollow rivet. Note: With this type of rivet, the stem is often designed to break above the rivet head, necessitating a further action, which entails cutting off the extra portion of the stem with snips (or a specialised pneumatic gun) and milling the exposed portion flush with the head. This type of rivet is going out of style because of the extra work involved with setting it.

Friction Lock Rivets Fig. 73


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13.4.2 MECHANICAL LOCK RIVETS

A mechanical lock-type of rivet (refer to Fig. 74), is similar in design to the friction lock rivet previously described, except in the manner in which the mandrel is retained in the rivet. This type of rivet has a positive mechanical locking collar, to resist the vibrations that may cause the friction lock rivet mandrels to loosen and fall out. In addition, the mechanical locking-type rivet stem breaks off flush with the head and, usually, does not require further stem trimming when properly installed. Self-plugging, mechanical lock rivets display all the strength of solid rivets and, in most cases, can be substituted rivet for rivet. Three operations are performed when the rivet is installed (generally using a pneumatic gun): • When pulling force is exerted on the stem, the stem is pulled in, forming the blind head and clamping the sheets of metal together. • At a pre-determined point, the inner anvil, incorporated in the gun, forces the locking collar into position. •

The rivet stem snaps off approximately flush with the head of the rivet.

Mechanical Lock Rivets Fig. 74


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13.4.3 HOLLOW/PULL-THROUGH RIVETS

When installed, the rivet mandrel is pulled through these rivets, leaving a hollow rivet of much lower strength than the self-plugging types. Different types of these rivets are supplied, either complete with individual mandrels or as individual rivets, used with a re-usable steel mandrel, which is drawn completely through the rivets. In some cases, the rivets may be plugged with sealing pins which, as previously stated, give them additional strength as well as sealing them. 13.4.4 GRIP RANGE

Unlike a solid rivet, the part of a blind rivet, available to form a head, cannot always be seen. It is, therefore, necessary to know the range of total material thickness that a given rivet can fasten together. This is known as the ‘Grip Range’ of the rivet and requires the use of a gauge to measure the material thickness (refer to Fig. 75), which is used in conjunction with a rivet data table.

1

3

5

7

3

Rivet Group to be Used = 4 Grip Measuring Gauge Fig. 75


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13.4.5 TUCKER ‘POP’ RIVETS

Tucker ‘Pop’ rivets (refer to Fig. 76) are supplied mounted on steel mandrels. The head of the mandrel is pulled into the rivet, expanding it, before the mandrel fractures at the waisted portion. This waisted portion may either be near to the head of the rivet, or part way up the stem. In the first case the rivet will be classified as ‘Break Head’ (BH) and in the second case, ‘Break Stem’ (BS) The rivets are set, using a pair of ‘Pop’ pliers or by the use of a hydro-pneumatic gun. ‘Pop’ rivets are less suitable for use on aircraft as they tend to loosen with vibration and then become increasingly difficult to remove, because of the looseness and the presence of the steel mandrel. (They also tend to spin when attempts are made to drill them out).

‘Pop’ Rivets Fig. 76


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Break head rivets must not be used if the structure is not accessible to retrieve the mandrel heads. It is sometimes permitted for the mandrels of Break Stem rivets to be dipped in an adhesive, so that they will not vibrate loose after installation. If Tucker ‘Pop’ rivets are to be used externally on aircraft, the heads must be sealed to prevent the ingress of dirt and moisture. Cellulose Metallic Filler is often recommended for this purpose. The rivets are manufactured in either aluminium alloy or cadmium-plated Monel metal, with either dome heads or 100º and 120º countersunk heads. The AGS reference number consists of the AGS number identifying the material and head type, a three figure size code and letters specifying Break Head or Break Stem. In the size code the first figure represents the diameter, in increments of 1/32 inch while the last two figures indicate the length in increments of 1/10 inch. Example: A rivet, with the designation code AGS2051/537/BS, would be a Tucker ‘Pop’, made from Monel metal, with a 120º Csk. Head. The figure 537 indicates that its diameter is 5/32 inch and its length is 0.37 inch. BS shows that it is a Break Stem. Note: Care must be taken to ensure all remaining stems and swarf, are totally removed from the aircraft, on completion of work, when using these rivets. 13.4.6 AVDEL RIVETS

Avdel rivets (refer to Fig. 77) are rarely used today, but may be found on older aircraft. To close the rivet, the stem is pulled through and, at a predetermined load, the stem breaks proud of the manufactured head of the rivet, plugging the rivet body. Whilst the stems can be milled off on alloy rivets, those manufactured of stainless steel or titanium break flush with the rivet head. A flush finish is required for aerodynamic reasons. Avdel rivets are pre-lubricated by the manufacturer, to facilitate forming the rivet. They should NEVER be de-greased in solvent before use.


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Avdel Rivets Fig. 77 13.4.7 CHOBERT RIVETS

Chobert rivets (refer to Fig. 78) are also similar to Tucker ‘Pop’ rivets, but have a tapered bore. The head of the mandrel is re-usable, and is pulled fully through the rivet on forming. This gives an advantage of no loose articles after the riveting operation is completed. The mandrel is drawn through the rivet using a special tool, which carries a number of rivets on the mandrel to allow repetitive and faster riveting. The tool simply feeds the next rivet into place after the closure of the previous one. Where additional shear strength or water-tightness is required, sealing pins or plugs of the same material are driven into the bore of the closed rivets

Chobert Rivets Fig. 78


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13.4.8 CHERRY RIVETS

Cherry Rivets (refer to Fig. 79) consist of a range of fasteners including CherryLok and Cherry-Max, which are manufactured in the USA. The primary difference between these and the rivets mentioned previously is that the mandrel is locked in position, after closing, instead of depending on friction alone. During the final stages of closing, a locking collar, located in a recess in the rivet head, is forced into a groove in the stem and prevents the stem from any further movement. This method means that, when closed, the rivets have a shear and bearing strength high enough to allow their use in place of solid-shank rivets.

Cherry Rivets Fig. 79

13.5 MISCELLANEOUS FASTENERS These fasteners are, basically, close-tolerance, metal pins that combine the best features of a rivet and bolt. They usually require access to both sides of the joint but are extremely strong in shear, with a shear strength equal to a standard AN bolt of the same size. Three typical types, considered here, are: •

Hi-Lok Fasteners

•

Hi-Tigue Fasteners

•

Hi-Shear Fasteners

13.5.1 HI-LOK FASTENERS

The Hi-Lok fastener (refer to Fig. 80) consists of a metal pin (made from heattreated steel) which has a thin, manufactured head at one end and a partthreaded shank at the other. The threaded end of the Hi-Lok fastener contains a hexagon-shaped recess, for the insertion of an Allen Key.


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After the pin is located in its prepared hole, a hexagon-headed collar is turned onto the threaded shank by a box wrench or an ordinary spanner. An Allen Key engages in the recess in the shank end, to prevent rotation of the pin whilst the collar is being tightened and, when a pre-determined load is reached, the hexagonal section of the collar shears off, leaving the pin securely fastened in the hole. Because the collar breaks off at a designated pre-load, the use of torque wrenches is eliminated and three primary design advantages are: • Accurate pre-load and torque to within 10%. • Minimum size and weight. • Rapid, quiet, single-handed operation.

Remaining Portion of Collar After Assembly Collar Wrenching Device Shears Off

Typical Collar

Collar Driving-Hex Pin Recess-Hex Hi-Lok Fastener Fig. 80

13.5.2 HI-TIGUE FASTENERS

Hi-Tigue fasteners (refer to Fig. 81) are similar to Hi-Loks, excepting that they possess a bead at the bottom of the shank, adjacent to the threaded portion of the fastener. The bead exerts a radial load to the side of the hole which serves to strengthen the area surrounding the fastener hole. This reduces the effect of cyclic loads on the fastener which, in turn, will reduce the effect of the cold working of the joint and minimise the likelihood of subsequent failure. Hi-Tigue fasteners are closed in exactly the same manner as the Hi-Lok types.


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Bead

Hi-Tigue Fastener Fig. 81

13.5.3 HI-SHEAR FASTENERS

A Hi-Shear fastener (refer to Fig. 82) is a close-tolerance pin, which is an interference fit and must be tapped into its hole before the locking collar is swaged on. There are two head styles; one being flat while the other is countersunk. The rivets are closed, either with a special pneumatic pulling tool or by a conventional riveting gun and a special, conical, gun-set.

Hi-Shear Fastener Fig. 82


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13.6 SPECIAL PURPOSE FASTENERS In addition to the fasteners already described, other rivet-type fasteners are often used in the manufacture and repair of aircraft. While some of these are designed for a specific use, others may be categorised as ‘High Strength Fasteners’. Typical examples of these special purpose-type fasteners include Jo-Bolts, Tubular Rivets and Rivnuts. 13.6.1 JO-BOLTS

This is the trade name for a fastener, which is used where a nut and bolt would normally be fitted, but where access is available only to one side of the work. Jo-bolts (refer to Fig. 83) consist of three components; an alloy steel nut (which may be of a hexagonal or countersunk headed type), a hollow steel bolt and a stainless steel sleeve. The fastener is installed with either a pneumatic or a hand-operated tool, with which the bolt is rotated and the nut is held stationary. This action expands the sleeve over the tapered end of the nut and draws the fastened items together. At a pre-determined torque, the bolt breaks off at a notch-weakened point, flush with the head of the nut. A different tool is required for each of the two head forms and for each particular diameter bolt.

