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Page 1: EASA PART 66   Module( 6) -Materials and Hardware

MODULE (6)Materials and Hardware

Page 2: EASA PART 66   Module( 6) -Materials and Hardware

CONTENTS

1 INTRODUCTION 1-12 PROPERTIES OF METALS 2-1

2.1 BRITTLENESS 2-1

2.2 CONDUCTIVITY 2-12.3 DUCTILITY 2-1

2.4 ELASTICITY 2-1

2.5 HARDNESS 2-12.6 MALLEABILITY 2-1

2.7 PLASTICITY 2-1

2.8 TENACITY 2-12.9 TOUGHNESS 2-2

2.10 STRENGTH 2-22.10.1 Tensile Strength 2-22.10.2 Yield Strength 2-22.10.3 Shear Strength 2-22.10.4 Bearing Strength 2-2

3 TESTING OF MATERIALS 3-13.1 TENSILE TESTING 3-1

3.1.1 Tensile Strength 3-13.2 LOAD/EXTENSION DIAGRAMS 3-4

3.2.1 Ductility 3-73.2.2 Proof Stress 3-7

3.3 STIFFNESS 3-9

3.4 TENSILE TESTING OF PLASTICS 3-9

3.5 COMPRESSION TEST 3-103.6 HARDNESS TESTING 3-10

3.6.1 Brinell Test 3-103.6.2 Vickers Test 3-113.6.3 Rockwell Test 3-113.6.4 Hardness Testing on Aircraft 3-12

3.7 IMPACT TESTING 3-13

3.8 OTHER FORMS OF MATERIAL TESTING 3-143.8.1 Creep 3-143.8.2 Creep in Metals 3-143.8.3 Effect of Stress and Temperature on Creep 3-153.8.4 The Effect of Grain Size on Creep 3-163.8.5 Creep in Plastics 3-163.8.6 Fatigue 3-163.8.7 Fatigue Testing 3-17

3.9 S-N CURVES 3-18

3.10 CAUSES OF FATIGUE FAILURE 3-20

3.11 VIBRATION 3-20

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3.12 FATIGUE METALLURGY 3-21

3.13 FATIGUE PROMOTERS 3-223.13.1 Design 3-223.13.2 Manufacture 3-233.13.3 Environment 3-23

3.14 FATIGUE PREVENTERS 3-233.14.1 Cold Expansion (Broaching) 3-24

3.15 DO'S AND DONT'S - PREVENTING FATIGUE FAILURES 3-25

3.16 STRUCTURAL HEALTH MONITORING (SHM) 3-253.16.1 Fatigue Meters 3-253.16.2 Strain Gauges 3-253.16.3 Fatigue Fuses 3-253.16.4 Intelligent Skins Development 3-25

4 AIRCRAFT MATERIALS - FERROUS 4-14.1 IRON 4-1

4.1.1 Cast Iron 4-14.1.2 Nodular Cast Iron 4-1

4.2 STEEL 4-14.2.1 Classification of Steels 4-24.2.2 Metallurgical Structure of Steel 4-34.2.3 Structure and Properties - Slow-Cooled Steels 4-34.2.4 Effects of Cooling Rates on Steels 4-4

4.3 HEAT-TREATMENT OF CARBON STEELS 4-44.3.1 Associated Problems - Hardening Process 4-54.3.2 Tempering 4-64.3.3 Annealing 4-64.3.4 Normalising 4-6

4.4 SURFACE HARDENING OF STEELS 4-74.4.1 Carburising 4-74.4.2 Nitriding 4-84.4.3 Flame/Induction Hardening 4-84.4.4 Other Surface Hardening Techniques 4-8

4.5 ALLOYING ELEMENTS IN STEEL 4-9

4.6 CARBON 4-94.6.1 Low-Carbon Steel 4-94.6.2 Medium-Carbon Steel 4-94.6.3 High-Carbon Steel 4-9

4.7 SULPHUR 4-94.8 SILICON 4-9

4.9 PHOSPHORUS 4-10

4.10 NICKEL 4-104.10.1 Nickel Alloys 4-10

4.11 CHROMIUM (CHROME) 4-114.11.1 Nickel-Chrome Steel and its Alloys 4-11

4.12 COBALT 4-11

4.13 VANADIUM 4-12

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4.14 MANGANESE 4-12

4.15 MOLYBDENUM 4-12

4.16 CHROME AND MOLYBDENUM 4-124.17 TUNGSTEN 4-13

4.18 MARAGING STEELS 4-135 AIRCRAFT MATERIALS - NON-FERROUS 5-1

5.1 PURE METALS 5-15.1.1 Pure Aluminium 5-15.1.2 Pure Copper 5-25.1.3 Pure Magnesium 5-25.1.4 Pure Titanium 5-2

5.2 ALUMINIUM ALLOYS 5-35.3 IDENTIFICATION OF ELEMENTS IN ALUMINIUM ALLOYS 5-3

5.4 CLAD MATERIALS 5-5

5.5 HEAT-TREATMENT OF ALUMINIUM ALLOYS 5-55.5.1 Solution Treatment 5-65.5.2 Age-Hardening 5-75.5.3 Annealing 5-75.5.4 Precipitation Treatment 5-8

5.6 IDENTIFICATION OF HEAT-TREATED ALUMINIUM ALLOYS 5-9

5.7 MARKING OF ALUMINIUM ALLOY SHEETS 5-105.8 CAST ALUMINIUM ALLOYS 5-11

5.9 MAGNESIUM ALLOYS 5-115.10 COPPER ALLOYS 5-12

5.11 TITANIUM ALLOYS 5-13

5.12 WORKING WITH TITANIUM AND TITANIUM ALLOYS 5-135.12.1 Drilling Titanium 5-14

6 METHODS USED IN SHAPING METALS 6-16.1 CASTING 6-1

6.1.1 Sand-Casting 6-16.1.2 Advantages/Disadvantages of Sand-Casting 6-36.1.3 Typical Casting Defects 6-36.1.4 Shell-Moulding 6-36.1.5 Centrifugal-Casting 6-36.1.6 Die-Casting 6-46.1.7 Investment-Casting (Lost Wax) 6-4

6.2 FORGING 6-56.2.1 Drop-Stamping 6-66.2.2 Hot-Pressing 6-66.2.3 Upsetting 6-6

6.3 ROLLING 6-76.4 DRAWING 6-7

6.5 DEEP DRAWING/PRESSING 6-7

6.6 PRESSING 6-76.7 STRETCH-FORMING 6-7

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6.8 RUBBER-PAD FORMING 6-7

6.9 EXTRUDING 6-86.9.1 Impact-Extrusion 6-8

6.10 SINTERING 6-8

6.11 SPINNING 6-9

6.12 CHEMICAL MILLING 6-96.13 ELECTRO-CHEMICAL MACHINING 6-9

6.14 ELECTRO-DISCHARGE MACHINING E.D.M 6-10

6.15 CONVENTIONAL MACHINING 6-116.16 SUPERPLASTIC FORMING 6-12

7 AIRCRAFT MATERIALS - COMPOSITE AND NON-METALLIC 7-17.1 PLASTICS 7-1

7.1.1 Thermoplastic Materials 7-27.1.2 Thermosetting Materials 7-37.1.3 Resins 7-47.1.4 Elastomers 7-6

7.2 PRIMARY ADVANTAGES OF PLASTICS 7-7

7.3 PRIMARY DISADVANTAGES OF PLASTICS 7-77.4 PLASTIC MANUFACTURING PROCESSES 7-8

7.5 COMPOSITE MATERIALS 7-97.5.1 Glass Fibre Reinforced Plastic (GFRP) 7-97.5.2 Carbon Fibre Reinforced Plastic (CFRP) 7-107.5.3 Aramid Fibre Reinforced Plastic (AFRP) 7-117.5.4 General Information 7-117.5.5 Laminated, Sandwich and Monolithic Structures 7-12

7.6 NON-METALLIC COMPONENTS 7-137.6.1 Seals 7-13

8 DETECTING DEFECTS IN COMPOSITE MATERIALS 8-18.1 CAUSES OF DAMAGE 8-18.2 TYPES OF DAMAGE 8-1

8.3 INSPECTION METHODS 8-38.3.1 Visual Inspection 8-38.3.2 Ring or Percussion Test 8-38.3.3 Ultrasonic Inspection 8-38.3.4 Radiography 8-3

8.4 ASSESSMENT OF DAMAGE 8-49 BASIC COMPOSITE REPAIRS 9-1

9.1 REPAIR OF A SIMPLE COMPOSITE PANEL 9-2

9.2 REPAIR OF A SANDWICH PANEL 9-39.3 GLASS FIBRE REINFORCED COMPOSITE REPAIRS 9-5

9.4 TYPES OF GLASS REINFORCEMENT 9-59.4.1 Uni-Directional Cloth 9-59.4.2 Bi-directional Cloth 9-69.4.3 Chopped Strand Mat 9-69.4.4 Resin 9-6

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9.5 POT LIFE 9-7

9.6 CURING 9-7

9.7 GEL COAT 9-89.8 STORAGE OF GFRP MATERIALS 9-8

9.8.1 Storing Resin 9-89.8.2 Storing Hardener 9-89.8.3 Storing Fabrics 9-8

9.9 PREPARATION FOR REPAIR 9-99.9.1 Surface Preparation 9-11

9.10 TECHNIQUES OF LAMINATING GLASS FIBRE 9-11

9.11 PRE-WETTING GLASS FIBRE 9-1210 ADHESIVES AND SEALANTS 10-1

10.1 THE MECHANICS OF BONDING 10-110.1.1 Stresses on a Bonded Joint 10-110.1.2 Advantages of Adhesives 10-310.1.3 Disadvantages of Adhesives 10-310.1.4 Strength of Adhesives 10-4

10.2 GROUPS AND FORMS OF ADHESIVES 10-410.2.1 Flexible Adhesives 10-410.2.2 Structural Adhesives 10-410.2.3 Adhesive Forms 10-4

10.3 ADHESIVES IN USE 10-510.3.1 Surface Preparation 10-510.3.2 Final Assembly 10-510.3.3 Typical (Abbreviated) Process 10-6

10.4 SEALING COMPOUNDS 10-610.4.1 One-Part Sealants 10-710.4.2 Two-Part Sealants 10-710.4.3 Sealant Curing 10-7

11 CORROSION 11-111.1 CHEMICAL (OXIDATION) CORROSION 11-1

11.1.1 Effect of Oxide Thickness 11-211.1.2 Effect of Temperature 11-311.1.3 Effect of Alloying 11-4

11.2 ELECTROCHEMICAL (GALVANIC) CORROSION 11-511.2.1 The Galvanic Cell 11-511.2.2 Factors Affecting the Rate of Corrosion in a Galvanic Cell 11-6

11.3 TYPES OF CORROSION 11-811.3.1 Surface Corrosion 11-811.3.2 Dissimilar Metal Corrosion 11-811.3.3 Intergranular Corrosion 11-911.3.4 Exfoliation Corrosion 11-1011.3.5 Stress Corrosion 11-1011.3.6 Fretting Corrosion 11-1111.3.7 Crevice Corrosion 11-1111.3.8 Filiform Corrosion 11-1111.3.9 Pitting Corrosion 11-12

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11.3.10 Corrosion Fatigue 11-1311.3.11 Microbiological Contamination 11-1311.3.12 Hydrogen Embrittlement of Steels 11-13

11.4 FACTORS AFFECTING CORROSION 11-1411.4.1 Climatic 11-1411.4.2 Size and Type of Metal 11-1411.4.3 Corrosive Agents 11-14

11.5 COMMON METALS AND CORROSION PRODUCTS 11-1511.5.1 Iron and Steel 11-1511.5.2 Aluminium Alloys 11-1511.5.3 Magnesium Alloys 11-1611.5.4 Titanium 11-1611.5.5 Copper Alloys 11-1611.5.6 Cadmium and Zinc 11-1611.5.7 Nickel and Chromium 11-17

11.6 CORROSION REMOVAL 11-1711.6.1 Cleaning and Paint Removal 11-1711.6.2 Corrosion of Ferrous Metals 11-1811.6.3 High-Stressed Steel Components 11-1811.6.4 Aluminium and Aluminium Alloys 11-1811.6.5 Alclad 11-1911.6.6 Magnesium Alloys 11-1911.6.7 Acid Spillage 11-2011.6.8 Alkali Spillage 11-2011.6.9 Mercury Spillage 11-21

11.7 PERMANENT ANTI-CORROSION TREATMENTS 11-2211.7.1 Electro-Plating 11-2211.7.2 Sprayed Metal Coatings 11-2211.7.3 Cladding 11-2211.7.4 Surface Conversion Coatings 11-23

11.8 LOCATIONS OF CORROSION IN AIRCRAFT 11-2311.8.1 Exhaust Areas 11-2311.8.2 Engine Intakes and Cooling Air Vents 11-2311.8.3 Landing Gear 11-2411.8.4 Bilge and Water Entrapment Areas 11-2411.8.5 Recesses in Flaps and Hinges 11-2411.8.6 Magnesium Alloy Skins 11-2411.8.7 Aluminium Alloy Skins 11-2411.8.8 Spot-Welded Skins and Sandwich Constructions 11-2511.8.9 Electrical Equipment 11-2511.8.10 Miscellaneous Items 11-25

12 AIRCRAFT FASTENERS 12-112.1 TEMPORARY JOINTS 12-112.2 PERMANENT JOINTS 12-1

12.3 FLEXIBLE JOINTS 12-1

12.4 SCREW THREADS 12-212.4.1 The Inclined Plane and the Helix 12-2

12.5 SCREW THREAD TERMINOLOGY 12-4

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12.5.1 Screw Thread Forms 12-612.5.2 Other Thread Forms 12-812.5.3 Classes of Fit 12-812.5.4 Measuring Screw Threads 12-9

12.6 BOLTS 12-1012.6.1 British Bolts 12-1012.6.2 Identification of BS Unified Bolts 12-1012.6.3 American Bolts 12-1312.6.4 Identification of AN Standard Bolts 12-1412.6.5 Special-to-Type Bolts 12-1612.6.6 Metric Bolts 12-17