Jo-Bolts and Installation Sequence Fig. 83


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uk engineering 13.6.2 TUBULAR RIVETS.

Tubular rivets are used primarily to save weight when riveting through tubular or hollow members, where a large part of the rivet is merely passing through space. They are often used on control rods for connecting end fittings. The rivets are made to AGS drawing specifications in several materials. The drawing number indicates the type of rivet and the following letter denotes the material. The number after the letter denotes the dimensions of the rivet, but has no particular significance as is the case with other types of rivet. Example: A tubular rivet with the designation code AGS 501/H/49 is made of mild steel, has a length of 1 inch, and has a wall thickness of 26 SWG. Table 17 shows the letters used to indicate different tubular rivet materials and the features by which the materials may be recognised. Table 17 IDENTIFICATION CODES FOR TUBULAR RIVETS Letter Material Identification Feature Identification Protective Physical Treatment Characteristic A

Aluminium (L54)

Anodic film

Dyed black

D

Duralumin (L37)

None

Natural colour

H

Mild steel (T26)

Cadmium plated

Magnetic

J

Nickel alloy (DTD268) or Monel metal (DTD204A)

Cadmium plated

Only slightly magnetic

K

Monel metal (DTD204A)

None

Only slightly magnetic


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13.6.3 RIVNUTS

These fasteners were produced to attach rubber de-icing boots to aircraft wing and tail leading edges. Rivnuts can be either of the countersunk or flat head types, of which, each can have open or sealed ends (refer to Fig. 84).

Rivnuts Fig. 84 Installation is achieved by drilling a hole into the skin and a small notch made on the edge of the hole to prevent the Rivnut rotating during closing The nut on the thread of the ‘puller’ is inserted into the hole (refer to Fig. 85), and the key aligned with the notch. The puller handle is squeezed, closing the nut and gripping the skin. The tool is then unscrewed from the Rivnut, leaving a threaded hole that accepts standard machine screws, for attaching the de-icer boots Rivnuts are supplied in American thread sizes and in BA or BSF thread forms, but to avoid confusion, only the American types are considered here. These Rivnuts are available in six grip ranges, the minimum grip Rivnut having a plain head while the next size has a radial dash mark on the head. Each succeeding grip range is indicated by an additional radial mark on the head with the largest size having five radial dash marks.


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Keyway

Rivnut Installation Fig. 85

.


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14 SPRINGS The invention of the wheel, for transport (and many other applications), is considered to be one of the major advances of mankind, but another, lesspraised, technical, innovation followed the development of devices employed to alleviate the discomfort of travelling on unmade or rutted roads. Using the fact that the elasticity, inherent in most materials, allows them to absorb energy by distorting or deflecting when under load - and then, to return to their original shape after the load has moderated (or has been removed), - early springs consisted of flat (and curved) sections of wood (and, later, metal), to which were attached the carriages of the respective eras. The dawning of The Industrial Revolution led to the mechanisation of practically every facet of civilised life, from the production of food and textiles to the mining and processing of minerals in order to provide many other materials and the various machines deemed necessary for sophisticated living conditions. In addition there has followed huge advances in transport, time-keeping, worldwide communication and (inevitably) military capabilities, in all of which can be found mechanisms involving the principle of the spring. 14.1 FORCES EXERTED ON, AND APPLIED BY, SPRINGS The three basic forces, which may be exerted on, and applied by, springs are: •

Compressional

•

Tensile

•

Torsional

Note: These forces may act singly or in combinations of any two or all three. 14.2 TYPES OF SPRINGS Springs have evolved into various shapes and sizes (and degrees of stiffness), which have been dictated by the uses to which they have been put. The more common forms are included here for consideration. 14.2.1 FLAT SPRINGS

Flat springs, while they were a development of flat, rectangular-sectioned, strips of metal, can actually be found in forms other than simply flat as, for instance, in the shape of the springs which control the contact breaker points in the magneto of an aircraft piston-type engine.


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14.2.2 LEAF SPRINGS

Leaf springs are formed by layers of flat springs and while very early aircraft embodied leaf springs in their landing gear, this type of spring is more familiar in the automobile and railway industries. 14.2.3 SPIRAL SPRINGS

Spiral springs may be found in the form of spirally wound flat springs (known as motor or power springs) or as spirally wound wire, such as the hair springs of many types of instruments. 14.2.4 HELICAL COMPRESSION AND TENSION SPRINGS

These are the most commonly found springs, which superseded the leaf spring when space and lightness of structure were the requirement. They are made in a wide variety of materials and sizes and may be found in a seemingly endless number of applications. 14.2.5 HELICAL TORSION SPRINGS

While being similarly wound to the previous two types, these springs have specially shaped ends, to permit a torque force to be applied, and transmitted, in a plane normal to the helix axis. 14.2.6 BELLEVILLE (CONED DISC) SPRINGS

Belleville springs are, in fact, shaped like the Cup Washers, which were previously discussed in the topic on Locking Devices. Belleville Springs are capable of exerting frictional or linear forces. 14.2.7 TORSION-BAR SPRINGS

Torsion-bar springs are, basically, straight bars of metal, with splined (or flanged) ends, that can accept and transmit torsional forces. 14.3 MATERIALS FROM WHICH SPRINGS ARE MANUFACTURED The materials, used for the manufacture of springs, cover a very wide range of metallic and non-metallic (plastic and elastomer) substances. These notes will, however, be confined mainly to the discussion of metallic types, with a small consideration being given to some composite materials. 14.3.1 STEELS USED FOR COLD-WOUND SPRINGS

Below a material (stock) cross-sectional diameter of approximately 9.53 mm to 18.42 mm (0.375 in to 0.725 in) certain steels are drawn into wires and coldwound to form the required shape of the spring. The wires are, then, usually, given some form of heat-treatment, to relieve the stresses imposed by the winding processes. Typical types of carbon- and alloy-steel stock, used for the manufacture of springs, include:


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• Hard-drawn Spring Wire: which is of a low-quality (and least expensive) carbon-steel stock. This wire is liable to posses fine seams in its surface and, as such, would only be used in applications of low stress and not where fatigue loadings could be exerted • Oil-tempered Spring Wire: which is of a better quality, high-carbon steel, stock, though it may also contain surface discontinuities and would be found where long fatigue life is not required • Music Wire: which is a carbon-steel stock of high quality and is suitable for small-sized, helical springs in applications involving high fatigue stresses • Chrome-Vanadium Steel Wire: which is a stock that has been used for pistontype aero-engine valve springs and is, therefore, suitable for high-temperature and high-stress conditions • Chrome-Silicon Steel Wire: which, when used in valve springs, has a higher fatigue life in the lower cycle ranges (10 kHz – 100 kHz) than other wires • Stainless-Steel Spring Wire: which, as is obvious from its name, is used in conditions where high corrosion-resistance is the requirement. This grade of wire would also be utilised in applications where resistance to creep at elevated temperatures is desired. Some grades of Stainless-Steel wires can be made to accept magnetism, where this characteristic is needed alongside the other qualities. 14.3.2 STEELS USED FOR HOT-WOUND SPRINGS

Above the cross-sectional diameters, previously mentioned, it is considered impractical to cold-wind and so, the larger diameter metals (bars or rods) are hotwound and then also subjected to various stress-relieving processes. Similar carbon- and alloy-steels to those already discussed are employed in the manufacture of hot-wound springs, with the necessary variations in their contents of carbon, chromium, manganese, molybdenum, nickel, silicon, and vanadium. 14.3.3 STEELS USED FOR COLD-ROLLED, FLAT SPRINGS

These steels vary in composition, depending on their location, but are, commonly, based on carbon and manganese as the main constituent elements and may be formed from oil-tempered steels (thin sections – clock-type springs) or from annealed steels which are subsequently heat-treated. 14.3.4 NON-FERROUS METALS USED FOR SPRINGS

Based mainly on copper alloys, where corrosion resistance and good electrical conductivity is required, and on nickel alloys where the ability to work at elevated temperatures is desirable, these alloys include: • Spring Brass: which is comparatively inexpensive, has good electrical conductivity, but is unsuitable for high-stress applications


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• Nickel Silver (also called German Silver): which has better characteristics than brass and is, in fact, made from different percentages of copper, zinc and nickel • Phosphor Bronze: which has a minimum percentage of 90% copper content and has, therefore, excellent electrical conductivity. It is suitable for applications of higher stress levels than those of brass • Silicon Bronze: which has similar characteristics to those of phosphor bronze but is less expensive to produce • Beryllium Copper: which has similar conductivity (and corrosion resistance) qualities to those of copper with the addition of beryllium (2.0& – 2.5%) imparting greater hardness and other superior mechanical properties • High-Nickel Alloys: which are the types more commonly found in aero-engine applications and which fall under various, familiar, trade names such as: • Monel • ‘K’ Monel (3% aluminium) • Permanickel • Inconel • Inconel ‘X’ (2.5% Titanium) Note: Another high-nickel alloy goes under the name of Ni-Span-C and does, in fact, contain almost 50% iron. All of these non-ferrous alloys can be found in the cold-rolled or drawn conditions for the manufacture of many types of springs. 14.3.5 COMPOSITE MATERIALS USED FOR SPRINGS

Some composite springs involve the joining of certain metals with elastomers to form the anti-vibration mountings (Metalastic Bushes and Housings) such as those found in aero-engine and auxiliary power unit (APU) installations Others combine synthetic rubber strands, encased within a sheath of braided cotton, nylon or similar materials. They are, usually, referred to as ‘Shock Absorbers’ or ‘Shock Cords’ rather than ‘Springs’ and are more familiarly known by the generic name of ‘Bungee Cords’. Bungee Cords may be encountered on many light- and medium-sized aircraft while their use on heavier aircraft is not unknown.