12.7 NUTS 12-1812.7.1 Stiffnuts and Anchor Nuts 12-19

12.8 SCREWS 12-2212.8.1 Machine Screws 12-2212.8.2 Structural Screws 12-2412.8.3 Self-Tapping Screws 12-24

12.9 STUDS 12-2512.9.1 Standard Studs 12-26Waisted Studs 12-2612.9.3 Stepped Studs 12-2712.9.4 Shouldered Studs 12-27

12.10 THREAD INSERTS 12-2712.10.1 Wire Thread Inserts 12-2712.10.2 Thin Wall Inserts 12-28

12.11 DOWELS AND PINS 12-2912.11.1 Dowels 12-2912.11.2 Roll Pins 12-2912.11.3 Clevis Pins 12-3012.11.4 Taper Pins 12-30

12.12 LOCKING DEVICES 12-3112.12.1 Spring Washers 12-3112.12.2 Shake-Proof Washers 12-3212.12.3 Tab Washers 12-3312.12.4 Lock Plates 12-3412.12.5 Split (Cotter) Pins 12-34

12.13 LOCKING WIRE 12-3512.13.1 Use of Locking Wire with Turnbuckles 12-3712.13.2 Use of Locking Wire with Locking Tabs 12-3712.13.3 Thin Copper Wire 12-38

12.14 QUICK-RELEASE FASTENERS 12-3812.14.1 Dzus Fasteners 12-3812.14.2 Oddie Fasteners 12-3912.14.3 Camloc Fasteners 12-4012.14.4 Airloc Fasteners 12-4112.14.5 Pip-Pins 12-4112.14.6 Circlips and Locking Rings 12-4212.14.7 Keys and Keyways 12-43

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12.14.8 Peening 12-44

12.15 GLUE/ADHESIVE BONDED JOINTS 12-4512.15.1 Locking by Adhesives 12-4512.15.2 Loctite 12-4612.15.3 Synthetic Resin Adhesives 12-4612.15.4 Testing of Adhesive Joining Techniques 12-46

12.16 METAL-TO-METAL BONDED JOINTS 12-4612.16.1 Welding 12-4612.16.2 Soft Soldering 12-4712.16.3 Hard Soldering 12-47

13 AIRCRAFT RIVETS 13-113.1 SOLID RIVETS 13-113.2 RIVET IDENTIFICATION 13-2

13.2.1 Solid Rivets (British) 13-213.2.2 Rivet Identification (British) 13-313.2.3 Rivet Material Identification (British) 13-313.2.4 Solid Rivets (American) 13-513.2.5 Rivet Identification (American) 13-613.2.6 Rivet Material Identification (American) 13-6

13.3 HEAT-TREATMENT/REFRIGERATION OF SOLID RIVETS 13-713.3.1 Heat-Treatment 13-813.3.2 Refrigeration 13-813.3.3 Use of Different Types of Rivet Head 13-8

13.4 BLIND AND HOLLOW RIVETS 13-913.4.1 Friction Lock Rivets 13-1013.4.2 Mechanical Lock Rivets 13-1113.4.3 Hollow/Pull-Through Rivets 13-1213.4.4 Grip Range 13-1213.4.5 Tucker ‘Pop’ Rivets 13-1313.4.6 Avdel Rivets 13-1413.4.7 Chobert Rivets 13-1513.4.8 Cherry Rivets 13-16

13.5 MISCELLANEOUS FASTENERS 13-1613.5.1 Hi-Lok Fasteners 13-1613.5.2 Hi-Tigue Fasteners 13-1713.5.3 Hi-Shear Fasteners 13-18

13.6 SPECIAL PURPOSE FASTENERS 13-1913.6.1 Jo-Bolts 13-1913.6.2 Tubular Rivets 13-2013.6.3 Rivnuts 13-21

14 SPRINGS 14-114.1 FORCES EXERTED ON, AND APPLIED BY, SPRINGS 14-1

14.2 TYPES OF SPRINGS 14-114.2.1 Flat Springs 14-114.2.2 Leaf Springs 14-214.2.3 Spiral Springs 14-214.2.4 Helical Compression and Tension Springs 14-214.2.5 Helical Torsion Springs 14-2

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14.2.6 Belleville (Coned Disc) Springs 14-214.2.7 Torsion-Bar Springs 14-2

14.3 MATERIALS FROM WHICH SPRINGS ARE MANUFACTURED 14-214.3.1 Steels used for Cold-Wound Springs 14-214.3.2 Steels used for Hot-Wound Springs 14-314.3.3 Steels used for Cold-Rolled, Flat Springs 14-314.3.4 Non-Ferrous Metals used for Springs 14-314.3.5 Composite Materials used for Springs 14-4

14.4 CHARACTERISTICS OF TYPICAL AEROSPACE SPRINGS 14-514.5 APPLICATIONS OF SPRINGS IN AIRCRAFT ENGINEERING 14-6

15 PIPES AND UNIONS 15-115.1 RIGID PIPES 15-115.2 SEMI-RIGID FLUID LINES (TUBES) 15-2

15.2.1 Flared End-Fittings 15-215.2.2 Flare-Less Couplings 15-3

15.3 FLEXIBLE PIPES (HOSES) 15-415.3.1 Low-Pressure Hoses 15-515.3.2 Medium-Pressure Hoses 15-515.3.3 High-Pressure Hoses 15-6

15.4 UNIONS AND CONNECTORS 15-715.4.1 Aircraft General Standards (AGS) 15-815.4.2 Air Force and Navy (AN) 15-815.4.3 Military Standard (MS) 15-8

15.5 QUICK-RELEASE COUPLINGS 15-816 BEARINGS 16-1

16.1 BALL BEARINGS 16-216.1.1 Radial Bearings 16-216.1.2 Angular-Contact Bearings 16-216.1.3 Thrust Bearings 16-216.1.4 Instrument Precision Bearings 16-2

16.2 ROLLER BEARINGS 16-316.2.1 Cylindrical Roller Bearings 16-316.2.2 Spherical Roller Bearings 16-316.2.3 Tapered Roller Bearings 16-3

16.3 BEARING INTERNAL CLEARANCE 16-416.3.1 Group 2 (‘One Dot’) Bearings 16-416.3.2 Normal Group (‘Two Dot’) Bearings 16-416.3.3 Group 3 (‘Three Dot’) Bearings 16-416.3.4 Group 4 (‘Four Dot’) Bearings 16-4

16.4 BEARING MAINTENANCE 16-516.4.1 Lubrication 16-516.4.2 Inspection 16-5

17 TRANSMISSIONS 17-117.1 BELTS AND PULLEYS 17-117.2 GEARS 17-3

17.2.1 Gear Trains and Gear Ratios 17-317.2.2 Spur Gears 17-4

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17.2.3 Helical Gears 17-417.2.4 Bevel Gears 17-417.2.5 Worm and Wheel Gears 17-417.2.6 Planetary (Epicyclic) Reduction Gear Train 17-517.2.7 Spur and Pinion Reduction Gear Train 17-617.2.8 Accessory Unit Drives 17-617.2.9 Meshing Patterns 17-7

17.3 CHAINS AND SPROCKETS 17-817.3.1 Typical Arrangements - Chain Assemblies 17-9

17.4 MAINTENANCE INSPECTIONS 17-1018 CONTROL CABLES 18-1

18.1 TYPES OF CABLES 18-118.2 CABLE SYSTEM COMPONENTS 18-2

18.2.1 End-Fittings 18-218.2.2 Turnbuckles 18-318.2.3 Cable Tensioning Devices 18-418.2.4 Cable Fairleads 18-518.2.5 Pulleys 18-6

18.3 FLEXIBLE CONTROL SYSTEMS 18-718.3.1 Bowden Cables 18-718.3.2 Teleflex Control Systems 18-9

19 ELECTRICAL CABLES & CONNECTORS 19-119.1 CABLE SPECIFICATION 19-1

19.2 CABLE IDENTIFICATION 19-119.3 DATA BUS CABLE 19-5

19.4 CONDUCTOR MATERIAL & INSULATION 19-6

19.5 WIRE SIZE 19-719.6 WIRE RESISTANCE 19-8

19.7 CURRENT CARRYING CAPABILITY 19-819.8 VOLTAGE DROP 19-10

19.9 WIRE IDENTIFICATION 19-11

19.10 WIRE INSTALLATION AND ROUTING 19-12

19.11 OPEN WIRING 19-1219.12 WIRE & CABLE CLAMPING 19-13

19.13 CONDUIT 19-1419.14 CONNECTORS 19-16

19.15 CRIMPING 19-19

19.16 CRIMPING TOOLS 19-2019.17 WIRE SPLICING 19-21

19.18 BEND RADIUS 19-22

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1 INTRODUCTION

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

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 ofother materials that alter the basic characteristics to suit the specific requirementsof the aircraft designer.

A metal’s usefulness is governed principally by the physical properties itpossesses. Those properties depend upon the composition of the metal, whichcan be changed considerably by alloying it with other metals and by heat-treatment. The strength and hardness of steel, for example, can be intensified byincreasing its carbon content, adding alloying metals such as Nickel andTungsten, 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 suchfeatures as permanent magnetism, corrosion resistance and high-strength whilstoperating at elevated temperatures, are desired.

Composites make up a large part of the construction of modern aircraft. In theearly days, composites and plastics were limited to non-structural, internalcosmetic panels, small fairings and other minor parts. Today there are manylarge aircraft, which have major structural and load-carrying parts manufacturedfrom composites. Composite materials, in addition to maintaining or increasingcomponent strength, contribute to the important factor of weight saving. There arealso many modern light aircraft that are almost totally manufactured fromcomposites 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 testson sample pieces. The terminology, associated with these properties, is outlinedin the following paragraphs.

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

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

2.3 DUCTILITYThe property of being able to be permanently extended by a tensile force. It ismeasured 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 ELASTICITYThe ability of a metal to return to its original shape and size after the removal ofany distorting force. The ‘Elastic Limit’ is the greatest force that can be appliedwithout permanent distortion.

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

2.6 MALLEABILITYThe ease with which the metal can be forged, rolled and extruded withoutfracture. Stresses, induced into the metal, by the forming processes, have to besubsequently relieved by heat-treatment. Hot metal is more malleable than coolmetal.

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

2.8 TENACITYThe property of a metal to resist deformation when subjected to a tensile load. Itis proportional to the maximum stress required to cause the metal to fracture.

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2.9 TOUGHNESSThe ability of a metal to resist suddenly applied loads. A metal’s toughness istested by impact with a swinging pendulum of known mass.

2.10 STRENGTHThere are several different measurements of the strength of a metal, as may beseen from the following sub-paragraphs

2.10.1 TENSILE STRENGTHThe 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 passedits yield point.

2.10.3 SHEAR STRENGTH

The ability to resist side-cutting loads - such as those, imposed on the shank of arivet, when the materials it is joining attempt to move apart in a direction normalto 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 canbe incorporated into any design. Mechanical property data is compiled fromextensive material testing. Various tests are used to determine the actual valuesof 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 asteadily increasing load to a test specimen, causing it to stretch until it eventuallyfractures. Accurate measurements are taken of the load and extension, and theresults are used to determine the strength of the material. To ensure uniformity oftest results, the test specimens used must conform to standard dimensions andfinish as laid down by the appropriate Standards Authority (BSI, DIN, ISO etc).

The cross-section of the specimen may be round or rectangular, but therelationship between the cross-sectional area and a specified "gauge length", ofeach specimen, is constant. The gauge length, is that portion of the parallel partof the specimen, which is to be used for measuring the subsequent extensionduring and/or after the test.

3.1.1 TENSILE STRENGTH

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

S tressLoad(N)

Origina lc.s.a. (mm2 )

Note: The units of Stress may be quoted in the old British Imperial (andAmerican) units of lbf/in2, tonf/in2 (also psi and tsi), or the European and SI unitssuch as kN/m2, MN/m2 and kPa or MPa.

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Example 1A steel rod, with a diameter of 5 mm, is loaded in tension with a force of 400 N.Calculate the tensile stress.

Load400

Stress 2

400 2 2037N

2/ mmArea r 25

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 05m2, is subjected to a load

of 10 MN. Calculate the stress in the member in;

Load 10 2

(a) MN/m2 and (b) N/mm2

(a)

(b)

Stress 20MN/mArea 05

2 2 21N/mm 1MN/m So Stress20 N/mm

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As the load in the tensile test is increased from zero to a maximum value, thematerial extends in length. The amount of extension, produced by a given load,allows the amount of induced STRAIN to be calculated. Strain is calculated bymeasuring 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

Example 3

ExtensionOriginalLength

An aluminium test piece is marked with a 20 mm gauge length. It is subjected to

tensile load until its length becomes 2115 mm. Calculate the induced strain.

E xtension 21 15 - 20 115 m m

Exte ns ion 115Strain OriginalLe ngth

Exercise 3

20

00575 (nounits)

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 DIAGRAMSIf a gradually increasing tensile load is applied to a test piece while the load andextension are continuously measured, the results can be used to produce aLoad/Extension diagram or graph (refer to Fig. 1). Obviously a number of differentforms of graph may be obtained, depending on the material type and condition,but the example shows a Load/Extension diagram which typifies many metallicmaterials when stressed in tension.

Load/Extension Diagram

Fig 1

The graph can be considered as comprising two major regions. Between points 0and A, the material is in the Elastic region (or phase), i.e. when the load isremoved the material will return to its original size and shape. In this region, theextension 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 stresscausing it.

Point A is the Elastic Limit. Between this point and point B, the material continuesto extend until the maximum load is reached (at point B). In this region thematerial is in the plastic phase. When the load is removed, the material does notreturn 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 materialcontinues to extend under reduced load until it eventually fractures at point C.

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Aircraft structural designers’ interest in materials does not extend greatly beyondthe elastic phase of materials. Production engineers, however, are greatlyinterested in material properties beyond this phase, since the forming capabilitiesof 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 anyincrease in load shortly after the elastic limit is reached. The onset of increasingextension, 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 BYield Point

Load/Extension Diagram for Mild Steel

Fig. 2

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

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

Point of Fracture

ZeroElongation

(a)

Plastic Region

SmallElongation

LargeElongation

(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 (high-carbon steel

(c) represents a material with some elasticity and good plasticity (e.g. softaluminium).