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14.4 CHARACTERISTICS OF TYPICAL AEROSPACE SPRINGS If a Load/Deflection graph, for a typical, helical-wound, spring, were to be plotted, it would be found (provided the spring was not loaded beyond the elastic limit of the material and ignoring the effects of temperature and constant or repeated loadings), that a straight-line graph would result (refer to Fig. 86 (a)). This indicates that the deflection is directly proportional to the load so that, if the load is doubled, then the deflection also doubles – a characteristic of which good use is made in so many aeronautical applications. Belleville springs, however, present a different form of graph (refer to Fig. 86 (b)) and, yet again, their particular characteristics also prove extremely useful in certain control and indicating functions of aircraft structures and components.

(a) Helical Springs

(b) Belleville Springs

Load/Deflection Graphs for Springs Fig. 86


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14.5 APPLICATIONS OF SPRINGS IN AIRCRAFT ENGINEERING If an examination were made, of virtually every Subject Topic of every Section of every Chapter of the many Aircraft Maintenance Manuals, complying with the ATA Specification No. 100 (from Air Conditioning through to Accessory Gearboxes), then numberless examples would be found, of the applications involving the use of springs in aircraft engineering Many applications have already been mentioned but some further examples, of the uses for springs, could include their use as: • Pressure Regulating/Limiting Devices: in Fuel, Hydraulic, Lubrication, and Pneumatic systems • ‘Fail Safe’ or ‘Return to Neutral Condition’ Devices: in Electrical Relays and Solenoids, and also in Electric, Hydraulic, Mechanical, or Pneumatic Actuators • Acceleration and Speed Control Devices: in Engine and Propeller control systems and in Power-Assisted Flight Controls and Wheel Braking systems • Shock Absorbing Devices: in Landing Gear systems and as Anti-Vibration Mountings for delicate instruments and components which are subject to movement • Devices which are capable of applying a constant force (linear or rotary) in a desired direction, as in the holding closed of an aero-engine valve spring for one example • Devices with the ability to accurately indicate (and control) the value of an applied force, as used in many instruments (Ammeters, Voltmeters, Fuel Flow Meters and Tachometers provide typical examples) Note: The subject of spring technology is vast and well beyond the scope of these notes, so it is sufficient for the student to appreciate the basic uses for springs in the aerospace environment and the functions that they fulfil.


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15 PIPES AND UNIONS The many different systems in an aircraft require the services of pipes and hoses, in a range of sizes. These can include fuel, oxygen, lubrication, hydraulic, instruments, heating, fire extinguishing, air conditioning and water systems. Loss of integrity in any of these systems could put the aircraft at risk. The pressures inside the pipes can vary from negative (suction) through ambient, in instrument piping, to as much as 4000 psi (27.58 x 10³ kN/m²) in a hydraulic system. Low-pressure fluid lines can be manufactured from metal or plastic (pipes and tubes) or, alternatively, from various forms of rubber (hoses). Highpressure fluid line can be made from a variety of materials, including aluminium alloy, stainless steel, copper, titanium and also reinforced flexible hoses. Fluid lines are made of rigid, semi-rigid and flexible tubes, depending on their use. A rigid fluid line would be one that is not normally bent to shape or flared. Direction changes and connections are made by the use of threaded end-fittings. Semi-rigid fluid lines are bent and formed to shape and have a relatively thin wall thickness in comparison to rigid lines. A variety of end-fittings may be used to make connections between semi-rigid tubes. Flexible fluid lines are made from rubber or synthetic materials and are usually called ‘hoses’. Depending on the pressure they are designed to carry, hoses may have reinforcing materials wrapped around them. Various types of end-fittings are used to attach hoses to each other and to other components. 15.1 RIGID PIPES These are usually manufactured in a standardised combination of length, outside diameter (OD), and wall thickness. The use of threads, cut into the pipe wall, and the need for special end-fittings means that, apart from some components, there are few, if any, rigid pipes used on aircraft.


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15.2 SEMI-RIGID FLUID LINES (TUBES) Semi-rigid fluid lines are usually referred to as tubes or tubing and can be bent to shape and are often flared for connectors. Sizing is also by length, OD and wall thickness. Various methods are used to connect semi-rigid tubes both to each other and to other connectors. These will depend upon the use, location and pressure being carried in the tube. The most common end-fittings are of the flared, flare-less, swaged or brazed types and are, often, standard parts. 15.2.1 FLARED END-FITTINGS

American flared end-fittings have a 74° flare (remember AGS are different and not compatible) on the end of the tube, which matches a cone of the same angle on the component (or adapter) to which it is being attached. A special nut and sleeve are used to pull the flare onto the cone and to form a fluid-tight metal-to-metal seal. The end-fittings are produced in a wide variety of types, depending upon their use. Examples are the ‘In-line-’, ‘Cross-’, ‘Elbow-’, and ‘T’-type of end-fittings, in addition to ‘Bulkhead’ fittings, which allow tubes to pass fluids through structural portions (bulkheads) of an aircraft or of an engine power-plant assembly. In-line connectors may be either of the pipe-to-pipe or pipe-to-adapter type of connectors and internally coned adapters usually require the use of adapter ‘nipples’ to provide an effective seal (refer to Fig. 87).

Pipe-to-Pipe and Pipe-to-Adapter Connectors Fig. 87 Where it is necessary to have fuel, oil or other tubes passing through structural bulkheads, it requires an end-fitting with a long body and provision for securing EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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the end-fitting to the bulkhead (refer to Fig. 88). Two typical bulkhead end-fittings, AN832 and AN833, are among those illustrated and they can be identified by the extra machine thread at one end, for attachment, to the bulkhead, by an additional, threaded, locking device.

Typical American (including ‘Bulkhead’) Connectors Fig. 88

15.2.2 FLARE-LESS COUPLINGS

The heavy-wall tubing, used in some high-pressure systems, is difficult to flare (and flaring tends to put the end of the tube in a stressed condition). For these applications the flare-less coupling is designed to provide leak-free attachments without flares. Although there is no need to flare the tube, in one of the methods used, it is necessary to pre-set the coupling, prior to its installation (refer to Fig. 89). Presetting is the process of applying enough pressure to a sleeve (also called a ferrule) to cause it to cut into the outside of the tube. The tube and ferrule are placed into a pre-setting tool and the action of tightening EASA Part 66 - Materials and Hardware.doc Issue 1 - 21 March, 2013

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the nut causes the ferrule to ‘bite’ into the tube. Depending on the size of the tube and its material, between one and one and a half turns of the nut is enough to form the pre-set. When complete, the tube can be inspected and, if satisfactory, attached directly to the appropriate union or adapter.

Pre-set Flare-Less Coupling Fig. 89 Two other methods of forming flare-less couplings involve the swaging of metal sleeves around the ends of the tubes, which are being connected and the joining of tubes by brazing. Both methods require specialist skills, which are beyond the scope of these notes. 15.3 FLEXIBLE PIPES (HOSES) The need for flexibility in many areas of aircraft construction means it is often necessary to employ hoses, instead of semi-rigid tubing, for the transmission of fluids and gases under pressure. Whilst a number of hoses were previously manufactured from rubber, most modern hose manufacturers use either Teflon or other elastomers.


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15.3.1 LOW-PRESSURE HOSES

An example of the type of construction used in these hoses is where the inner and outer tubes are made from synthetic rubber, with the inner having a braided cotton reinforcement (refer to Fig. 90). These hoses are used on instrument systems, vacuum systems, autopilots and other low-pressure systems, usually operating at pressures below 300 psi (2.07 x 10³ kN/m²).

Low-Pressure Hose Fig. 90 A typical marking on this type of hose could be a yellow line with the letters ‘LP’ along it. The line (lay line) is used to ensure that the hose is not assembled with a stress-inducing twist in it. Other markings could include the hose manufacturer’s code and part number, its size and the date of manufacture 15.3.2 MEDIUM-PRESSURE HOSES

Medium-pressure hoses are generally used with fluid pressures up to 1500 psi (10.34 x10³ kN/m²). Their maximum pressure varies with diameter, so that whilst smaller diameter hoses will be able to withstand such pressures, larger sizes may be restricted to lower pressures. Typical construction of this type of hose could be a seamless inner liner made from different materials, a layer of cotton braid, a layer of stainless-steel reinforcement and an outer layer of tough, oil-resistant, rubber-impregnated cotton.


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15.3.3 HIGH-PRESSURE HOSES

All high-pressure hoses (refer to Fig. 91) have a maximum working pressure of at least 1500 psi to 3000 psi (10.34 x 10³ kN/m² to 20.68 x 10³ kN/m²) and use a synthetic rubber liner to carry petroleum products. The inner liner is usually wrapped with two or more steel braids as reinforcement. To distinguish highpressure from medium-pressure hose, the entire hose usually has a smooth outer cover

High-Pressure Hose Assembly Fig. 91

The end fittings on a flexible hose assembly are made of steel or light alloy, depending on their application. They are designed to exert a grip on the tubes and wire braids, so as to resist the high pressure twisting and vibrating loads, as well as providing an electrical bond throughout the assembly. Flexible hoses have their sizes identified by their inner bore diameter and the overall length. With pre-assembled hoses, the overall length of the assembly, from the centres of the nipple extremities, regardless of the shape of the end fittings, is used for identification purposes (refer to Fig. 92). Flexible hoses, used in engine bays and other high temperature areas, will often have a metallic stainless braid as the outside layer, to make the hose fireresistant.


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Effective Length

Effective Length

Effective Length

Effective Length of Hose Assemblies Fig. 92 15.4 UNIONS AND CONNECTORS Very few pipes and hoses are manufactured at company engineering facilities, the majority being obtained direct from manufacturers and specialist suppliers. It is important that engineers be aware of the variety of different types of unions and connectors that are available for rigid pipes and flexible hoses on aircraft. These may be of British, European or American manufacture with the different standards that these entail.


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15.4.1 AIRCRAFT GENERAL STANDARDS (AGS)

It has already been discussed, in earlier topics, how this British standard comprises a wide range of small parts, which includes items such as bolts, nuts, rivets and taper pins. The standard also includes pipe end-fittings (union nuts and adapters), sleeves, collars, and nipples. The cones (flares) on AGS end-fittings (unions and adapters) have an included angle of 32º, with the pipe flaring machines being shaped accordingly. 15.4.2 AIR FORCE AND NAVY (AN)

This standard may also be found in a wide range of aircraft and components, but it should be noted that the flares and other hardware for this standard have an included angle of 74º. 15.4.3 MILITARY STANDARD (MS)

This standard (as previously discussed) has replaced the standards from the AN system. Many AN part numbers have been incorporated into the MS system and now appear with MS designations Other specifications in current use with aircraft manufactured in the USA include National Aerospace Standards (NAS) and Military Specifications (Mil Specs). These may have an equivalent civilian or Military Standard. The Society of Automotive Engineers (SAE), and the Aeronautical Materials Division of SAE specifications (AMS) are yet another set of standards to which aerospace materials may be produced. The Society of Automotive Engineers has a second standard - referred to as the Aeronautical Standard (AS) – which is for components that do not qualify for an AMS standard. All these specifications provide for a range of fasteners with Unified threads in the UNC, UNF and UNJF series and, whereas British aircraft fasteners are manufactured in a selected range of Unified threads, American fasteners are in some instances supplied in both UNC and UNF threads. From all this it can be seen that great care must be taken when matching up union assemblies with these many different forms of thread. 15.5 QUICK-RELEASE COUPLINGS Quick-release couplings are required at various points in aircraft systems. Typical uses are in fuel, oil, hydraulic and pneumatic systems. Their purpose is to save time in the removal and replacement of components; to prevent the loss of fluid and to protect the fluid from contamination. The use of these couplings also reduces the maintenance cost for the system involved.


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A coupling consists of a male and female assembly (refer to Fig. 93). Each assembly has a sealing piston (poppet valve) that prevents the loss of fluid when the coupling is disconnected. Three checks may be used to verify a positive connection. These involve an audible, visual and tactile indication. A click may be heard at the time the coupling is locked and indicator pins will extend from the outer sleeve upon locking, which can be seen and felt.

Typical Quick-Release Coupling Fig. 93


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16 BEARINGS Bearings are, broadly, classified by the type of rolling element used in their construction. Ball bearings employ steel balls, which rotate in grooved raceways, whilst Roller bearings utilise cylindrical, tapered and spherical rollers running in suitably shaped raceways (refer to Fig. 94). Although these notes give information on the uses of the various types of ball and roller bearings, - together with general information on installation, maintenance and inspection, - the Aircraft Maintenance Manual (AMM) should be the final arbiter for specific installations. Ball bearings and tapered roller bearings accept both radial and axial loads, whilst the other types of roller bearings may accept only radial loads. Those bearings, which are contained in cages, are, in general, used for engine and gearbox applications with rotational speeds in excess of approximately 100 rpm. Most other bearings, on an aircraft or in an engine, are intended for oscillating or slow rotation conditions and do not have a cage. They are generally shielded or sealed and pre-packed with grease, although some have external lubrication facilities.

Ball and Roller Bearings Fig. 94


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16.1 BALL BEARINGS Ball bearings may be divided into four main types that define the way in which the bearings are used. The main types of Ball bearings are: •

Radial Bearings

•

Angular-Contact Bearings

•

Thrust Bearings

•

Instrument Precision Bearings

16.1.1 RADIAL BEARINGS

Radial bearings are the most common type of bearing and can be found in all types of transmission assemblies such as shafts, gears, control rods and end fittings. They are manufactured with either a single or double row of balls, rigid for normal applications and self-aligning for positions where accurate alignment cannot be maintained, such as in control rod ends. 16.1.2 ANGULAR-CONTACT BEARINGS

Angular-Contact bearings are capable of accepting radial loads and axial loads in one direction only. The outer ring is recessed on one side to allow the ball and cage assembly to be installed, thus enabling more balls to be used and the cage to be in one piece. The axial load capacity depends on the contact angle. In applications where axial loads will always be in one direction, a single angularcontact bearing may be used but, where they vary in direction, an opposed pair of bearings may be used. 16.1.3 THRUST BEARINGS

Thrust bearings are designed for axial loading only. They will usually be found in use together with roller or radial ball bearings. The balls are retained in a cage and run on flat or grooved washers. These bearings are adversely affected by centrifugal force and so work best under high-load, low-speed situations. 16.1.4 INSTRUMENT PRECISION BEARINGS

Instrument Precision Bearings are manufactured to high accuracy and finish. They are generally of the radial bearing type and can be found in both instruments and communication equipment.


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16.2 ROLLER BEARINGS Roller bearings may be divided into three main types that define their use. They are: •

Cylindrical Roller Bearings

•

Spherical Roller Bearings

•

Tapered Roller Bearings

16.2.1 CYLINDRICAL ROLLER BEARINGS

Cylindrical Roller bearings will accept greater radial loads than ball bearings of the same size. This is due to the greater contact area of the rolling elements and, if they have ribs on both rings, cylindrical roller bearings will also accept light, intermittent, axial loads. Normally the rollers have a length equal to their diameter, although some rollers have a length greater than their diameter to cater for special applications. Roller bearings, which have a length much greater than their diameter, are normally called needle roller bearings. These are designed for radial loads only and are best used in situations where the movement is oscillatory rather than rotary, such as in universal joints and control rod ends. 16.2.2 SPHERICAL ROLLER BEARINGS

Spherical Roller bearings can be found with single or double rows of rollers, which run in a spherical raceway in the outer ring, thus enabling the bearing to accept a small degree of misalignment. These bearings will accept high radial loads and moderate axial loads. 16.2.3 TAPERED ROLLER BEARINGS

Tapered Roller bearings are designed so that the axes of the rollers form an angle to the shaft axis. They are capable of accepting radial and axial loads simultaneously, in one direction only. It is common to find tapered roller bearings mounted in pairs, - back to back - so that loads can be accepted in both directions.


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16.3 BEARING INTERNAL CLEARANCE Due to the heat, generated during operation, Radial Ball and Cylindrical Roller bearings are manufactured with different amounts of internal clearance. The bearings are produced in four grades (groups), and are usually marked in some way to indicate each particular group. A system of dots (or circles or letters) is often used as identification and it is most important that replacement bearings are to the same standard as those removed. 16.3.1 GROUP 2 (‘ONE DOT’) BEARINGS

Group 2 bearings have the smallest radial internal clearance and are, normally, used in precision work, where minimum axial and radial movement is required. These bearings should not be used in applications where high temperatures could reduce the internal clearance and are not suitable as thrust bearings nor for high-speed situations. 16.3.2 NORMAL GROUP (‘TWO DOT’) BEARINGS

Normal Group bearings are used for most general applications, where only one ring, of the bearing race, is an interference fit and where no appreciable amount of heat, is likely to be transferred to the bearing. 16.3.3 GROUP 3 (‘THREE DOT’) BEARINGS

Group 3 bearings have greater internal clearance than Normal Group bearings and are employed where both race rings are interference fits, or where one ring is an interference fit, and some transfer of heat must be accepted. These bearings are also used for high speed and in applications where axial loadings are predominant. 16.3.4 GROUP 4 (‘FOUR DOT’) BEARINGS

Group 4 bearings have the greatest internal clearance and are found where both rings are interference fits and where the transfer of heat reduces internal clearances. Standard bearings are produced in all four groups while instrument precision bearings are supplied only in the first three groups


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16.4 BEARING MAINTENANCE Ball and roller bearings, if properly lubricated and installed, have a long life and require little attention. Bearing failures may have serious results and, to avoid such problems, Aircraft Maintenance Manuals and approved Maintenance Schedules include full lubrication and inspection instructions, which MUST be followed in order to limit the likelihood of bearing failure. 16.4.1 LUBRICATION

As has already been stated, most bearings, used in airframe applications, are shielded (sealed) to prevent the entry of dirt or fluids, which could affect bearing life. These cannot, normally, be re-greased and must be replaced if there are signs of wear, loss of lubricant or brinelling. (brinelling is the indentation of the surfaces of the bearing races). In some places, where there is risk of loss of lubricant, a grease nipple will be provided to permit recharging with fresh grease. Greasing should only be done after the nipple has been wiped clean of all dirt and, on completion, all excess grease must be wiped away with a clean cloth. 16.4.2 INSPECTION

Bearings are designed to operate with little or no maintenance, but they must be inspected regularly because, if corrosion or wear begins, the bearing will deteriorate rapidly. Bearings are usually inspected without removing them from the component (in situ), as continued removal and installation of bearings can cause wear and damage. Wheel bearings are inspected when the wheel is returned to the Wheel Servicing Bay for maintenance. Other items might also be inspected when their major assembly is removed for ‘off-aircraft’ maintenance. ‘On-aircraft’ checks can include checking for smoothness of operation, for wear (by moving the assembly both axially and radially) and also for any signs of interference or fouling with (or from) adjacent components.