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3.2.1 DUCTILITY

After fracture of a specimen, following tensile testing, an indication of materialductility is arrived at, by establishing the amount of plastic deformation whichoccurred. 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 ofthe gauge length than the reduction in area. The standard measure of ductility isto establish the percentage elongation after fracture.

Percentage elongation

Example 4

Final Extension

O riginalGauge Length100

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?

ElongationFinal Extension

Gauge Length100

176.1-150.5

1505100 17.009% 17%

3.2.2 PROOF STRESS

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

The tensile stress, which is just sufficient to produce a non-proportional elongation, equal to a specified percentage of the originalgauge length.

Usually a value of 0.1% or 0.2% is used for Proof Stress, and the Proof Stress isthen 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 (referto Fig 4) as follows:

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

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

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

Acquiring the Proof Stress from a Load/Extension GraphFig. 4

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3.3 STIFFNESS

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

StressieStrain

is a m e as ur eof s tiffne s s

This value, which is of great importance to designers, is known as ‘the Modulusof Elasticity, or Young’s Modulus’, and is signified by use of the symbol E.

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

The actual numerical value is usually large, as it is a measure of the stressrequired 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 usuallymade from sheet material. Although the basic load/extension curve for some

plastics is somewhat similar to metal curves, changes in test temperature or therate of loading can have a major effect on the actual results.

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

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3.5 COMPRESSION TESTMachines for compression testing are often the same as those used for tensiletesting, 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 tothat in the tensile test. The value of `E' is the same in compression as it is intension. Compression testing is seldom used as an acceptance test for metallic orplastic materials (except for cast iron).

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

3.6 HARDNESS TESTINGThe hardness of materials is found by measuring their resistance to indentation.Various methods are used, but the most common are those of the Brinell, Vickersand Rockwell Hardness Tests.

3.6.1 BRINELL TEST

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

Diameter (Area)of resulting

Indentation

The Brinell Hardness TestFig. 5

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3.6.2 VICKERS TEST

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

The Vickers Hardness TestFig. 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 forhard 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 thediamond hardness number is known as Rockwell C (e.g. RC 65).

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

The Rockwell Hardness TestFig. 7

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

Table 1COMPARATIVE HARDNESS VALUES

MATERIAL BHN HV ROCKWELLAluminium alloy 100 100 B 57

Mild steel 130 130 B 73Cutting tools 650 697 C 60

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

3.6.4 HARDNESS TESTING ON AIRCRAFTIt is not normal to use Brinell, Rockwell or Vickers testing methods on aircraft inthe hangar. There are, however, portable Hardness Testers, which may be usedto test for material hardness on items such as aircraft wheels, after an over-heatcondition, because the over-heat condition may cause the wheel material tobecome soft or partially annealed.

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3.7 IMPACT TESTINGThe 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 struckby a fast-moving, weighted pendulum. The energy, which is absorbed by the testpiece on impact, will give a measure of toughness. A brittle material will breakeasily and will absorb little energy, so the swing of the pendulum (which isrecorded against a calibrated scale) will not be reduced significantly. A toughmaterial will, however, absorb considerably more energy and thus greatly reducethe recorded pendulum swing.

Most materials show a drop in toughness with a reduction in temperature, thoughsome materials (certain steels in particular) show a rapid drop in toughness asthe temperature is progressively reduced. This temperature range is called theTransition Zone, and components, which are designed for use at lowtemperature, should be operated above the material’s Transition Temperature.

Nickel is one of the most effective alloying elements for lowering the TransitionTemperature of steels

Test Piece

Impact TestFig. 8

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3.8 OTHER FORMS OF MATERIAL TESTINGAlthough some of the more important forms of material testing have already beendiscussed, there are several other forms of material testing to be considered, notleast 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, ofmaterials subjected to prolonged stress. This deformation is plastic and occurseven 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 inKelvin), the creep rate is very low, but, at higher temperatures, it becomes morerapid. For this reason, creep is commonly regarded as being a high-temperaturephenomenon, associated with super-heated steam plant and gas turbinetechnology. However, some of the soft, low-melting point materials will creepsignificantly at, or a little above, ambient temperatures and some aircraftmaterials may creep when subjected to over-heat conditions.

3.8.2 CREEP IN METALSWhen 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 withtime 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 torupture. 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 CreepFig. 9

3.8.3 EFFECT OF STRESS AND TEMPERATURE ON CREEP

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

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

Although actual design data are based on the long-time tests, short-time creeptests 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 grainboundaries, creep resistance is improved by increased grain size, due to the

reduced grain boundary region per unit volume. It is mainly for this reason thatsome modern, high-performance turbine blades are being made from directionallysolidified (and, alternatively, improved single-crystal) castings.

3.8.5 CREEP IN PLASTICSPlastics are also affected by creep and show similar, though not identical,

behaviour to that described for metals. Since most plastics possess lower thermalproperties than metals, the choice of plastic for important applications, particularlyat 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 ofengineering components were caused by fatigue. Consequences of modernengineering have led to increases in operating stresses, temperatures andspeeds. This is particularly so in aerospace and, in many instances, has madethe fatigue characteristics of materials more significant than their ordinary, staticstrength properties.

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

Experience in the aircraft industry has shown that the stress cycles, to whichaircraft are subjected, may be very complex, with occasional high peaks, due togust loading of aircraft wings. For satisfactory correlation with in-servicebehaviour, full-size or large-scale mock-ups must be tested in conditions as closeas 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 couldsafely sustain a mass of 12 tons, broke when a mass of only 3 tons was raisedand lowered on the girder some 3x106 times.

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

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

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

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

Test Piece

Load

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3.9 S-N CURVESOne of the most useful end-products, from fatigue testing, is an S-N curve, whichshows, graphically, the relationship between the amount of stress (S), applied toa material, and the number of stress cycles (N), which can be tolerated beforefailure of the material.

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

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

Stress(S)

Fatigue Limit

40 - 60 % UTS

Number of Cycles(N)

A S-N Curve for a Steel MaterialFig. 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 beseen that these materials will fail if subjected to an appropriate number of stressreversals, even at very small stresses. When materials have no fatigue limit anendurance limit together with a corresponding number of cycles is quotedinstead.

It follows that components made from such materials must be designed with aspecific life in mind and removed from service at the appropriate time. Theservice fatigue lives of complete airframes or airframe members are typicalexamples of this philosophy.

An S-N Curve for an Aluminium AlloyFig. 12

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

The importance of fatigue strength can be illustrated by the fact that, in a high-cycle fatigue mode, a mere 10% improvement in fatigue strength can result in a100-times life improvement.

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3.10 CAUSES OF FATIGUE FAILUREAs the fatigue characteristics of most materials are now known (or can be

ascertained), it would seem reasonable to suppose that fatigue failure, due tolack of suitable allowances in design, should not occur.

Nevertheless, fatigue cracking occurs frequently, and even the mostsophisticated engineering product does not possess immunity from this mode offailure. Such failures are often due to unforeseen factors in design, environmentalor 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 anapplied steady stress).

A sufficient number of cycles of that fluctuating stress.

The stress fluctuations may be separated by considerable time intervals, asexperienced in aircraft cabin pressurisation, during each take-off (e.g. daily), orthey may have a relatively short time interval, such as encountered during theaerodynamic buffeting/vibration of a wing panel. The former example would beconsidered 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 causecracking of a given material, are affected by many other variables, such as stressconcentration points (stress raisers), residual internal stresses, corrosion, surfacefinish, material imperfections etc.

3.11 VIBRATIONVibration has already been quoted as being a cause of high-cycle fatigue and,because most dynamic structures are subjected to vibration, this is undoubtedlythe most common origin of fatigue. All objects have their own natural frequency atwhich they will freely vibrate (the resonant frequency). Large, heavy, flexiblecomponents vibrate at a low frequency, while small, light, stiff components vibrateat 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, criticalcomponents are not vibrated at their natural frequencies and so avoid creatingresonance.

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

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3.12 FATIGUE METALLURGYUnder the action of fatigue stresses, minute, local, plastic deformation on anatomic scale, takes place along slip planes within the material grains. If the

fatigue stresses are continued, then micro cracks are formed within the grains, inthe area of the highest local stress, (usually at or near the surface of thematerial). The micro cracks join together and propagate across the grainboundaries 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 FractureFig. 13

The resultant fractured surface often has a characteristic appearance of:

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

A rougher, typically crystalline section, which is the final rapid fracture whenthe 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 theloading 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 failureinvestigators when attempting to estimate the crack initiation and/or propagationlife. The striations are often so fine and indistinct that electron beam microscopesare required to count them.

In normal circumstances, a great deal of energy is required to `weaken' thematerial sufficiently to initiate a fatigue crack, and it is not surprising, therefore, tofind 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 relativelyquickly, and gaining in speed as the crack front not only increases in size, butalso reduces the component cross-sectional area.

A point is eventually reached (known as the 'critical crack length') at which theremaining cross-section is sufficiently reduced to cause a gross overloadingsituation, 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 themajor reasons for operators of equipment being relatively unsuccessful indetecting 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 considerableinfluence on its susceptibility to fatigue. A good designer will therefore minimisestress concentrations by:

avoiding rapid changes in section and

using generous blend radii or chamfers to eliminate sharp corners

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3.13.2 MANUFACTURE

While the designer may specify adequate blend radii, the actual product may stillbe prone to fatigue failure if the manufacturing stage fails to achieve thissometimes-seemingly unimportant drawing requirement.

Several other manufacturing-related causes of premature fatigue failure exist, themost 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 surfaceprotection, 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 thecomponent is operating in a corrosive medium. Steel (normally), has a well-defined fatigue limit on the S-N curve but, if a fatigue test is conducted in acorrosive environment, not only does the general fatigue strength dropappreciably, but the curve also resembles the aluminium alloy curve (e.g. thefatigue failure stress continues to fall as the number of cycles increases).

Other environmental effects such as fretting and corrosion pitting, erosion orelevated temperatures will also adversely affect fatigue strength.

3.14 FATIGUE PREVENTERSIf a component is prone to fatigue failure in service, then several methods ofimprovement are available, in the form of:

Quality. Correct and eliminate any failure-related manufacturing orprocessing shortcomings.

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

Geometry.

a) Increase the size (c.s.a.) to reduce the general stress level ormodify the local geometry to reduce the change in section (largeradius).

b) Modify the geometry to change the vibration frequency orintroduce 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, alternatingstress. If the most fatigue-prone areas, such as spar fastener holes, have acompression stress applied (refer to Fig. 14), they are significantly more resistantto fatigue failure.

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

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

Area around hole pre-stressedin compression

Tapered Mandrel pulledthrough fastener hole

Cold Expansion of Fastener HoleFig.14

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3.15 DO'S AND DONT'S - PREVENTING FATIGUE FAILURESDO

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 sothat the respective fatigue lives are not exceeded. Several methods have beendeveloped 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 monitorthe fatigue history of the aircraft. Fatigue meters also allow a check to be madeon the moment in time when the aircraft exceeds the design limits imposed on it.

3.16.2 STRAIN GAUGESStrain gauges may be used to monitor stress levels on specific aircraft structures.Strain gauges are thin-foil, electrical, resistor elements, bonded to the aircraftstructure. 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 failat different fatigue stresses. The electrical current, flowing through the fuse, willvary 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 bedesigned and built with a variety of sensors and systems embedded into the

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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. Themost common ferrous metal, in aircraft construction is steel, which is an alloy ofiron with a controlled amount of carbon added.

4.1 IRONIron 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 naturallyin its metallic state, but as iron ores which only contain in the range of 25% to60% iron and are mined in open-cast or open-pit mines. Iron has a great affinityfor 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 morecommonly known as rust. Iron usually has other materials added to improve itsproperties.

The first smelt from the raw ore is poured into troughs (which are said toresemble piglets suckling on a sow) and the iron is referred to as ‘pig iron’. Thepig 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 fewaircraft 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 ‘SpheroidalGraphite Iron’. It is produced by adding magnesium and nickel (or magnesium,copper and silicon) and is a tough, strong, hard-wearing material which can beused in applications where only wrought materials were used in the past (aclassic example being piston engine crankshafts).

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

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

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

The high quality steels, used in aircraft construction, are usually produced inelectric furnaces, which allow better control, than do gas furnaces, when alloying.The carbon electrodes produce an intense arc and the steel, when molten, canhave 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 metalincreases. The effect of varying amounts of carbon is truly dramatic. If carbon isprogressively 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 furtherincrease but ductility is reduced.

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

The (American) Society of Automotive Engineers (SAE) has classified steel alloyswith a four-digit numerical index system. A small extract from the SAEclassification system is shown in Table 2, where it can be seen, for example, thatone common steel alloy is identified by the designation SAE 1030. The first digitidentifies it as a Carbon-Steel, while the second digit shows that it is a PlainCarbon-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 differentclassification system.

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Table 2EXTRACT FROM THE SAE CLASSIFICATION FOR STEEL ALLOYS

1xxx Carbon Steels10xx Plain Carbon Steels2xxx Nickel Steels3xxx Nickel Chromium Steels40xx Molybdenum Steels41xx Chromium Molybdenum Steels5xxx Chromium Steels6xxx Chromium Vanadium

4.2.2 METALLURGICAL STRUCTURE OF STEEL

The amount of carbon present in steel has a major effect on the mechanicalproperties. 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 distributedin a solid solution, the metallurgical structure is called ferrite. At roomtemperature only a very small amount of carbon (0.006%) can be containedin solid solution, therefore this ferrite structure is almost pure iron. It is (notsurprisingly) soft, weak and ductile.

When 1 carbon atom chemically combines with 3 iron atoms the result iscalled 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. Aspearlite is half cementite and half ferrite, it is not surprising to find thatpearlite combines the properties of ferrite and cementite I.e. Whereas ferritewas too soft and weak - and cementite was basically strong but too hardand 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 contentexceeds this value, the excess carbon forms carbon-rich cementite areas alongthe grain boundaries, and this is known as free cementite. Such high-carbonsteels as already stated are very hard and strong but very brittle.