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17 TRANSMISSIONS In mechanical engineering terms, transmissions consist of a series of connected parts (or mechanisms) whereby a source of power can be applied to another component, which is, then, able do the required work in the form of motion. Transmissions can be used to: •

Connect two (or more) shafts so that one provides drive to another (or others)

•

Change the speed of one shaft relative to another

•

Change the direction of rotation of one shaft relative to another

•

Convert one type of motion to another (rotary to linear or vice versa)

Aerospace transmissions involve the use of a wide range of sources, to provide the power, which eventually results in the desired motion in a particular system and those power sources include (singly or in combinations): •

Muscle and (where possible), assisted muscle power

•

Hydraulic, pneumatic and electrical power

•

External and (most frequently) internal combustion engines

The means of transmitting power, from a source to provide eventual motion, is achieved by such devices as: •

Belts and Pulleys

•

Gears

•

Chains and Sprockets

17.1 BELTS AND PULLEYS Whilst some forms of pulley are covered in the section on controls, there are a few situations where (lighter and less expensive) belts and pulleys are used to transmit movement/power in place of cables. Nominally flat belts and pulleys use only friction to transmit the power from input to output shafts. These are, unfortunately, prone to slippage so, to reduce the problem, vee-section belts were devised and yet a further improvement has seen the development of serrated or ‘toothed’ belts and pulleys, which use the principle of ‘engagement’, rather than ‘friction’, to provide drive. Some of the uses to which belt drives are put can include a change of ratio, usually in a step-down situation, as well as a simple connection between input and output shafts which are displaced by some distance. Issue 1 - 20 March 2001

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The simple belt and pulley system (refer to Fig. 95), has a step-up or step-down facility, depending on which pulley is driven. It will give a mechanical advantage of 2:1 if the smaller pulley is driven, due to it being half the diameter of the larger pulley. The larger pulley will rotate at half the speed of the smaller one, and can be driven using half the torque.

Simple Belt and Pulley System Fig. 95 Some uses of belt and pulley installations in aviation can include the driving of propellers on micro-light aircraft, which use high-revving engines. These engines rotate about 6000 rpm whilst propellers are most efficient at around 2000 – 2500 rpm. Therefore the drive from the crankshaft pulley, via a strong wide belt to the propeller pulley, gives a step down ratio of about 2.5:1 on most of this type of aircraft. Another application of belt drives is on certain piston-engined helicopters, which use a belt to connect the output pulley on the end of the crankshaft to the transmission and rotor. The tension pulleys, which bear onto the belt, keep it at the correct tension for normal use. When starting-up, the tension can be totally released, allowing the engine to be started without the load of the rotors and transmission. In an emergency the released tension allows the rotors to free-wheel (autorotate) and, thus, enables a safe landing. There are a number of places inside piston engines where toothed belts, are used to drive camshafts and other accessories from the crankshaft.

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17.2 GEARS There are a number of different types of gears (refer to Fig. 96), all of which are designed for specific tasks. They will all transmit the rotary motion of the input shaft to an output shaft, but the angle between them, their direction of rotation and the ratio of their speeds, depends on the type of gears being used.

Spur

Bevel

Key Worm and Wheel

Basic Forms of Gears Fig. 96

17.2.1 GEAR TRAINS AND GEAR RATIOS

A gear ‘train’ consists of two (or more) gear wheels, running in series, on separate, parallel, shafts such that one gear transmits its drive to the other. Gear trains can change the direction of rotation and can also alter the speed of the output shaft. The speed of rotation is dependent on the ratio between the number of teeth of the input gear to that of the output gear (the Gear Ratio). Issue 1 - 20 March 2001

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If, for example, the input gear has 25 teeth and the output gear has 75 teeth, then the output speed will be in the ratio of 25:75, or one third of the input speed. Conversely, if the input gear has 20 teeth and the output gear has 10 teeth, then the output speed will be in the ratio of 20:10, or twice that of the input speed. Gear trains may be used in a variety of ways, to change the direction of rotation or to increase or decrease the speed of the relevant output gear (and its shaft). The design of a gear train will be influenced by the amount of space available to accommodate the desired effect and by the power which is to be transmitted through the gears. 17.2.2 SPUR GEARS

The teeth of Spur gears are ‘straight cut’, which means that the teeth are cut parallel with the axis of the shaft. Straight cut spur gears are comparatively easy to manufacture but are noisy in operation. Spur gears form the simplest of gear ‘trains’. 17.2.3 HELICAL GEARS

Helical gears are also used to transmit drive between parallel shafts. They are more complex to manufacture and are quieter in operation than spur gears but (unlike spur gears), helical gears produce an axial load on their respective bearings. Another advantage of helical gears however, is that there are more teeth in mesh, to provide a larger contact area than straight cut gears, on wheels of the same width. This means that helical gears can transmit more power than straight gears of the same axial width. 17.2.4 BEVEL GEARS

Bevel gears are, generally, used to transmit the drive between shafts which have intersecting axes. The angle of intersection (and thus the drive) will vary with individual applications. Bevel gears can be found in many places, an example of which could be that, taken from the main drive shaft of an aircraft engine, .to drive an accessory gearbox. 17.2.5 WORM AND WHEEL GEARS

The worm and wheel gear set consists of a helically-cut, worm gear, on an input shaft, driving a spur gear-mounted wheel, on an output shaft. The axes of the two shafts cross at 90° and are in different planes. The main difference between this configuration and the bevel gears is that the worm and wheel combination gives a much larger ‘step-down’ between the ‘driver’ and ‘driven’ shaft speeds where space is limited, though frictional losses are higher with the worm and wheel arrangement. This configuration can only be used to drive one way; i.e. the input and output are always the same. This allows the input system to drive the output slowly and with a high mechanical advantage (higher torque), without any back loads being able Issue 1 - 20 March 2001

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to drive the system in reverse. This is ideal for aircraft Flap Control systems, which have to be ‘driven’ in both directions (up and down)), via an electric or hydraulic motor, but in which the air loads, on the flaps, must not be allowed to drive them in an opposite direction. 17.2.6 PLANETARY (EPICYCLIC) REDUCTION GEAR TRAIN

The Planetary or Epicyclic gear train (refer to Fig. 97), is typical of a gear train which is used to reduce the speed of an aircraft engine’s output shaft to a more acceptable speed for its propeller. It has the advantage of putting the output shaft (the propeller), in line with the input shaft (the engine shaft). This configuration is far more efficient than a series of spur gears, as it results in a smaller frontal area being necessary for the power unit and the subsequent reduction in aerodynamic drag. It should also be made clear, that neither the number of teeth on the planetary gears, nor the number of gears on the spider affect the actual gear reduction. For example, if the ring gear has 72 teeth and the sun gear has 36 teeth, then the overall ratio remains at 2:1.

SPIDER

Planetary (Epicyclic) Reduction Gear Train Fig. 97


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17.2.7 SPUR AND PINION REDUCTION GEAR TRAIN

The smaller, of a high-ratio pair of spur gears, is referred to as the ‘Pinion’, while the larger remains the ‘Spur’ and spur and pinion gear arrangements also vary, depending on the desired results. Where the drive pinion is located inside the spur-cut ring gear (refer to Fig. 98) it has the advantage of not only stepping down the ratio of input to output but also (as can be seen), both gears rotate in the same direction. Considerable space is also saved, compared to a system using two, externallycut gears, for a similar reduction in output speed. Drive Gear (Pinion)

Direction of Rotation

Driven Gear (Spur)

Spur and Pinion Reduction Gear Train Fig. 98 17.2.8 ACCESSORY UNIT DRIVES

Aircraft engines also employ multiple gear trains (refer to Fig. 99), in their internal and external gearboxes. These provide the drives for accessories such as fuel, hydraulic and oil pumps, electrical generators, engine speed indicators and many other devices Here it can be seen that ‘idler’ gears are added to reverse the rotation and possibly to alter the final ratio of several drives and, while the majority of the gears are of spur and helical configuration, the drive from the engine shaft, to the gearbox, has bevel gears.


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Typical External Accessory Gearbox Fig. 99 17.2.9 MESHING PATTERNS

Because of the high power being transmitted by gears in certain situations and keeping in mind that (using spur gears) only one tooth at a time can be subjected to that power, then the point of contact between the teeth in mesh is very important. Helical gears may have as many as 5 teeth in contact at any one time, therefore power will be spread across more teeth. The loads must be applied mid-way between the front and rear faces of the gear wheel. They must also be exerted between 1/3 and 2/3 of the distance between the root and tip of the gear tooth. These settings and adjustments have to be attended to during the build-up of the gearbox and are usually achieved with the use of appropriately sized shims. Issue 1 - 20 March 2001

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17.3 CHAINS AND SPROCKETS Chains, for aircraft use, are generally the simple roller type that consist of outer and inner plates, rollers, bearing pins and bushes (refer to Fig.100). Chains may be one of four standard sizes but, for most aircraft installations, the manufacturer dictates the size and type of chains used. They are obtained as complete, proof-loaded, units from manufacturers, and are identified by their allocated part numbers in the relevant aircraft IPC. Chain links or attachments should never be drilled and re-riveted. Where chains have bolts in place of rollers and rivets, then the split pins must be replaced BUT, if the nuts have been ‘peened’, then the nut and bolt must be replaced before re-assembly

Typical Chain Parts and Terminology Fig. 100 The chain’s main purpose is to transfer motion from one point, to another, remote, point where the input motion is replicated. An example of this would be found in the input action of moving a control lever, on the flight deck of an aircraft, and the subsequent output action of the movement of a control surface. Most installations use chains to generate and convert rotary motion at each end, but use cables to connect the chains together over long distances. After installation in the aircraft, the chains should be examined for freedom from twist. Particular attention must be paid in instances where the attachment is made to threaded rods by means of screwed end connectors. Care should also be take to ensure the chain is not pulled out of line by the chain wheel. The wheel should engage smoothly and evenly with the wheel teeth.


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17.3.1 TYPICAL ARRANGEMENTS - CHAIN ASSEMBLIES

Chain assemblies may be used in various arrangements (refer to Fig. 101) and can be employed to provide simple rotary-to-straight line motion or to change the direction of straight line motion in one plane. A change of direction in two planes can be achieved by the use of a special ‘bi-planar block’.

Typical Arrangements of Chain Assemblies Fig. 101


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17.4 MAINTENANCE INSPECTIONS Chain assemblies should be inspected for serviceability at periods specified in the AMM. The continued smoothness of operation between chain and chain wheel or pulley should be checked for wear. This can be confirmed by trying to lift the links off the wheel teeth and checking the links for looseness. The chain should also be checked for damage, cleanliness, correct lubrication and freedom from corrosion. If a chain is suspected of becoming elongated, it should be removed, cleaned and subjected to a specified tensile load. Its length is then measured and this measurement is compared with its nominal length when it was new. Should the difference indicate an extension of 2% or more, in any section of the chain, then the chain assembly will be required to be replaced. One of the most common operations done on chain assemblies is that of checking the tension of the chain. Care must be taken not to twist the end fittings when re-tightening the lock nuts, which butt up against the end connectors


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18 CONTROL CABLES Cables, used in aircraft control systems, comply with a number of British and American Standards and are ‘preformed’ during manufacture. Preforming is a process in which each strand is formed into the shape that it will take up in the completed cable. This makes the cable more flexible, easier to splice and less prone to kinking. Another advantage of preformed cables is that, in the event of a wire breaking, it will lie flat within its strand, so that the cable should be less likely to jam in its pulleys and fairleads. Preformed cables are manufactured from either galvanised carbon steel or corrosion-resistant steel, and are impregnated with friction-preventive lubricant during manufacture. Non-preformed single strand cable may be found on some minor aircraft systems. Aircraft cables are usually classified by either their minimum breaking load or nominal diameter. It is very rare for a cable to be manufactured by an operator. They are normally ordered through the aircraft’s IPC, and the aircraft manufacturer supplies the cable fully formed with the necessary end-fittings and to the correct load factor. 18.1 TYPES OF CABLES The construction of the cable is determined by the number of strands it contains, and the number of wires in each strand (refer to Fig. 102). For example a cable designated as 7 x 19, consists of 7 strands, each containing 19 wires. The two most common forms of construction are the flexible and the extra-flexible types.

1x7 Non-Flexible

1x19 Non Flexible

7x19 Extra Flexible

7x7 Flexible

Common Forms of Aircraft Control Cables Fig. 102


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18.2 CABLE SYSTEM COMPONENTS There are many components associated with cable systems and a selection is presented here merely for information. They include: •

End-Fittings

•

Turnbuckles

•

Tensioning Devices

•

Fairleads

•

Pulleys

18.2.1 END-FITTINGS

Whilst cables were, previously, ‘spliced’ or ‘whipped’, to form end-fittings, the majority of modern cables have a ‘swaged splice’ end-fitting. Most end-fittings, on control cables, are special-to-type and end-fittings such as fork, threaded (internal and external), and ball end-fittings (refer to Fig. 103) can be found in various locations. The nominal overall length of a cable will depend on the type of endfitting which is being employed.

Overall Length

Aircraft Cable End-Fittings Fig. 103


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18.2.2 TURNBUCKLES

Turnbuckles are devices which are attached (via internal or external threads) to appropriately designed end-fittings of aircraft cables and are used to join lengths of cables and to adjust the tension of those cables. Cable runs that are too tight will make the controls stiff to operate and, conversely, cables that are too slack will make the controls sloppy and unresponsive. Turnbuckles are adjusted by the use of a ‘left-hand’ thread in one end of the turnbuckle, and a ‘right-hand’ thread in the other end (refer to Fig. 104). When the centre part of the turnbuckle is rotated, its length will increase or decrease, and so it will adjust the cable tension. The groove, around one end of the turnbuckle barrel, indicates the ‘left hand thread’. Once the correct tension has been obtained and confirmed (using a cable tensiometer), the turnbuckle is checked for ‘safety’ (sufficient threads are engaged in the turnbuckle) and the device is then securely locked. The spring type of locking clip (used in place of locking wire) can only be inserted into the turnbuckle when the corresponding longitudinal grooves in the barrel and end fittings are aligned. Spring Locating Clips

Groove

Spring-Locked Turnbuckle Fig. 104


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18.2.3 CABLE TENSIONING DEVICES

Although the cable tension can be correctly adjusted on the ground, that set tension may alter once the aircraft is in flight. This can be due to the large temperature differentials involved - particularly with larger aircraft, which fly at high altitudes and are capable of experiencing various climates in one flight – and the consequences of an expanding, contracting and flexing airframe. To overcome these problems a tension regulator is installed in some control runs. As previously stated, engineers will use a tensiometer to set and check the tension of a cable. The tension regulator (refer to Fig. 105) is a device which has springs, incorporated within the mechanism, to ensure that the cable tension remains constant, regardless of the flexing and temperature changes of the airframe. Cable Tension Regulators can be very dangerous, when disconnecting cable runs, so it is important to ensure that they are locked or ‘snubbed’, in accordance with the AMM, before any work is done on the controls.

Tension Regulator and Cable Tensiometer Fig. 105


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18.2.4 CABLE FAIRLEADS

The cables of a control run must be supported otherwise they may foul the airframe structure. They are supported by fairleads (refer to Fig. 106), which are usually made from fibre. These fairleads should not be lubricated as this will collect dirt and dust, which will cause extra wear on the cable and fairlead. Where a change in direction of the cable is required, a pulley is normally used, due to its low friction in comparison with fairleads. Guards are fitted to pulleys when the risk of the cable riding off the pulley is high. The fairleads, already mentioned, simply allow the cable to pass through the bulkheads without chafing. If, however, the bulkhead is the divider between the pressure cabin and the outside air pressure, then the fairlead will be designed to be an airtight seal, as well as a cable guide.

Cable Fairleads Fig. 106


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18.2.5 PULLEYS

Cables that run from the flight deck to the control surfaces, require the ability to change direction (possibly a number of times). If the cable needs to change direction to another angle, the conventional method of a pulley allows this change with little friction. The example of the elevator flying control run of a simple aircraft, (refer to Fig. 107), has pulleys that can change the direction of the cable through a large range of angles.

A Simple Elevator Control Run Fig. 107


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18.3 FLEXIBLE CONTROL SYSTEMS Normal aircraft cables are only capable of performing a pulling action, due to their lack of rigidity, so, where a two-directional movement (push/pull) is required it would be necessary either to employ the use of rods, with the attendant weight penalty, or to use flexible control systems. The two most common are: •

Bowden Cables

•

Teleflex Control Systems

18.3.1 BOWDEN CABLES

The Bowden system of control consists of a stainless steel wire, housed in a flexible sleeve or conduit (refer to Fig. 108). The control is intended for pull operation only, with the cable being returned, on release of the control lever, by a return spring. The transmitting end of the cable is attached to the actuating lever whilst, at the receiving end, the cable is secured to the component to be operated. The flexible cable is made up of several strands of stainless steel wire with nipples soldered onto the end of the wire. The nipples are of different shapes, depending on their use. The flexible conduit consists of close-coiled wire, covered with cotton braiding and a waterproof coating. For long runs, or runs not requiring flexibility, the Bowden cable is fed through rigid metal tubing, which can be bent over large radius curves if required.

Bowden Cable Fig. 108


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The control fittings are used at each end of the cable to transmit and actuate the movement required. These fittings are the hand levers and adjustable stops (refer to Fig. 109). The illustration shows a simplified set-up of a Bowden cable control, with an operating lever and an adjustable stop. The double-ended stop is used if the component does not permit access to the stop at that end of the cable. At points along the conduit, connectors may be found which allow the conduits to be separated for maintenance. Junction boxes are also used, to permit either more than one input, to actuate a single operating lever, or one input to operate a number of operating mechanisms.

Bowden Cable Controls Fig. 109


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18.3.2 TELEFLEX CONTROL SYSTEMS

The Teleflex control system differs from most other cable control systems in that, rather than have a pair of cables (both of which operate in tension only), the Teleflex system allows a single, flexible cable to operate in both push and pull mode, without the need for a return spring. Examples of the types of systems, operated by Teleflex controls, are engine and propeller controls, trimming controls and fuel valves. Teleflex controls can also be used to transmit movement from one place to another, such as in a mechanical Flap Position indicator or as interlocks between controls and throttles during control lock operation. Like the Bowden system, described previously, the Teleflex system consists of a flexible transmitting cable operating inside a rigid or flexible metal conduit. The main advantages are that it provides a more accurate and positive control throughout the range of movement and the controlled component can be temporarily locked in any desired position. The control cable is a unique design of a helically-wound high-tensile steel wire (‘left’ or ‘right handed’ coil). The ‘pitch’ of the wire coil is designed to engage with gear teeth of the control units and the end-fittings (refer to Fig. 110).