Mild steel has a metallurgical structure comprising approximately one thirdpearlite 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 hasbeen considered. If such steels are, however, rapidly cooled from relatively hightemperature the metallurgical structure and properties can be somewhat different.

4.3 HEAT-TREATMENT OF CARBON STEELSIf a ‘straight’ carbon steel is progressively heated from cold, a steady rise in

temperature occurs. However, at approximately 700C, there is a reduction in therate of temperature rise (a ‘hesitation’), even though the heating is continued(refer to Fig. 15). This hesitation starts at 700C and finishes at up to 200Chigher (depending on the percentage of carbon present) and, eventually, thetemperature rise speeds up and the rate of rise is similar to that which occurredbefore the hesitation.

Temperature/Time Graph for Steel Heat-TreatmentsFig. 15

The start of the hesitation is known as the ‘lower critical point’ and the end iscalled the ‘upper critical point’, and the phenomenon of the temperatureresponse is due to a change in the crystalline structure of the steel in between thetwo critical points.

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If carbon steel is heated to just above its Upper Critical Point the structure iscalled ‘Austenitic’. This structure is a solid solution of carbon in iron (i.e. all thecarbon 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 UpperCritical Point it becomes hardened.

The more carbon present, the harder the steel will be after quenching. This rapidcooling causes a change in the metallurgical structure and is called ‘Martensite’.Martensite is extremely hard but is not suitable for most engineering purposesdue to it being very brittle. For most applications it is necessary to carry out afurther 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 suitabletemperature below its Lower Critical Point followed by cooling (usuallyquenching).

The effect of this heat-treatment is to slightly reduce the hardness whilst at thesame time greatly increasing the toughness. The actual tempering temperatureused 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 hardenedcarbon steel is achievable when 0.83% carbon is present. If an even greateramount of carbon is present, the hardness continues to increase but strength willdecrease.

4.3.1 ASSOCIATED PROBLEMS - HARDENING PROCESSThe effective hardening of carbon steels depends not only on the amount of

carbon present but also on very rapid cooling from high temperature. The coolingrate mainly depends on the cooling medium, the size of tank, and the mass of theobject to be cooled.

Agitation in the cooling bath can also speed up the cooling rate and, in terms ofcooling 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 isnormally used, whereas oil or air-cooling is used on certain alloy steels. The rapidcooling rates, involved in the hardening of carbon steel, cause enormous thermalstresses in the component and distortion is commonplace. Cracking may alsooccur in some cases.

To achieve relatively uniform cooling it is sometimes necessary to immerse theobject 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 hardenedsteel whilst suffering only a modest drop in strength. Accurate temperaturemeasuring equipment, in addition to well equipped facilities, are required to dothese procedures on aerospace metals.

Table 3,however, shows that, when carbon steel is polished to a bright, cleansurface, and then slowly heated, a range of colours appears, due to a thin oxidefilm forming during the heating process. These colours are related fairly closely totemperatures. The higher temperature achieved during the tempering process,the softer (and tougher) the material will become and vice-versa.

Table 3COLOUR/TEMPERATURE RELATIONSHIP OF CARBON STEELS

COLOUR TEMPERATUREStraw 230/240cPurple 270°CBlue 300°CDark red 500C

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 theUpper Critical Limit followed by very slow cooling. In practise the slow coolingrates are achieved by cooling in the furnace or by immersing in a poor thermalconductor such as ashes. The end result is a stress-free, fully softened material,suitable for major forming operations such as deep pressing, drawing, extrudingetc.

4.3.4 NORMALISING

This process is similar to annealing, except that the cooling is done in still air. Theend result, again, is a stress-free, soft material with uniform fine grain structure.Normalising is commonly used on actual components after heavy machiningoperations (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 veryhard surface. This would, typically, be required of a component, which issubjected to high dynamic stresses, yet also has to resist surface wear andwould 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 methodsare 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 900C, while the

component is in contact with a carbon-rich medium followed by a suitable heat-treatment.

Carbon is generally absorbed into the surface of the heated steel and the rate ofpenetration is approximately 1mm in 5-6 hours. Low-carbon steels are particularlysuited to this type of treatment, as it increases the carbon content and hence thehardness locally. Various methods of carburising are used, the most commonones being:

Pack Carburising. The object is sealed in a container containing a carbon-rich (charcoal based) powder and heated in a furnace. The metal is nextquenched 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 isheated.

Gas Carburising. The object is placed in a basket in a furnace, throughwhich is passed a suitable, carbon-rich gas (e.g. methane, propane).

Liquid Carburising. The object is heated to a suitable temperature and thenimmersed in a hot, salt bath at 900C. The salts are usually based onsodium cyanide and the process is often called ‘cyanide hardening’. Themetal is quenched in water (not oil-which would react unfavourably with thesalts).

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4.4.2 NITRIDING

This process involves the absorption of nitrogen (instead of carbon) into thesurface of the steel. Suitable "Nitralloy" steels are necessary for this process andthey usually contain 1% Aluminium, 1.5% Chromium and 0.2% Molybdenum.

A special furnace is used and ammonia gas is circulated through it. The furnacetemperature of 500C converts the ammonia into a nitrogen-rich gas and formshard 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 isnecessary to achieve the final hardness, and no finish machining is required afternitriding. This, relatively low-temperature process, results in negligible distortionand is much cleaner than the carbon methods. Aircraft piston engine cylinderbarrels are particularly suitable for nitriding, as are some crankshaft bearingsurfaces and the stems of some aero-engine induction and exhaust valves.

Nitrided surfaces must be protected against pitting corrosion, usually (as withengine gears and shafts) by keeping the surface oiled.

Note: If certain surfaces of a component are not to be case-hardened, it isnecessary to protect them during the carburising or nitriding processes, tolocally 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 HARDENINGUnlike carburising and nitriding, flame and induction hardening do not add a

hardening agent into the surface of a basically softer material. Instead, they aremerely techniques for hardening the surface of material by a `local heat-treatment'.

Steels suitable for these processes already contain sufficient carbon (or otherelements) 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 theheated surface is then immediately quenched by water jets. The flame orinduction 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 hardsurfaces 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 STEELAs discussed earlier, iron has few practical uses in its pure state. Adding smallamounts of other materials to molten iron, however, dramatically changes itsproperties. Some of the more common alloying elements include carbon, sulphur,silicon, phosphorus, nickel and chromium (also referred to as chrome).

4.6 CARBONCarbon is the most common alloying element found in steel. When mixed withiron, compounds of iron carbide form and it is the carbon in steel that allows it tobe heat-treated to obtain varying degrees of hardness, strength and toughness.The greater the carbon content, then the more receptive the steel becomes toheat-treatment and, while its strength and hardness increases, its malleability andweldability decreases.

4.6.1 LOW-CARBON STEEL

Low-carbon steels contain between 0.1% and 0.3 % carbon and are classifiedas SAE 1010 to SAE 1030 steels. They are used in such items as locking wireand cable bushings and, in sheet form, they are used for low-loadapplications. Low-carbon steels weld easily but do not accept heat-treatmentvery well.

4.6.2 MEDIUM-CARBON STEEL

These steels contain between 0.3% and 0.7 % carbon. The increased carbonassists in heat-treatment while still retaining reasonable ductility. Medium-carbon steels are used for machining or forging and where surface hardness isrequired.

4.6.3 HIGH-CARBON STEEL

The carbon content of these steels, ranges between 0.5% and 1.5 % and thismakes them very hard. High-carbon steels are primarily used in springs, filesand in most cutting tools.

4.7 SULPHURSulphur causes steel to be brittle when rolled or forged and so it must beremoved during the refining process. If it proves impossible to remove all of thesulphur, then manganese, which is harmless to the steel can be added to themetal (to form manganese sulphide),. The manganese also improves forging bymaking the steel less brittle during the forming processes.

4.8 SILICONWhen silicon is alloyed with steel, it acts as a hardener and, used in smallquantities, it also improves ductility.

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4.9 PHOSPHORUSPhosphorus raises the yield strength of steel and improves a low-carbon steel’sresistance to atmospheric corrosion. The steel tends to be brittle when cold, sono more than 0.05 % phosphorus is normally used in steel production.

4.10 NICKELNickel is used extensively for alloying with steel as follows:

In the range of 1% - 5% there is a marked improvement in strength (andhardness) without lowering ductility. This high-strength, tough steel iswidely used for highly stressed parts.

At about 25% nickel, the steel becomes highly corrosion-resistant, heat-resistant and non-magnetic.

At 36% nickel, a unique steel (known under its trade name as ‘Invar’) iscreated. This has the lowest coefficient of expansion of any metal (1/20ththat of steel) and is excellent for master gauges and instruments.

Because of the effect of such amounts of nickel on the expansion propertiesof steel, a range of nickel-steels can be purpose-made, to trim thecoefficient of expansion to specific needs. These alloys are used inthermostats, spark plug electrodes etc.

4.10.1 NICKEL ALLOYS

When the amount of nickel present is predominant, then the material becomesknown 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 area family of alloys, containing 50% - 80% nickel, with the balance being mainlychromium (chrome) with some titanium and aluminium.

Nimonic alloys are used in hot air control ducting, for gas turbine enginecombustion chambers and turbine blades because of their extremely lowcoefficient of expansion at elevated temperatures.

Other ranges of nickel-based alloys come under the trade names of Inconel andHastelloy, 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 bothcorrosion and chemical attack, is tough, ductile, reasonably strong (equivalent tomild steel) and is non-magnetic. It is used in many marine applications, forsurgical apparatus and for aircraft rivets. Normally Monel does not respond toheat treatment but, when alloyed with a small amount of aluminium (2% - 4%), itcan 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 yieldstrength. By slowing the rate of hardening during heat-treatment, the depth ofhardening can be increased and the steel’s grain structure made finer. SAE 2330steel, containing 3.0 % nickel and 0.3 % carbon, is used in the manufacture ofbolts, nuts, rod ends and pins.

4.11 CHROMIUM (CHROME)When small amounts of chrome are added to steel, the strength and hardnessincreases, but there is some loss of ductility.

1.5% chrome, in a high-carbon (1%) steel, results in a very hard material which isused 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 hard-wearing 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. Oneparticular stainless steel is designated ‘18/8 Stainless’, which containsapproximately 18% chrome and 8% nickel. These stainless steels are usedextensively in engine parts, particularly for hot applications and for exhaust areaswhere their corrosion resistance is vital.

4.11.1 NICKEL-CHROME STEEL AND ITS ALLOYSThis term is used when the amount of nickel present is greater than the chromecontent. A wide range of such steels exists, but the low nickel-chrome alloys aresuitable for through-hardening or case-hardening. The nickel content is around3%-5% and the chrome ranges from 0.5%-1.5%. Crankshafts and connectingrods 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 heaterelements.

By adding both metals, in appropriate percentages, steel, which is suitable forhigh-strength structural applications, is produced. Nickel-chromium steels areused for forged and machined parts requiring high strength, ductility, shock-resistance and toughness.

4.12 COBALTCobalt is often included in High-Speed Steel (HSS) in addition to chrome,

vanadium, molybdenum, and tungsten (to improve still further the ability to cut athigh working temperatures). Cobalt is included in high-strength, permanentmagnets, in some of the nimonic alloys used for high-temperature components ingas turbine engines and cobalt is also found in a range of temperature-resistantalloys called ‘Stellite’ (used in piston engine valves and for cutting tools)

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4.13 VANADIUMWhen added to steel, vanadium improves the strength without loss of ductility, butalso greatly improves its toughness and its resistance to fatigue. Because of theimproved tensile and elastic properties, Valve (and many types of other) Springs,usually include vanadium. Small amounts of vanadium are included in certainnickel-chrome steels and good quality engineering tools.

Vanadium, when combined with chromium, produces a strong, tough, ductilesteel-alloy. Amounts of up to 0.2 % vanadium improve grain structure, ultimatetensile strength and toughness. Ball bearings are also made from chrome-vanadium steel.

4.14 MANGANESEWhen small amounts of manganese are added to steel (up to 1.5%) the result isa steel which is strong and hard (similar to nickel-chrome steel). Such steel isoften used for shafts and axles

11%-14% manganese steel has very unusual properties and is extremely useful.When this material is heated to approximately 1000C and water-quenched, itsstructure becomes austenitic and, although it is only moderately hard, anyattempt to cut it, or abrade it, results in the local formation of hard martensite andit thus becomes highly resistant to cutting or abrasion. Because of this peculiarproperty, it is used extensively for rock drills, stone crushers, and railway lines atjunctions etc.

Small amounts of manganese are used in steel production and in welding rodssince it acts as a purifying agent by reducing oxidation.

4.15 MOLYBDENUMOne of the most widely used alloying elements for aircraft structural steel ismolybdenum. It reduces the grain size of steel, which increases its impact-strength and elastic limit. Other advantages are an increase in wear-resistanceand high fatigue-resistance, which is the reason why molybdenum-steels arefound in structural members and engine parts.

4.16 CHROME AND MOLYBDENUMChrome-molybdenum steel is, probably, the most commonly used alloy steel inthe aircraft industry. Its SAE4130 designation denotes an alloy of1 0%molybdenum and 0.3 % carbon. It machines well and is easily welded by gas orelectric arc methods, as well as responding well to heat-treatment. Its use inaircraft construction includes landing gear, engine mountings and many enginecomponents.

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4.17 TUNGSTENTungsten has an extremely high melting point and adds this characteristic to thesteel with which it is alloyed. Tungsten steels retain their hardness at elevatedtemperatures, and are typically used for contact-breaker contacts (in magnetos),and also for high-speed cutting tools.

4.18 MARAGING STEELSConventional 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 combatthese shortcomings, maraging steels were developed. These steels are over50% stronger than normal high tensile steels and yet are very tough and easy tomachine. These properties are achieved by the almost total elimination of carbonand by alloying with nickel, cobalt and molybdenum in such a way that it can beprecipitation hardened.Maraging steels can only be used for special, high-stressed applications (due tocost, which is about three times that of conventional alloy steels). They are usedfor 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 non-ferrous 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 thiscourse, 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 thecommercially ‘pure’ state for engineering purposes - usually in the form ofsheets, tubes, wires or as thin coatings on other metals.