Two Types of Teleflex Control Cable Fig .110


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The conduits operate in a similar manner to the Bowden system and are made from copper, aluminium or steel. The linings are of PTFE on most conduits except in high temperature areas like engine bays. To operate the system, the cable and conduit are connected to control units at each end of the control run and, in between, to other units and fittings, which are used to direct the run. In many locations, the cables are attached to leveroperated wheel units or to push-pull handles. At the receiving end of the run, another wheel unit or sliding end-fitting is used to actuate the mechanism. The Teleflex system allows a variety of controls to operate a wide selection of end-fittings (refer to Fig. 111).

Examples of Teleflex Control Cable Runs Fig. 111


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19 ELECTRICAL CABLES & CONNECTORS An electrical circuit has at least three elements: 1. Source of electrical power. 2. Load device to use the electrical energy. 3. Cables incorporating a "Conductor" to connect the source to the load. Cables must provide a path for the flow of electrons from the source, through the load and back to the source with the minimum resistance. Additionally, two other important factors for a conductor are: Ability to carry a specific load. Reliability under operating conditions. Two materials considered to be excellent electrical conductors are "Copper" and "Aluminium", both are used extensively in aircraft installations. 19.1 CABLE SPECIFICATION A large number of specifications exist for aircraft electrical cables. The majority of cables used on British built aircraft now in service will have been produced to "Aerospace G" series of British Standards. 19.2 CABLE IDENTIFICATION This covers cable type, size, manufacturer and year of production. It is important to be able to distinguish between the different types of cable and the size of the core. One of the main difficulties is the extensive use of nylon and terylene braids over the basic insulation of many cables giving them a similar appearance. Cables are stamped with the name and size of the cable at intervals along its length. If the cable is too thin to be printed on, the code will be printed on a nonmetallic sleeves positioned along the cable.


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There are many different types of wire used for special applications in aircraft electrical systems, but the majority of the wiring is achieved with MIL-W-5086 or MIL-W-22759 stranded tinned copper wire with PVC, nylon or Teflon insulation. Figure 1 shows an example of MIL-W-5086 copper wire.

TINNED COPPER CONDUCTOR

POLYVINYL CHLORIDE INSULATION

EXTRUDED NYLON JACKET

MIL-G-5086 Copper Wire Figure 1 Where large amounts of current must be carried for long distances, MIL-W-7072 aluminium wire is often used. This wire is insulated with either "Fluorinated Ethylene Propolene (FEP), nylon or fibreglass braid. Aluminium wire smaller than six-gauge is not recommended because it is so easily broken by vibrations. Anytime a wire carries a current, a magnetic field surrounds the wire, and this field may interfere with some aircraft instrumentation. For example, the light that illuminates the compass card of a magnetic compass is powered with low-voltage DC. The field from this small voltage can deflect the compass. To minimise this occurrence, a two-conductor twisted wire is used to carry the current to and from this light. By using a twisted wire, the fields cancel each other out and thus do not interfere with the compass.


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AC or pulsating DC has an especially bad effect on electronic equipment, as its conductor’s radiate electrical energy much like the antenna of a radio. To prevent radio interference, wires that carry AC or pulsating DC are often shielded. Encasing the conductor in a wire braid carries this out. This ensures that the radiated energy is received by the braid and is then passed to the aircraft's ground where it can cause no interference. Figure 2 shows a shielded wire.

TINNED COPPER CONDUCTOR

POLYVINYL CHLORIDE INSULATION

EXTRUDED NYLON JACKET

TINNED COPPER SHIELD

Shielded Wire Figure 2


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Antennas are connected to most of the radio receivers and transmitters with a special type of shielded wire called "Coaxial Cable". This consists of a central conductor surrounded by an insulator and a second conductor. The spacing and concentricity of the two conductors are critical for the most efficient transfer of energy through the cable. This second conductor is normally the wire braid, which is then covered in an outer insulator. Figure 3 shows a coaxial cable.

SOLID CENTER CONDUCTOR

INNER INSULATOR

OUTER INSULATOR JACKET

BRAID OUTER CONDUCTOR

Coaxial Cable Figure 3


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uk engineering 19.3 DATA BUS CABLE

One special type of cable used exclusively for various digital electronic systems is called “Data Bus Cable”. Data bus cable typically consists of a twisted pair of wires surrounded by electrical shielding and insulators. Digital systems operate on different frequencies, voltages and current levels. It is extremely important to ensure that the correct cable is used for the system installed. The cable should not be pinched or bent during installation and data bus cable lengths may also be critical. Refer to current manufacturer’s manuals for cable specifications. Figure 4 shows an example of a data bus cable.

TINNED COPPER CONDUCTORS

DATA BUS CABLE “B” DATA BUS CABLE “A”

ETFE TEFZEL® INSULATION

ETFE TEFZEL® JACKET

TINNED COPPER BRAID SHIELD

Data Bus Cable Figure 4


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19.4 CONDUCTOR MATERIAL & INSULATION The wires installed in an aircraft electrical system must be chosen on the basis of their ability to carry the required current without overheating and to carry it without producing an excessive voltage drop. There are a number of factors to consider when choosing the correct wire, these are: 1. Conductor material. 2. Flexibility of the wire. 3. Insulation material. 4. Diameter of the wire (American Wire Gauge – AWG). 5. Length of wire. 6. Type of installation. For aircraft, the wire material could be either copper or aluminium. If the conductor is made from copper, the individual strands of wire are typically plated to protect the copper from corrosion. Figure 5 shows two types of conductor found in aircraft systems.

STRANDED CONDUCTORS

SOLID CONDUCTOR

Electrical Wire Type Figure 5


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The number of strands that make up the wire and the type of insulation on the wire typically determine the flexibility of a conductor. The type of insulation is very important; various insulations have different ratings for heat, abrasion and flexibility. The length and type of installation are factors established by the aircraft manufacturer. 19.5 WIRE SIZE The wire used for aircraft electrical installations is sized according to the “American Wire Gauge” (AWG). The size of the wire is a function of its diameter and is indicated by a unit called “Circular Mil”. One circular mil is equal to the cross-sectional area of a 1-mil (0.001-in) diameter wire, measured in thousandths of an inch. To determine the size in circular mils of a wire, simply square the wire's diameter measured in thousandths of an inch. Figure 6 shows this concept.

1 mil2

0.001 IN

1 CIRCULAR MIL (1 cmil)

American Wire Gauge (AWG) Figure 6 In AWG only even numbers are used, small wires have higher numbers, typically starting at AWG 24. Large wires have smaller numbers, down to AWG 0000. AWG size 20 is approximately 0.032in. in diameter, and AWG 0 is approximately 0.325in. in diameter.


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To determine the size of any given wire, a wire gauge tool may be used. Figure 7 shows a typical wire gauge tool.

NOTE: AWG FOR ELECTRICAL WIRE IS 24 - 0000

Wire Gauge Tool Figure 7 19.6 WIRE RESISTANCE Resistance is the opposition to current flow and is measured in Ohms (Ω). The resistance of a wire will increase with an increase of length, but will decrease with an increase of cross-sectional area. 19.7 CURRENT CARRYING CAPABILITY A wire fitted to an aircraft system should be able to carry the required current without overheating and burning. Also it must be able to carry the required current without producing a voltage drop greater than that which is permissible for the circuit. Most aircraft wiring that is required to carry large amounts of current for long distances, is generally made up of aluminium wire. Tables 1 and 2 shows the characteristics of MIL-W-5086 copper wire and MIL-W7072 aluminium wire.


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uk engineering Wire Size

Single Wire Max Amps (In free Air)

Bundled Wire Max Amps (Conduit)

AN-20 AN-18 AN-16 AN-14 AN-12 AN-10 AN-8 AN-6 AN-4 AN-2 AN-0 AN-00 AN-000 AN-0000

11 16 22 32 41 55 73 101 135 181 245 283 328 380

7.5 10 13 17 23 33 46 60 80 100 150 175 200 225

Max Resistance Ohms/1,000ft (20°C) 10.25 6.44 4.76 2.99 1.88 1.1 0.7 0.44 0.27 0.18 0.11 0.09 0.07 0.06

Weight Pounds per 1,000ft 5.6 8.4 10.8 17.1 25 42.7 69.2 102.7 162.5 247.6 382 482 620 770

MIL-W-5086 Table 1

Wire Size

Single Wire Max Amps (In free Air)

Bundled Wire Max Amps (Conduit)

Max Resistance Ohms/1,000ft (20°C)

Weight Pounds per 1,000ft

AL-6 AL-4 AL-2 AL-0 AL-00 AL-000 AL-0000

83 108 152 202 235 266 303

50 66 90 123 145 162 190

0.64 0.43 0.27 0.17 0.13 0.11 0.09

........ ........ ........ 166 204 250 303

MIL-W-7072 Table 2


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We need to supply an actuator with 100 amps of current from a 28V system. Using tables 1 and 2, select both a copper and aluminium single wire to carry out this task. Copper wire gauge

-

..........................................

Aluminium wire gauge

-

..........................................

Note; The higher the number the smaller the wire. Now select a wire for the above task that will be routed within a bundle. Copper wire gauge

-

..........................................

Aluminium wire gauge

-

..........................................