Cadmium, chromium, nickel, tin and zinc are also often used to provide protectivecoatings on other metals in order to retard the effects of corrosion.

Precious metals, such as gold, platinum and silver have been used for specialwork in high-grade electrical instruments, aircraft windshields and, of course,space vehicles.

Mercury (quicksilver) - the only metal to remain liquid at room temperature - maybe found in certain types of barometers, discharge lamps, small, electrical circuitbreakers, pressure gauges and vacuum pumps (it can also be found in thedetonators of some explosive devices).

In a similar manner to steels, it has been discovered that tremendous advantagesare to be gained by alloying non-ferrous metals with each other and, indeed, withother (ferrous) metals and elements.

Aluminium, copper, magnesium and titanium alloys are among the more commonnon-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 townof Les Baux, in France, where it was first found) . It is a soft, weak, ductile andmalleable metal. Aluminium is approximately one third the weight of steel and hasapproximately one third the stiffness of steel.

While its strength may be improved by cold working, it remains a low-strengthmaterial. Aluminium is highly corrosion-resistant, due to the rapid formation of athin, but very dense oxide surface film, which limits further corrosion and it is anexcellent 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 apatina 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 theability to be beaten, cast, drawn, forged, pressed, rolled or spun into manydifferent (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 beused in electrical systems (in bus bars, bonding, electrical motors, wiring etc.),though neither copper, nor its alloys, find much use as structural materials in theconstruction of aircraft.

5.1.3 PURE MAGNESIUM

WARNING;- WATER MUST NOT BE USED TO EXTINGUISH MAGNESIUMFIRES.

Two thirds the weight of aluminium (with a relative density of 1.74), no metal canbe cut, drilled, filed or shaped so easily as magnesium - provided that certainprecautions are taken to prevent it over-heating.

Magnesium burns readily, especially in small particles and dust. Great care mustbe taken when filing and grinding this metal and, if a fire should occur, it must beextinguished with dry sand or an appropriate powder extinguisher but WATERMUST NOT BE USED.

Magnesium is obtained primarily from electrolysis of seawater or brine from deepwells. In its pure state it lacks sufficient strength and characteristics for use as astructural metal. It can, however, be alloyed with a range of other elements togreatly improve its strength. These elements include aluminium, manganese,thorium, zirconium, and zinc.

5.1.4 PURE TITANIUM

WARNING:- TITANIUM FIRES MUST BE EXTINGUISHED WITH THECORRECT EXTINGUISHANT(DRY ASBESTOS WOOL AND CHALKPOWDER) AND NOT WATER.

Pure titanium at approximately 56% the weight of stainless steel, has almost thesame strength as iron. It is highly resistance to corrosion, non-magnetic and isreadily shaped by all of the methods, which relate to steel. Titanium is also softand ductile.

Care should be taken when working with titanium. Titanium fires usually startthrough high-speed rubbing. The low thermal conductivity of titanium prevents the

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rapid dissipation of heat, which progressively builds up locally, until ignition finallyoccurs. Accumulations of small particles of swarf and dust are a possible fire riskand all such accumulations should be avoided.

5.2 ALUMINIUM ALLOYSBecause pure aluminium lacks sufficient strength to be used for aircraft

construction and, to achieve medium/high-strength properties, aluminium must bealloyed with other elements. The most common alloying elements in the wroughtaluminium alloys are copper, manganese, magnesium and zinc. A commonelement used when casting aluminium is silicon.

Aluminium alloys may be designated as being either heat-treatable or as non-heat-treatable, though both types can be strengthened and hardened through

work-hardening (or strain-hardening). This process requires mechanically workingan 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, arereferred to as Light Alloys, while the remainder are termed Heavy Alloys.

5.3 IDENTIFICATION OF ELEMENTS IN ALUMINIUM ALLOYSVarious national Standards Institutions have evolved their individual systems foridentifying the many variants of aluminium alloys (in a similar manner to thatshown with SAE Steels).

While it would be impossible (and unsafe) to attempt to memorise them all, thesenotes provide examples of the American system of identifying aluminium (oraluminum) alloys.

American aluminium alloys are classified by a code, which refers to the elementthat makes up the major percentage of the alloy

As previously stated, the elements most commonly used for alloying withaluminium are copper, manganese, silicon, magnesium, and zinc.

Table 4 shows a four-digit number, which identifies aluminium, either in itscommercially ‘pure’, or in its alloyed state. The first digit of the designating coderepresents the major alloying element, while the second digit of the codeindicates a specific alloy modification, such as controls over impurities.

The last two numbers of the 1xxx group indicate the hundredths of 1% above the99% of pure aluminium. For example, if 75 were the last two digits, the metalwould be 99.75%pure.

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The 2xxx to 8xxx groups use the last two digits to identify the different alloys inthe group.

Table 4American System of Identifying Alloying Elements with Aluminium

Code Major AlloyingElement

1xxx aluminium2xxx copper3xxx manganese4xxx silicon5xxx magnesium6xxx magnesium & silicon7xxx zinc8xxx other elements

In the 1xxx group, commercially ‘pure’ aluminium (over 99% pure) is good forcorrosion resistance, has good electrical and thermal conduction properties, iseasy to work but is not very strong.

The 2xxx group uses copper as its major alloying element. The major benefit ofcopper is a large increase in strength, although if the alloy is not correctly heat-treated, intergranular corrosion can occur between the aluminium and coppergrains within the metal. These are probably the commonest aluminium alloysused in aircraft construction.

The 3xxx group has manganese as its major alloying agent and it is not possibleto heat-treat.

The 4xxx series utilises silicon as its major element. This lowers its melting pointand improves its welding and brazing capabilities.

The 5xxx group has magnesium as the main alloying element. This is good forwelding and corrosion resistance although, if exposed to high temperature or coldworking, it can corrode quite badly.

The 6xxx group has silicon and magnesium added to the aluminium. This makesthe alloy heat-treatable and with good forming and corrosion resistanceproperties.

The 7xxx alloys are made harder and stronger by the addition of zinc. These aredifficult to bend and are more often used where flat plates are required.

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5.4 CLAD MATERIALSThough 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 thinlayer of pure aluminium hot-rolled onto the surfaces. These are then known asclad materials with commercial names such as Alclad, and Pureclad.

Alclad is a ‘pure’ aluminium coating that is rolled onto the surface of an aluminiumalloy, which may, then, be heat-treated. The thickness of the coating isapproximately 5% of the material thickness on each side. For example, if analclad 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 aluminiumalloy. If, however, the cladding is penetrated, corrosive agents can attack thealloy under the cladding. For this reason, sheet metal should be protected fromscratches and abrasions. In addition to providing a starting point for corrosion,abrasions can create potential ‘stress raisers’ (points from which cracking caninitiate).

5.5 HEAT-TREATMENT OF ALUMINIUM ALLOYSWARNING:- 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 ANDWEAR ADEQUATE PROTECTIVE CLOTHING (APRONS, FACEMASKS, GOGGLES AND GLOVES) WHERE NECESSARY ANDENSURE THE CORRECT FIRE-FIGHTING APPLIANCES AREAVAILABLE.

Heat-treatment is a series of operations involving the heating and subsequentcooling 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 andmore ductile for working into a required shape (bending etc.). One treatmentcannot give all of these properties. Some treatments are achieved at the expenseof 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 timeand 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 propertiesare 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 inwhich the aluminium alloy is placed.

The salt bath employs molten mineral salts (water would evaporate long beforethe required temperatures were reached. The salts (usually nitrate of soda orsimilar) are solid at room temperature, but become liquid when they areelectrically heated. Gradual heating of the bath is necessary to avoid spattering orspitting. The aluminium alloy (pre-dried, also to avoid spattering) can then besubmerged within the heated liquid. Another precaution when using a salt bath isto avoid any adjacent flames or sparks, because the salts are inflammable.

Accurate thermostatic control is vital, as narrow tolerances on temperatures arespecified (typically plus or minus 5ºC).

Quench tanks must be sited nearby the furnace or salt bath, to avoid delaybetween removing from the heating source and quenching. Most quench tankscontain cold water but hot water is sometimes specified (especially for heavysections e.g. large forgings). Limits are also stipulated for the permissible periodbetween heating and quenching which is known as the lag-time (typically 10seconds max.). If these lag-times are exceeded, material properties or corrosionresistance may be adversely affected. If the cooling rate, during quenching, is tooslow this may also affect the corrosion resistance.

Thorough washing of the material is essential after salt bath heat-treatment toremove any salt residue.

There is no limit to the number of times that heat-treatment may be carried out onnormal aluminium/copper alloys but, if the material is clad with pure aluminium,for corrosion resistance (Alclad), then a maximum of three treatments isimposed.

This is to limit the migration of copper, from the alloyed material, into the purealuminium cladding, which would significantly reduce its corrosion resistance.

5.5.1 SOLUTION TREATMENT

Solution treatment is sometimes called ‘re-crystallisation H.T’. This operationserves to distribute the copper uniformly throughout the aluminium (i.e. to createa solid solution). The heating may be achieved (as previously stated) in an oven

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or, more commonly (to obtain better overall heating), in a bath of special, moltensalts. However, although the aluminium can accommodate 5% or so of copper insolid solution at high temperature, this condition is unstable at lower temperaturesand, after the alloy has cooled to room temperature, most of the copper slowlycomes out of solution and separates into local `islands' of copper aluminide.

By cooling the alloyed metals very quickly (quenching), the copper becomestrapped '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 reachmaximum values after several days (or weeks in some instances). Because of thetime lapse involved, this gradual hardening is termed ‘age-hardening’. Althoughcopper may be the major alloying element (in the ‘2000 series’ alloys) otherelements, including magnesium and manganese can also be present.

Although the aluminium/copper alloys are the most common age-hardened, high-strength 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 alloyis used in heavy loaded applications such as Main Spars, Landing Gear andMainplane Attachment brackets etc..

5.5.3 ANNEALING

Annealing, as with steel, serves to soften the aluminium alloy, to enable it to beworked without cracking. Even in this condition, ageing will gradually occur and

24 hours is the normal limit for working after annealing, although this can beextended if the material is stored under refrigerated conditions to slow the ageingprocess. A temperature of -5ºC will provide approximately 2 days’ delay while oneof -20ºC will provide approximately 1 week’s delay in the age-hardening processThe maximum for refrigeration is approximately 150 hours at -20°C.

Typical annealing procedure may be achieved by raising the temperature of thealloy to between 340°C and 410C. The alloy is then cooled slowly at about 10Cper hour (rates will differ with each particular alloy), until it reaches a pre-determined temperature. At this point it is allowed to cool naturally.

These, heat-treatable type, alloys must never be installed in an aircraft structurewhile in the annealed state, since material properties and corrosion resistance willbe 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 followingquenching although, with time, the metal gradually becomes harder and gainsstrength.

When the alloys are left at room temperature, after quenching, the hardeningprocess (natural ageing), and can take from several hours to several weeks. Analuminium/copper alloy, for example, is only at 90% strength within 30 minutes ofquench, but is at maximum strength after four or five days.

We have already discussed how the natural ageing process can be drasticallyretarded (allowing the metal to be kept in a soft condition until required for use),by storing the alloys at sub-zero temperatures(refrigeration) for prescribedperiods of time.

Alternatively, following quenching, by re-heating the metal to a lower temperaturethan that employed for the solution treatment and allowing it to ‘soak’ at that heatfor a period of time, the ageing process (and, thus, the hardening of the alloy) canbe accelerated. This process is referred to as artificial ageing or precipitationtreatment.

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5.6 IDENTIFICATION OF HEAT-TREATED ALUMINIUM ALLOYS

Aluminium alloys that have been subjected to heat-treatment are usuallyidentified by markings that indicate the heat-treatments involved. Three typicalidentification systems are those of the British Standards Institute(BS), theMinistry of Supply (MoS), and the American systems as can be seen in Table 5.

Table 5IDENTIFICATION MARKINGS OF HEAT-TREATED ALUMINIUM ALLOYS

BS System Meaning

M As manufactured stateO Annealed state

OD Annealed and lightly drawnT Solution-treated, no precipitation requiredW Solution-treated, can be precipitatedWP Solution-treated and precipitation treated

MoS System Meaning

A Annealed stateN Solution-treated, no precipitation requiredW Solution-treated, and requires precipitationWP Solution-treated and precipitation treated

AmericanSystem Meaning

T3 Solution-treated and cold workedT4 Solution-treated only (naturally aged)T6 Solution-treated and artificially agedT8 Solution-treated, cold worked and artificially agedT9 Solution-treated, artificially aged and cold worked

An example of one of these marking systems would be an alloy with thedesignation 2024-T4, which indicates an aluminium/copper alloy that has beensolution-treated only, and then naturally aged

Apart from these systems, many other exist world-wide, but the British systemsare, broadly, confined to three basic ones for light alloys.

British Standards for general engineering use BS 1470 -1475. In this seriesthe prefix N is used to denote non-heat-treatable aluminium alloys andprefix 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 commonlyabbreviated to L71, L72, L 73 etc.

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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, originallymanufactured by the Duren Aluminium Company (Germany), but it tendsto be used as a generic name for similar alloys, regardless of source ofmanufacture.

D.T.D. Specifications: - these are material identification numbers issued bythe Directorate of Technical Development (a Ministry Department) forspecialised applications. i.e. when widespread use is not anticipated.

If such a material finally becomes commonly used, a British Standardsspecification is compiled and issued.

5.7 MARKING OF ALUMINIUM ALLOY SHEETSSheet material, for aerospace use, is marked ‘all over’ with the specificationidentification, 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 similarlooking metals.

Some sheets may also have alternate lines of red numbers/letters, which indicatethat heat-treatment is needed before assembly. These red numbers/letters thendisappear when the necessary heat-treatment is done.

It is imperative that only the correct specifications and thicknesses ofmaterials are used in the construction of aircraft structures.

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5.8 CAST ALUMINIUM ALLOYSThese are not used extensively on airframes mainly due to their lack of strength,poor fatigue characteristics and lack of elasticity when compared to the wroughtaluminium alloys.