Note; The rule of thumb says that when substituting copper for aluminium wire, we should use wire that is two gauge numbers larger. The FAA does not allow aluminium wire smaller (in size, larger in number), than 6-gauge to be used on aircraft. 19.8 VOLTAGE DROP When we add any electrical equipment to an aircraft, we must be sure that the current flowing in the wiring does not drop the voltage below a set level. Table 3 shows an example of the allowable voltage drop for various systems using various supply voltages.

Nominal System Voltage

14 28 115 200

Allowable Voltage Drop - Volts

Continuous Operation 0.5 1 4 7

Intermittent Operation 1 2 8 14

Allowable Voltage Drop Table 3


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19.9 WIRE IDENTIFICATION Wire identification should identify the wire, with respect to, type of circuit, size of cable and location within the circuit. Coded letters identify wires within systems; Figure 8 shows a typical example of a code.

26TH WIRE IN THE CIRCUIT

22 GAUGE WIRE

FLIGHT INSTRUMENTATION 4TH SEGMENT (A= 1ST SEGMENT)

GROUND

Wire Code Figure 8 The numbers on the wire greatly facilitate troubleshooting of an electrical system. Maintenance manuals list the various codes (they can vary between aircraft).


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uk engineering 19.10


WIRE INSTALLATION AND ROUTING

In aircraft there are two groups of wiring that may be installed: Open Wiring - (Wire Groups, Bundles and Routing). Conduit - (Mechanical Protection). 19.11

OPEN WIRING

This is where the wires are bundled together and installed with no external protection. This method is used when there is no great danger of mechanical damage (Chafing, Rubbing). This type of installation is easy to install and maintain, and is lighter in weight. Wires are grouped and tied together in bundles for the neatest and most efficient routing. No one bundle should carry wires from circuits that would disable both main and back-up systems. The bundles should be routed so as not to interfere with any of the controls or moving components. They must be routed where they cannot be damaged by persons entering or leaving the aircraft or by baggage or cargo moving over them or resting on them. Figure 9 shows an example of an Open Wire bundle fitted to an aircraft sidewall. “P” CLIPS ATTACHING BUNDLE TO AIRCRAFT FRAME

WIRE BUNDLE

½ INCH MAXIMUM WITH NORMAL HAND PRESSURE

CABLE BUNDLE

“P” CLIP

Open Wire Bundle Figure 9


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WIRE & CABLE CLAMPING

Electrical cables or wire bundles are secured to the aircraft structure by means of metal clamps (P Clips/clamps), lined with a synthetic rubber or similar material. In the installation of cable clamps, care must be taken to assure that the stress applied by the cable to the clamp is not in a direction that will tend to bend the clamp. When a clamp is mounted on a vertical member, the loop of the clamp should always be at the bottom. Correct methods for installing clamps is shown in Figure 10.

DANGEROUS ANGLES

° 45 X MA

45 MA ° X

SAFE ANGLES

Correct Methods of Installing Cable clamps Figure 10


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CONDUIT

Mechanical protection can be provided for the wire by routing the bundles through either flexible or rigid conduit. The size of the conduit is normally an inside diameter 25% larger than the diameter of the wire bundle being encased. Figure 11 shows the two types of conduits.

MINIMUM BEND RADIUS (FOUR TIMES INSIDE DIAMETER)

CABL E CONDUIT

CONDUIT CLAMP

ADAPTOR

INSIDE DIAMET ER

CLAMP ADAPT OR

FLEXIBLE CONDUIT BRACKET

METALLIC CONDUIT

CABLE CLAMP

CLAMP

RIGID CONDUIT

Cable Conduit Figure 11


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All conduit, rigid and flexible, should have drain holes at the lowest point in each run, and these holes and the edges of the conduit, should have no rough edges that could damage the wiring. Figure 12 shows a bundle fitted inside conduit.

LINE REPLACEMENT UNIT (LRU)

CABLE CONDUIT

DRAIN HOLE

PLUG CONNECTION

Conduit Drain Hole Figure 12


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CONNECTORS

Most of the electrical components in an aircraft are designed so that they may be serviced with a minimum amount of time needed for their removal and installation. The electrical wiring is usually connected through quick-release plugs. There are many different types of plugs, but they are all somewhat similar. The individual wires are fastened to pins or sockets inside the plugs and are clamped tight to prevent mechanical strain on the cable being transmitted into the connectors themselves. The most commonly used connector is the Military Standard (MS), type. Each MS connector has an identification number on it, Figure 13 shows a connector and identification number.

MILITARY STANDARD TYPE NUMBER CLASS SIZE INSERT ARRANGEMENT NUMBER CONTACT STYLE INDEX SLOT

INSERT NUMBER

Connector Identification Number Figure 13


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The MS type number is the basic configuration of the connector: MS3100 - Wall Receptacle. MS3101 - Cable Receptacle. MS3102 - Box Receptacle. MS3106 - Straight Plug. MS3108 - Angle Plug. The letter following the configuration tells the class of connector: A - General purpose, solid aluminium alloy shell. B - General purpose, split aluminium alloy shell. C - Pressurized, solid aluminium alloy shell. D - Environmental-resistant, solid aluminium alloy shell. E - Fire and flame proof, solid steel shell. The size of the connector is indicated with a code number, the higher the number, the larger the connector. The insert arrangement is a code number to identify the number and size of the connector and its physical arrangement. The contact style may be either an "S" or "P" to indicate a "socket or "pin" (female or male), arrangement. The final letter in the identification is one of the last letters in the alphabet, "W", "X", "Y" or "Z". These letters indicate the rotation of the insert in the connector. It is possible to connect the wrong plug to a receptacle, so to prevent this, the inserts may be rotated in their relationship to the index slot. This ensures only the correct plug may be inserted into the receptacle.


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Figure 14 shows typical MS type connectors. These connectors can carry either pins or sockets in the form of inserts. It is normal practice that, if a connector carries power supplies, it will use sockets. Pins will be used for the receiver equipment. This is to eliminate the possibility of shorts circuits to ground.

MS 3100 BULKHEAD RECEPTACLE

MS 3101 CABLE RECEPTACLE

MS 3108 BULKHEAD PLUG

MS 3102 BOX RECEPTACLE

MS Quick-Release Connectors Figure 14


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CRIMPING

Crimping is a method of firmly attaching a terminal end to an electrical conductor by pressure forming or reshaping a metal barrel, together with the conductor. The forming of a satisfactory crimped joint depends on the correct combination of conductor, crimp barrel and tool. When applied with the correctly matched tool, a joint would be established which has both good electrical and mechanical properties. Figure 15 show a crimped terminal.

STRIPPED WIRE CONDUCTOR

TERMINAL RING

DIAMOND GRIP CRIMP FOR INSULATION SUPPORT

CRIMP INSULATION

CROSS CRIMP FOR GRIPPING WIRE STRANDS

WIRE INSULATION

Pre-Insulated Crimped Terminal Figure 15


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CRIMPING TOOLS

There are a number of types of crimping tool available, but the best ones have a ratchet mechanism that will not allow them to open until they have crimped the terminal to the proper size. These tools, often referred to as "Precision Termination Tools (PTT), require periodical calibration checks. If a terminal is properly crimped on the wire, the wire will break before the terminal slips off. Figure 16 shows a heavy-duty crimping tool, this is used to install pre-insulated wire terminals.

CRIMPING HEAD CRIMPING JAWS CONDUCTOR BEING CRIMPED

RATCHET MECHANISM

HANDLE

Heavy-duty Crimping Tools Figure 16


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WIRE SPLICING

The splicing of electrical wires may be done if approved for a particular installation. Typically, the splice is made with an approved crimp type connector. The “Splice connector” is a metal tube with a plastic insulator on the outside or a plain metal tube that is covered with a plastic tube after the splice has been made. The stripped wire is inserted into the end of the tube and then crimped with a terminal crimping tool. When splices are made in wires that are in a cable bundle, the spliced wires are placed on the outside of the bundle. If several splices are to be made in any cable bundle, the splices should be staggered to reduce the bundle diameter. Figure 17 shows various situations of splices in a cable bundle.

DO NOT PUT CABL E LACING ON T OP OF T HE SPL ICES

2 CM MIN IMU M

DISTRIBUT E SPILCES IN A CBLE BUNDLE EVENLY ON THE OUTSIDE OF TH E BUNDLE

3 - PH ASE POW ER SU PPLY

CABLE SIZ E AWG 8 OR LARGER

1 CM M IN IM U M

ME TAL TUB E

3 - PH A SE POW ER S UPPL Y

CABL E SIZE AWG 8 OR L ARGER

PLA STIC IN S ULATIO N

CA BLE SPLIC E CO NSTRU CTION

Cable Splices Figure 17


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BEND RADIUS

To protect the cable from undue stress, it is important to ensure that when the cable has to bent, the radius of the bend is not less than six time the radius of the cable bundle. Figure 18 shows the bend radius for a cable with connector.

CONNECTOR RADIUS AT LEAST SIX TIMES OUTER DIAMETER

STRAIGHT STRAIN RELIEF

Bend Radius Figure 18

If the cable bundle is supported at the bend (example on a terminal block, then the bend radius can be reduced to a minimum of three times the diameter of the cable bundle. Figure 20 shows a terminal block connection.

TERMINAL BLOCK

RADIUS MINIMUM OF THREE TIMES THE OUTER DIAMETER OF CABLE

Bend radius (Supported) Figure 20 Issue 1 - 20 March 2001

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EASA Part 66 - Module 6 - Materials and Hardware

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