The lack of elasticity is particularly relevant, as the very nature of an airframestructure requires the ability to flex considerably without cracking.

Although their use is obviously limited on airframes, cast aluminium alloys areused extensively on engines, where there is a need to produce complex coredshapes such as crankcases, drive casings, cylinder heads etc. No other methodthan casting would be viable for such items. The stresses can be kept to amodest 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, temperature-resistant alloys exist. One of the most common in the category is RR 58(sometimes known as `Y' Alloy), which is an age-hardening material containingapproximately 2½% copper, l½% magnesium, 1½% nickel, and l% iron. Aderivative of this material was also used (in wrought form) for the skin of thesupersonic 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 linearexpansion, so is often included in piston castings.

5.9 MAGNESIUM ALLOYS

WARNING;- WATER MUST NOT BE USED TO EXTINGUISH MAGNESIUMALLOY FIRES.

Magnesium alloys are used for castings and, in their wrought form, are availableas sheet, bar, tubing and extrusions. They are among the lightest metals havingsufficient strength and suitable working characteristics for use in aircraftstructures.

There are some serious disadvantages to using magnesium alloys in aircraftconstruction. These include a high susceptibility to corrosion and cracking.

The corrosion problem is minimised by treating the surface of the metal withchemicals, which form an oxide film, to prevent oxygen reaching the metal.

Another way of minimising corrosion is to use hardware such as rivets, nuts, boltsand screws that are made from compatible materials.

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The cracking problem contributes to the difficulty in shaping magnesium alloysand, thus, limits its use. One method used to overcome the tendency for crackingis to form the metal whilst it is hot.

Magnesium alloys can also be solution heat-treated, which will improve theirtensile strength, ductility and resistance to shock. To improve their hardness andyield strength they can also be precipitation heat-treated after the solution heat-treatment.

5.10 COPPER ALLOYSOf those (Heavy) alloys that use copper as a base; brasses, and various bronzesare the primary types used on aircraft.

Brasses may contain zinc and small amounts of aluminium, iron, lead and otherelements 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 tinand 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 valveseats. 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 stillreferred to as‘bronzes’. High-Tensile Brass, for instance, because of itsmanganese content is called ‘Manganese Bronze’, while Phosphor and Siliconbronzes also contain practically no tin. Wrought aluminium bronzes are almost asstrong as medium-carbon steel while cast aluminium bronzes are found inbearings 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 ithas been heat-treated, beryllium bronze is very strong (300-400 Brinell) and isused for diaphragms, precision bearings and bushings, ball bearing cages andspring washers.

Leaded Bronze is found in the bearings of some aero engines. The very highpressures (and speeds) tend to squeeze the lubricant out of normal journalbearings, so the addition of lead acts as a sort of lubricant in the event of the oilfilm breaking down.

Solder is a general term frequently used for joining metals together. The principaltypes are ‘soft solder’ (which is a mainly lead-tin alloy), and ‘hard solder’ which isan alloy of copper, silver and zinc.

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5.11 TITANIUM ALLOYS

WARNING:- TITANIUM ALLOY FIRES MUST BE EXTINGUISHED WITH THECORRECT EXTINGUISHANT(DRY ASBESTOS WOOL AND CHALKPOWDER) AND NOT WATER.

Titanium alloys, apart from being light and strong, also have excellent corrosionresistance, particularly in a salt-laden atmosphere. To prevent reaction withoxygen and nitrogen, in its pure form, titanium is treated with chlorine gas and areducing 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 heat-treatments. Commercially pure titanium is ‘non-heat-treatable’ (It can beannealed, but its strength/hardness cannot be improved by heat-treatment.).

When suitably alloyed, titanium based materials are heat-treatable. Thestrengthening is immediate i.e. it is not an age-hardening material.

Titanium alloys are used extensively in aerospace gas turbines, but their use islimited on subsonic civil airframes to fasteners, and high temperature areas suchas engine bays, heat shields, hot zone bulkheads, air ducts etc.

In appearances titanium is similar to 18/8 stainless steel. Two practical methodsof identification apart from weight are:

spark test - a light touch of a grinding wheel will produce a brilliant whitetrace, ending in a brilliant white burst.

moisten the titanium and draw a line on a piece of glass - this will leave adark line similar to a pencil mark.

5.12 WORKING WITH TITANIUM AND TITANIUM ALLOYSCAUTION: DO NOT STAND ON, OR PUSH AGAINST, THIN-WALLED

TITANIUM STRUCTURES. LOCAL HARDENING WILL CREATE STRESSRAISERS WHICH COULD LEAD TO STRUCTURAL FAILURE.

DO NOT ALLOW HOT OIL TO DRIP ONTO TITANIUM. IT WLL CAUSEHYDROGEN EMBRITTLEMENT. THERE IS NO REPAIR PROCEDEURE FORTHIS TYPE OF CONTAMINATION OF TITANIUM.

AN INCORRECT CLEANING AGENT CAN ALSO CAUSE HYDROGENEMBRITTLEMENT IN TITANIUM COMPONENTS

Titanium materials are, generally, not susceptible to normal corrosion attack, butit has been established that stress corrosion cracking can take place in somewelded structures which are exposed to trichloroethylene and other chlorinated

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hydro-carbons (the alloys most affected in practise being the titanium-aluminium-tin family).

Titanium may also show evidence of deterioration in the presence of salt depositsor metal impurities, especially at high temperatures. It is, therefore prohibited touse steel wool, iron scrapers or steel brushes for the cleaning of, or for theremoval of corrosion from, titanium components

If titanium surfaces need cleaning, then hand-polishing, or the use of soft bristlefibre brushes, with aluminium oxide compound or a mild abrasive may bepermissible. Use only the recommended procedures outlined in the relevantMaintenance or Overhaul Manual

When it is necessary to machine a welded titanium structure, or doubt existsregarding the use of cutting fluids with a particular titanium alloy, the materialmanufacturer should be consulted

5.12.1 DRILLING TITANIUM

Rigidity is essential when drilling titanium and titanium alloys so that thin-wallstructures must always have a backing support.

Centre drilling should always be used, instead of centre punching, as the localwork-hardening caused by centre punching will cause difficulty in starting the drilland 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 helixangle 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 90 º

or ‘double-angled’ point is better. Drills must be precision ground and special caremust be taken to ensure that the drill tip is completely central, as any off-set ofthe 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 frictionaltroubles. High quality soluble oils, used in the diluted form recommended by themanufacturers, or chlorinated or sulphured oils, should be used in generousquantities 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 holesbelow 6 mm.(0.25 inch) diameter, and of 0.1 to 0.2 mm (0.005 to 0.010 inch) perrevolution for larger holes is recommended. Positive power feed must be

employed whenever possible.

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6 METHODS USED IN SHAPING METALS

There are four basic methods of converting raw material into the requiredmanufactured shape whilst also achieving the desired material structure. Theyare casting, deformation, machining, and various forms of fabrication (i.e. thejoining together of smaller pieces or particles of material to form a larger object).Welding, adhesive bonding, mechanical fasteners or even powder metallurgycome 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 flowplastically in the solid state without deterioration of their properties. Processessuch as these, result in a minimum of material waste.Machining processes provide excellent precision, but the process generates alarge amount of waste material. Fabrication techniques enable complex shapesto be constructed from simpler particles or units.

6.1 CASTINGThis involves the pouring of molten material into a shaped mould and allowing itto solidify to that shape. It is an ancient process, which enables complex shapesto be produced in a wide range of materials in a single-step operation. Castcomponents can range in size from the small teeth of a zip, to large casings ofseveral metres in diameter. Ocean-going ships’ propellers, up to 10 metres indiameter, 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, byfar, 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 themetal is poured).

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Although sand-casting is simple in principle, there are many vital aspects of thetechnique, which are necessary to produce good castings. The sand, forexample, must have:

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

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

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

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

Adequate hardness (the hardness may be checked by the resistance toindentation 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 mustalso meet certain standards, some of which are:

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

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

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

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

The incorporation of vents, where necessary, to permit the escape of gas andsteam 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 notrequire elaborate equipment and is economical for small batches. It is alsosuitable for most metals. The major shortcomings are that the process is not veryrapid, it is not particularly accurate (due to lack of sand rigidity) and it is notsuitable for thin-wall sections.

6.1.3 TYPICAL CASTING DEFECTS

Casting defects vary to some extent, depending on the casting process used, butthe 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 beforesolidification)

Cold Shuts (when local areas of metal are not molecularly joined, due tosolidification 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 amixture 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 castingprocess is then the same as for normal sand-casting. The advantages of shell-moulding 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. Theprocess is used for the manufacture of hollow cylinders (e.g. cylinder liners),bronze or white metal bearings etc. The rotation can result in acceleration forcesof up to 60g and this produces high-quality, dense castings, since all of the slagmigrates to the bore (due to it being of lower density than the metal) and it canthen be machined out.

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6.1.6 DIE-CASTING

This process uses a permanent metal mould, which results in more accurate, andbetter finished, castings than those produced in sand. Die-casting, can be sub-divided into ‘gravity’ or ‘pressure’ processes, depending on how the metal is fedinto the mould.

Gravity Die-Casting - sometimes known as ‘Permanent-Mould Casting’.This casting process is virtually identical to sand-casting except that themould (die) is metal. A wide range of metals can be cast and hollowcastings are possible if a sand core is used. Fine grain structures areproduced, due to the more rapid rate of cooling, compared to that achievedin 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 innon-ferrous materials (aluminium, magnesium, zinc, copper and theiralloys), because steels have too-high a melting temperature for the metaldies to accommodate. The dies are, usually, made from hard, tool-steelsand are water cooled. This process can achieve excellent detail, superfinish, 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. Hollowcastings 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), butit only became of great industrial importance in the 1950's, when gas turbinemanufacturing began to increase. The process was ideally suited to theproduction of complex-shaped nozzle guide vanes and turbine blades which,often, contained tortuous inner passages, very thin sections and had to be cast inexotic 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 ofinvestment material (a ceramic slurry or paste), usually by dipping a numberof times.

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

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

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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 precisioncastings with a dimensional accuracy of less than 0.1 mm. Surface finish is alsoexcellent, but the major advantage, that the process offers, is the ability toproduce accurate, complex shapes which would be impossible by machining.

6.2 FORGINGThis is a squeezing/hammering technique, which is intended to achieve largedeformation/shaping of the material. The process is usually carried out hot (i.e.above the re-crystallisation temperature), so that these large deformations can beattained without being accompanied by any massive, residual stresses.

Sometimes a cold forging operation may be necessary but, in this instance, thematerial will be harder, stronger and pre-stressed (i.e. still containing unrelievedinternal stresses).

Forging ranges from the simplest form of the hand operations, conducted by theblacksmith, to the massive, mechanical, powered rams, used for very largeforgings. The forging hammer will often have a relatively low strike rate, butsometimes high-speed, pneumatic hammers are used for High-Energy-RateForming.

Forging not only shapes the metal, but also reduces grain size and produces adirectional control of grain flow. Both of these are desirable features for manyengineering applications, particularly for highly-stressed components, such ascrankshafts and especially if they are subject to a mechanical fatigueenvironment.

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6.2.1 DROP-STAMPING

Drop-stamping, or drop-forging (refer to Fig. 16), involves the use of shaped diesand a heavy drop-hammer, which usually falls under gravity. The piece ofmaterial, to be forged, is placed between the top and bottom dies and the drop-hammer is allowed to fall the necessary number of times for the contact faces ofthe dies to come together. ‘Flash gutters’ are provided, to accommodate theexcess 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 ProcessFig. 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 processtends to affect the whole structure of the component, whereas some forgingprocesses, using multi- (but light) blows will, mainly, affect the material closest tothe surface.

6.2.3 UPSETTING

Upsetting is, sometimes, called ‘Heading’ and usually involves locally heating ofthe end or ends of the material, immediately prior to forging. Poppet valves areformed 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 areformed in this way.

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6.3 ROLLINGRolling can be carried out hot or cold.

When done hot ,it is capable of achieving major re-forming/re-shaping, and slabscan be reduced to plate or sheet while bars of circular or rectangular crosssection can also be produced. Hot rolling can also produce structural shapessuch 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 thehot scale from steel prior to cold rolling. The rolling process would also be used toproduce the clad (and unclad) sheets of aluminium alloys.

6.4 DRAWINGDrawing 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 ashaped, hardened die. A ductile material is essential.

6.5 DEEP DRAWING/PRESSINGThis process uses a ram, to deform a piece of sheet metal into a recessed dieand is usually done hot.

6.6 PRESSINGPressing 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 apowered ram. Pressing is usually done hot (except for the soft, ductile materials).

6.7 STRETCH-FORMINGThis 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 overthe former (by moving the clamps or the former) and the material is stretchedbeyond its elastic limit so that permanent deformation occurs.

This process is convenient for small batches of material (and is particularlyfinancially 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 FORMINGIn 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 criticalmatching and alignment problems of conventional pressing, When used for small

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batches (e.g. aircraft production), low-cost, easy to machine, materials can beused for the forming block.

Rubber-Bag forming (Hydro-forming) uses the same principle, but incorporates aflexible diaphragm and hydraulic pressure in place of the rubber pad.

6.9 EXTRUDINGThe extrusion process, forces hot metal through a shaped die, to producecircular, rectangular, tubular, angular, half-round sections etc.

In some respects, the process is similar to drawing, but extruding forces metalfrom a heated billet, through hardened dies by compression, whereas, in drawing,it is achieved by tension. Malleability is, therefore, an essential material propertyfor the extrusion process.

Extruding is normally restricted to aluminium alloys and copper alloys, whereextrusion 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 softand malleable materials (e.g. aluminium). The shaped component is formed, byforcing a punch onto a ‘blank’ of material within a shallow recess. The extrudedshape results from the metal being forced to escape through the small gap,between the punch and the recess.

6.10 SINTERINGSintering; 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 veryhigh 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 ornitrogen) to prevent oxidation. Under the high pressures used, a metallurgicalbond occurs (diffusion bonding), between the particles of powder. The sinteredend-product is, typically, around 10%-20% porous and can then be impregnatedwith graphite (or high melting-point grease), to provide excellent, self-lubricatingproperties 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 carrythe 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 incorporatingtungsten carbide particles within a cobalt matrix.

Hot, Isostatic-Pressing, uses a similar technique to sintering, but uses highertemperature and very much higher pressures to produce zero porosity. Thetechnique is sometimes used to heal micro-porosity in super-critical castings.)

6.11 SPINNINGSpinning 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. Thenecessary force to deform the sheet metal is generated by a long tool, which islevered about a suitably positioned fulcrum.

For thin gauge, soft metals, the tool can be manipulated by hand, while, forthicker 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 MILLINGChemical milling is, sometimes, referred to as chemical etching. It is a purelychemical process, not electro-chemical.

Although simple in principle, chemical milling offers a method of producingcomplex patterns and lightweight parts and is used for incorporating integral ribsand stiffeners in sheet metal. Tapered sections can also be easily formed - theunwanted 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 slowprocess, 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 athin layer of plastic, which can be brushed or sprayed on.

Although the chemically etched surface is not very rough, a drop in fatiguestrength does result and, in critical applications, restoration of fatigue strength isdesirable. A light, peening operation, using glass beads or steel shot, achievesthis.

6.13 ELECTRO-CHEMICAL MACHININGUsing electrolysis and, by making the workpiece the anode of the dc electricalcircuit, an electrolyte is pumped rapidly (under pressure) through the gapbetween 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 isprogressively removed from the workpiece, until the desired shape is achieved

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The process is ideal for metals, which are difficult to machine by conventionalmethods, 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 goodconductor (copper or brass) and it must resist corrosion, because the electrolyteis 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 theenergy (and heat) of electrical sparks, which travel from the electrically negativetool electrode, through a dielectric fluid, and explosively strike the electricallypositive workpiece.

The intense heat of the strike, causes local particles of metal to instantaneouslyvaporise, without a molten metal phase (a process known as ‘sublimation’),though, away from the actual centre of the explosion, molten fragments of metalare washed away, with the vapour, by the dielectric fluid.

A suitable fluid (usually kerosene) is fed, under pressure, between the electrodeand the workpiece, to maintain a uniform electrical resistance. The spark rate isaround 10,000 per second and the gap between the tool and the workpiece iscritical 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 aredifficult to machine conventionally, but it also excels in its ability to produce high-aspect 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 mmdiameter 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. Thewire, 0.05 mm - 0.25 mm in diameter (0.002 in - 0.010 in), is positioned by, andfed over, two pulleys and resembles a simple band-saw operation. The workpieceis mounted on a table, which can be moved in two axes and, when the table iscomputer controlled, the wire-cutting process can cut accurate, complex shapesin metals (e.g. dovetails, fir-trees etc.) which are difficult to machine withconventional 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 theadvances, until relatively recently, have been confined to tooling improvementswhich have permitted higher material removal rates. The early, high-carbon steeltools, 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 ofautomatic tool changes and of doing difficult types of machining without the needfor transferring work to a different machine and re-setting up. In this way a muchmore versatile machine tool has evolved. However, the biggest single machiningadvance in modern times (especially with regard to aircraft manufacture) hasbeen the introduction of Numerically Controlled (NC) machines. NC milling, inparticular, 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 onthe Binary System (or a variant) which is common to most electronic computingdevices. The primary advantage of NC machining is the ability to accuratelycontrol the spindle, the tool or the workpiece movements in three directions (x, yand z axes) independently or simultaneously. NC machines are capable ofproducing compound shapes and contours, and are especially suited to the taskof generating integral spars, ribs, and stiffeners in slabs or forgings.

NC machines usually incorporate a feed-back system, which ‘tells’ the control unithow much actual movement is made, analysis is then done and finalcompensation eliminates any error (i.e. the motion ceases when the input andfeed-back signals agree). Electrical control of the machine servo-motors, cancontrol movements as small as 0.0005 mm (0.00002 in).

CNC machines (i.e. Computer Numerically Control) differ from NC machines onlyin that the electronic control unit on the CNC machine is more sophisticated in

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that it is adaptable to a wide variety of software and can accommodate a diverserange of programs. Although the capital cost of NC/CNC machines is high, thefollowing advantages make such machines technically desirable andeconomically viable, where super-light, complex, high-tech, manufacture isconcerned:

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 FORMINGSome Titanium alloys, when heated, become extremely ductile and can

plastically deformed without necking occurring This superplasticity can beexploited in the forming process (refer to Fig. 17), when an inert gas is used toblow the material into the required shape

Superplastic Forming ProcessFig. 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 termcould be applied to a wide range of engineering materials. These would includenot only the metallic alloys, but also the most earliest of all composite materialsused by man, - wood (the tough, fibrous, xylem, or water-conducting tissue, ofshrubs and trees, which contains lignin and cellulose). Brick, concrete, and glassare among the many other materials, which could be considered as composites.

In the aerospace industry, the term ‘composite’ is used when referring tomaterials, which, in turn, are a combination of fibrous and synthetic resinmaterials 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 materialswill be found in use on modern aircraft.

7.1 PLASTICSThe word plastic comes from the Greek plastikos - to mould, and plasticity (aswas discussed in The Properties of Metals) is the ability to retain a deformationafter the load, producing it, has been removed. Plastics are particularly useful forapplications, which involve relatively low-stress levels, where lightness isimportant, and where low electrical or thermal conductivity is required.

The earliest plastic materials (before the synthetics) were those made from thesap, 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 syntheticplastic material (used as an inexpensive substitute for ivory), from the cellulose ofplants (and called it Celluloid), while the chemist, L H Baekeland (in 1909)developed the first entirely synthetic plastic material (Bakelite), from phenol-formaldehyde. Bakelite is hard and fairly brittle. It is often used with a suitablefiller material (mica, or wood flour) and is widely used for various electricalmouldings 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, thatcan be cast, extruded or moulded into various shapes or drawn out into filamentsto be used as fibres.

While the two major groups of plastics are the Thermoplastic and Thermosettingcompositions, the manufacture of synthetic rubbers (called Elastomers) is alsoconsidered to be part of the plastics industry.

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7.1.1 THERMOPLASTIC MATERIALS

Thermoplastic materials, in their normal state, are hard but become soft andpliable when heated(the Greek word thermo- heat). When softened,thermoplastic materials can be moulded and shaped, and they retain their newshape when cooled. Unless their heat limit is exceeded, this process can berepeated many times without damaging the material.

Two types of transparent thermoplastic materials are used for aircraft windshieldsand side windows, and are usually referred to as cellulose acetate and acrylic.

Older aircraft used cellulose acetate plastic because of its transparency and light-weight. A disadvantage of cellulose acetate is its tendency to shrink and discolourwith time, which has led to it being phased out almost completely.

Cellulose acetate can be identified by its slight yellowish tint (especially whenaged), and by the fact that a scrap of it will burn with a sputtering flame and giveoff 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) andPlexiglass(USA). It is stiffer than cellulose acetate, more transparent andpractically colourless. Acrylic burns with a clear flame and gives off a fairlypleasant odour. Acetone, if applied, will cause white marks but will leave thematerial as hard as it previously was.

7.1.1.1 Use of Thermoplastics

Thermoplastics are, normally, used where there are no unusual temperaturechanges and the majority of all plastics production is thermoplastics, whichinclude:

Acetate - widely used for tool handles, and electrical goods.

Poly-Ethylene - commonly known as polythene. Its uses include flexibletubing, cable insulation and packaging.

Poly-Propylene - stronger, harder and more rigid than polythene. Used forsuch items as high-pressure air piping.

Poly-Vinyl-Chloride - commonly known as PVC. Varying degrees ofrigidity/flexibility are achievable by varying the amount of plasticiser used.Rigid, moulded sections or piping can be produced and also flexible electriccable insulation

Polystyrene - can be produced in rigid form, but is more familiar in theexpanded form, when it is useful for thermal insulation, buoyancy or shock-resistant 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 lighttransmission properties and are also resistant to splintering. There is atendency for some fine craze-cracking to develop if exposed for longperiods to ultra violet light. These transparent plastics may be solid orlaminated. When laminated two or more layers are bonded together with aclear adhesive and, in this form, they are more shatter-resistant and areideally suited to pressurised aircraft windows.

An even stronger and more shatterproof transparent plastic can be achievedby stretching the acrylic in both directions before final shaping. Theseimproved properties, result from the stretching operation causing apreferential alignment of the long-chain molecules. Extreme care should betaken when handling acrylics, as they are they are easily scratched. Theacrylics are supplied with a paper or rubberised film, which should not beremoved, until required for use. If dirty, they should be cleaned with coldwater or soapy water. Care should also be taken when using solvents in thevicinity of acrylics. Some solvents, or their vapours, may cause crazing ofthe material. , Reference to the appropriate Manuals or manufacturers’specification sheets are essential.

Poly-Carbonates - these have similar uses to the acrylics (Perspex etc) butare 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 andversatile material. It is strong, tough and also has low friction properties. Itcan be used as a fibre or produced as a moulding. Popular uses includetextiles, furnishings, ropes, tyre reinforcement, bushes, pulleys, gears, andlightweight mouldings such as brackets, handles etc.

Poly-Tetra-Fluoro-Ethylene - commonly known as ‘PTFE’, it is similar tonylon in appearance but is denser, whiter and much more expensive. It hasa wax-like surface and this characteristic results in very low frictionproperties, which make it suitable for bushes and gears. It also has a hightemperature capability (over 300ºC) and is extensively used as a non-stickcoating e.g. Teflon. PTFE tape is often used as a thread sealant for oxygenpipe threads, and as backing rings for hydraulic seals

7.1.2 THERMOSETTING MATERIALS

Thermosetting materials (also called Thermosets) will, initially, soften whenheated, but will remain soft for only a short time and will set (and harden) if theheat continues to be applied.

The process of Thermosets becoming hard, when heated, is called ‘curing’ andcuring 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 plasticbecomes hard and cannot be re-softened by heating.

Thermosets are, thus, chosen where a plastic component will be exposed torelatively high temperatures, as some of them can tolerate temperatures inexcess of 250C before beginning to char.

Note: Thermosetting materials are generally stronger, have a lower ductility andlower impact properties than the Thermoplastics.

7.1.3 RESINS

Natural resins are obtained from the exudations from certain trees and otherplants and as clear, translucent, yellow (amber), brown, solid, or semisolidagents, they are used in inks, lacquers, linoleum, varnishes and, of course,plastics.

While the words plastics and resins are often used synonymously, they are, infact, quite different, in that plastics refers to the material in the finished itemswhile 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 employedwith them, to assist in the moulding characteristics, or to enhance the propertiesof 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 micro-balloons, cotton and glass flock and aerosil (fumed silica).

Reinforcing agents, plasticizers, stabilisers, colorants, flame-retardants, smokesuppressants and processing aids, such as lubricants and coupling agents, areamong the other additives used with resins.

Resins have little strength in themselves and are generally used to impregnatelinen, 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 hightemperature, and have high strength without causing wear to the control cables.

When layers of paper are impregnated with a thermosetting resin such as phenol-formaldehyde or urea-formaldehyde, they can be moulded into flat sheets orother shapes. Once hardened, the material makes an exceptional electricalinsulator and can be found in use as terminal strips and printed circuit boards.

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7.1.3.1 Polyester Resin

Polyester resin can be extruded into fine filaments and woven into fabric (likenylon) 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 lackrigidity so, to convert glass fibre into a useful structural material, it is impregnatedwith polyester resin and moulded into a desired form.

Polyesters cure by chemical action, and, so, differ from materials, which cure bythe evaporation of an oil or solvent. As polyester is thick and unmanageable, astyrene 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 solidmass, so inhibitors are added to delay this curing process and to improve shelflife.

A catalyst then has to be used, when the inhibitors are no longer wanted and thecuring process is to be started and an accelerator will appreciably shorten thecuring 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 thecatalyst and accelerator generates heat within the resin. This(exothermicreaction 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 viscousand cause it to ‘run’ (particularly if it is on a vertical surface). To overcome thisproblem, a thixotropic agent is added to the resin after mixing, to increase itsviscosity. The increased viscosity allows the resin to remain in place no matterwhere it may be used.

7.1.3.3 Epoxy Resin

Another type of resin that can be used in place of polyester in laminatedstructures is epoxy resin. Epoxy resin has a low percentage of shrinkage, highstrength 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 orcuring agent without recourse to heating.

There is also a difference in the mixing ratios between polyester and epoxyresins. For polyester resin, the ratio is 64:1, resin to catalyst whilst, for epoxyresin, 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, syntheticmaterials (polymers) which have considerable elastic properties.

Because they may also be moulded into shapes, which they retain, they qualify tobe included in the category of plastics. Elastomers will tolerate repeatedelongation and return to their original size and shape, in a similar way to naturalrubber

Some of the more common elastomers, to be found in the aerospace industryinclude:

Buna ‘N’ - also known as Nitrile. A synthetic rubber, made (initially inGermany) by the polymerisation of butadeine and sodium (hence BuNa), ithas 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 incontact with metal, and is, therefore, particularly suited to ‘moving-seal’applications.

Buna - ‘S’ relatively cheap material, also with a performance similar tonatural rubber. It is often used for tyres and tubes, but its poor resistance tofuels/oils/cleaning fluids makes it unsuitable for seals.

Fluoro-Elastomers - these have exceptional high-temperature propertiesand can be used at 250ºC. They are also solvent-resistant and are mainlyused for high-temperature seals. A common name for these materials isViton. These materials are expensive.

Neoprene - has very good tensile properties and excellent elastic recoveryqualities. It is also solvent-resistant and, therefore, has a wide range ofapplications as fuel and hydraulic seals and gaskets. However, because ofits special elastic recovery properties, it is also ideally suited to diaphragmsand hydraulic seals.

Poly-Sulphide Rubber - although it possesses relatively poor physicalproperties, 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 stressedseals and hoses, which come into contact with fuels or oils. This compoundis 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 pottingof electrical circuits, because of its ability to retain its rubbery state, even atlow temperatures.

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7.2 PRIMARY ADVANTAGES OF PLASTICSPlastics are being used on an ever-increasing scale and are frequently replacingsome of the more conventional materials such as metals, wood and naturalrubbers. Plastics have properties, which make them a popular choice overconventional aircraft materials. Some of the more important characteristics ofplastics, which help to explain their popularity, are:

Lightness - most plastics have specific gravities of 1.1 to 1.6 whereas themore common engineering materials, such as aluminium and steel, havevalues of 2.7 and 7.8 respectively.

Corrosion Resistance - plastics will tolerate hostile corrosion environmentsand many of them resist acid attack.

Low Thermal Conductivity - this property makes many plastics ideal forthermal insulators.

Electrical Resistance - plastics are used in enormous quantities for electricalinsulation applications.

Formability - many plastics are easily formed into the finished product, bycasting moulding or extrusion, often in a single operation.

Surface Finish - excellent surface finishes can be achieved in the basicforming operation, so finishing operations are not necessary.

Relatively Low Cost - because, although some of the materials may not beparticularly cheap, the lack of machining necessary and the high productionrates possible, keeps the costs down.

Light Transmission - some plastics are naturally clear, whilst other areopaque. These characteristics, consequently, provide the possibility for arange of light-transmission properties. Optical properties can also beachieved with some plastics.

Vibration Damping - many plastics are naturally resistant to fatigue and,because of the high value of internal damping present, resonances will tendto be of relatively low amplitude.

7.3 PRIMARY DISADVANTAGES OF PLASTICSAlthough plastics are extremely useful materials, some shortcomings inevitablyexist, particularly when compared to some metals. Plastics major deficienciesare:

Lack of Strength - most plastics are much weaker than metals and mildsteel 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 arecomparable.

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Low Stiffness - plastics have a very inferior value of Young’s Moduluscompared with the common metals.

Low Impact Strength - many plastics have poor impact strength, but thereare a few exceptions, such as with certain polycarbonates.

Poor Dimensional Stability - mainly due to high values of thermal coefficientof expansion.

Poor High-Temperature Capability - metals are generally capable ofretaining reasonable strength at much higher temperatures than the

plastics. The long-term maximum operating temperature, for the betterplastics, is not usually above 250ºC. High-temperature metals can operatefor long periods well in excess of 800ºC.

Moisture Absorption - many types of plastic absorb moisture, which canresult in a significant loss of strength in a humid environment.

Ultra Violet Light - some plastics deteriorate when exposed to UV light forlong periods. Increased brittleness and loss of strength can occur.

7.4 PLASTIC MANUFACTURING PROCESSESThe most common manufacturing methods are:

Casting - the molten material is simply poured into a mould and allowed toset.

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 mouldor around a former.

Sandwich-Construction - plastic facings have, sandwiched between them, ahoneycomb or foam core. Very stiff, but light, structures are achieved bythis 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, theplastic is sucked into contact with the shaped die (a method often used tomanufacture aircraft interior trim).

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7.5 COMPOSITE MATERIALSAs previously stated with Plastics, the main reason for utilising composite

materials, in aerospace structures, is to reduce weight, which has a direct benefitin lowering operating costs. Composites also provide further benefits in theirability to be easily formed, comparatively lower production costs, resistance tocorrosion and reduced maintenance costs.

The principal types of composite materials are those involving fibrous elementswhich may be used as strands, or be woven into fine ‘tapes’ and ‘cloths’ (orcoarser ‘mats’), held in a suitable resin matrix and formed into the requiredshapes

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 ableto transmit light over long distances.

The ultimate tensile strength of undamaged, very small diameter glass fibres isextremely high, although the strength is reduced significantly if the fibres areslightly damaged.

In its structural use it is often merely referred to as glass fibre or fibreglass, whenglass fibres (in various forms) are bonded together by appropriate resins.

When moulded with resin, the resulting composite is, also, of considerably lowerstrength but, nevertheless, good GFRP structures are stronger than mild steeland, on a simple strength-for-weight basis, can be comparable to high tensilesteel if the fibre form and lay-up is near optimum. It is however, considerably lessstiff than steel or even aluminium.

A graphic example of GFRP flexibility is the enormous deflection, which takesplace in the pole during a pole vault. As the glass fibres are about a hundredtimes stronger than the resin, it is obviously necessary to get as much fibrepacked into the moulding as possible.

Non-structural items may be made from, or include, a percentage of choppedstrand mat, (i.e. glass fibres in a random, non- woven state) but, whereconsiderable strength is required, uni-directional glass cloth is used.

To provide all round strength, sheets of uni-directional cloth can be layed up at90º to each other, in a similar manner to the grain in plywood. Sometimes suchsheets are used as facings for an internal honeycomb of plastic-impregnatedpaper, 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 impregnatedwith resin and partially cured (‘Pre-preg’), in which case it is necessary to keepthe material in refrigerated storage. Resin curing is usually done at elevatedtemperatures (120C - 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 non-electrical 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 notsuitable 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 ofglass fibre, used in high-temperature applications. They can be used attemperatures up to 1650C and are suited for use around engine and exhaustsystems. Ceramic fibres are heavy (and expensive) and are only used where noother materials are suitable.

7.5.2 CARBON FIBRE REINFORCED PLASTIC (CFRP)CFRP (also referred to as ‘Graphite’) is a composite material, which was primarilydeveloped to retain (or improve upon) the high strength-to-weight ratiocharacteristics exhibited by GFRP, but with very much greater stiffness values.

Carbon fibres are very stiff and, when formed into a composite, the Young'sModulus (‘E’) value can be higher than steel. CFRP is not only six times stifferthan GFRP but is also over 50% stronger. It also has twice the strength of high-strength aluminium alloy and three times the stiffness.

Carbon fibres are typically less than 0.01 mm (0.0004 in) in diameter and areproduced by subjecting a fine thread of a suitable nylon-type plastic to a very hightemperature (to decompose the polymer), and driving off all of the elements withthe exception of carbon. The carbon thread is then stretched, at white heat(2000C-3000ºC), to develop strength. Unfortunately, the process is complex andvery costly.

Nevertheless, where the high cost can be justified, CFRP can offer considerableweight savings over conventional materials. CFRP components are generallymade from ‘Pre-preg’ sheet (fibres impregnated with resin and a hardener, whichonly require heat and pressure to cure). Some specialist items are made by a

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laborious, but ideal, process called ‘Filament Winding’, in which a carbon fibrestring is wound over a former in the shape of the workpiece whilst bonded withresin.

Because of CFRP's high stiffness modulus, it is also used extensively to stiffenGFRP or aluminium alloy structures.

A material known as Carbon-Carbon (where the resin is also graphitised), is usedfor the rotors and stators on brake units. It offers a significant weight saving, aswell as high efficiency, due to the fact that it dissipates the heat generated veryquickly.

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 manyaircraft.

The use of composites, in the manufacture of helicopter rotor blades, has led tosignificant increases in their life and, in some cases, they may have an unlimitedlife span (subject to damage). The modern blade is highly complex and may becomprised 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 arewell known for their superior toughness, strength-to-weight characteristics andheat-resistance. Tyres, reinforced with aramid fibres are comparable to thosereinforced 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-resistancemakes it useful in areas, which are liable to be struck by debris, as experienced inareas around engine reverse-thrust buckets. Kevlar is used to manufacture bullet-proof 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 propertiesdepend on the direction of the fibres. Random direction fibres would result in amuch 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 verylow, 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 ofthe finished component (which can be affected by careless handling, variations incleanliness 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.

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It has already been stated that composites usually have good internal dampingcharacteristics and are less prone to vibration resonances. Where high strength,combined with stiffness is required, then a CFRP is used but, when lesser levelsof stiffness are necessary, then GFRP or AFRP are used.

Composites have very low elongation properties and toughness. Aluminium alloyhas a typical elongation-to-fracture value of 11%, whereas composites rangefrom 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 therange 220C-250ºC.

Some composites, such as carbon fibre in a carbon matrix, have very highpermissible operating temperatures (around 3000ºC), and are used for high-energy braking applications and as thermal barriers for space vehicles).

Boron, Tungsten, Silicon Carbide and Quartz may also be used to provide fibresfor high-temperature composites

7.5.5 LAMINATED, SANDWICH AND MONOLITHIC STRUCTURES

Laminated plastics consist of layers of synthetic resin-impregnated fibres (orother, coated, fillers), which are bonded together (usually heated and underpressure), to form a single laminate or sheet of composite material. Plasticlaminates 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 ofkeeping a surface clean is required

To provide a light-weight structure, which possesses strength and rigidity, one ofseveral structural materials, is sandwiched between two laminated composites.

The sandwiched material (the core) may be made of a solid material, such aswood, or a series of thin corrugations of a material, which are joined and placedend-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 balsawood, which has been cut across the grain and sandwiched between two layersof reinforced resin (or a metal). This construction makes an extremely light, yetstrong material, which can be used as floor panels, wall panels and, occasionally,aircraft skins.

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The cellular core, used for laminated honeycomb material, may be made fromresin-impregnated paper, or from one of the many fibre cloths. The core is formedor 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 forradomes of all kinds.

Metal honeycomb cores (made from light alloy or stainless steel), are alsosandwiched between two face sheets of fibre-reinforced resins. On otheroccasions the metal honeycombs may be found sandwiched between sheets oflight alloy, stainless steel or titanium. This type of core is referred to as ‘metal-faced honeycomb’ and is used where abrasion- and heat-resistance is importantor when sound-absorption qualities are desired.

In monolithic structures, angle sections (‘Top Hat’, ‘U’, ‘I’ and ‘Z’), frames ribs andstringers are fashioned from similar materials to the outer layers of the sandwichstructure, then covered with the appropriate surface ‘skin’, before the stronger,metallic spars and hinges are attached, Such a structure can save manykilograms (or pounds) in the weight of the flying control surfaces (or the finstructure) of a large aircraft.

7.6 NON-METALLIC COMPONENTSIn addition to the non-metallic materials, used in the aircraft structure, non-metallic materials are used in many aircraft components and systems. Many ofthese materials require specialist knowledge and understanding, during aircraftmaintenance.

7.6.1 SEALSSeals or packing rings (refer to Fig. 18) serve to retain fluids and gases, withintheir 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 andmaterials.

The most common materials, from which seals are manufactured, are naturalrubber, synthetic rubber and Teflon (trade name for polytetrafluoroethane orPTFE). O-ring seals effectively seal in both directions of movement. They areused to prevent both internal and external leakage, and are the most commonlyused seals in aviation

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Examples of Seals and Packing RingsFig. 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 outor extruded. These back-up rings are usually made from Teflon, which does notdeteriorate with age, is unaffected by system fluids and vapours and toleratestemperatures well in excess of those found in high-pressure hydraulic systems.

O-rings are available in many different materials and sizes (both diameter andthickness). They are supplied in individual, hermetically-sealed, envelopes with allthe necessary information marked on the packaging. This system has generallyreplaced the previously used, colour-coding of seals, which had severelimitations.

For applications (such as in actuators) that subject a seal to pressure from twosides, two back-up rings can be used but, when the pressure is from one sideonly, a single back-up ring is adequate.

Other seals, commonly found are V-ring and U-ring seals. The V-ring has anopen ‘V’ facing the pressure and is located by the use of a male and femaleadapter. 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) areencountered. 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, associatedwith metals and also have good fatigue characteristics, they do, however, requireregular 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 therelevant chapter (51-57 for Airframe topics, 61-61 for Propellers) of the AircraftMaintenance Manual (AMM). The inspection methods to be used will be found inthe Non-destructive Testing Manual (NTM) and the approved repair procedureswill be outlined in the Structural Repair Manual (SRM).

Repairs in unexpected areas, or damage, which is not covered in the SRM, willnecessitate the request of specific repair drawings from the aircraft manufacturer.

8.1 CAUSES OF DAMAGEIf a sharp object strikes a thermosetting plastic, the plastic is liable to crack andshatter, like glass, with straight sharp edges. The reason for this is that, once acrack 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 surfacefinish applied to it. An important point about this type of damage is that there islittle loss of strength in the overall material, in addition to the absence of theshattering 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-strikedamage can also require extensive repairs. Damage to composite structures mayresult 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 andexcessive ‘g’ loading.

Lightning strikes and static discharge.

Chafing against internal fittings such as pipes and cables.

8.2 TYPES OF DAMAGEThe types of damage, which may affect fibre-reinforced structures are:

Cracks which may simply affect the outer lamination or may penetratethrough the skin.

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Fibre reinforced plastics however, apart from being much stronger thannormal plastics, have different failure modes. Each strand of fibre acts as atrap, to stop cracks travelling through the plastic(refer to Fig.19). Atravelling crack quickly reaches a fibre, which is difficult to break so, instead,the crack travels along the fibre. Eventually the crack reaches another fibreand is deflected again. This process continues until the failure is divided intomany small cracks, which will not have propagated far from the initialdamage.

Fibre

(a) (b) (c)Crack travelling Crack travelling Cracks around fibre

towards fibre along fibre

Crack Propagation within a CompositeFig. 19

Delamination - which involves separation of the fibreglass layers and mayaffect single or multiple layers.

Debonding - when honeycomb sandwich structures are damaged, theeffect usually entails separation of the honeycomb from the skin. Thereason for this is that the bonding of the skin to the honeycomb walls isalong very fine lines, and this bond is fairly easily broken.

Once there is separation, the strength of the whole structure is reduced bya significant amount. Greater damage can be due to the crushing of thehoneycomb core itself, which may require extensive repair or evenreplacement of the complete component.

Blisters - which usually indicate a breakdown in the bond within the outerlaminations and may be caused by moisture penetration through a smallhole, 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 causedby lightning strikes or by static discharge.

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8.3 INSPECTION METHODSAreas 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 internalflaw) and surface defects such as delamination and blistering. A lamp and amagnifying glass are useful in detecting cracked or broken fibres. A smallmicroscope or a x20 magnifier may be helpful in determining whether the fibres ina cracked surface are broken, or if the cracks affect only the resin.

Delamination may sometimes be found by visual inspection. If the area isexamined at an angle, with a bright light illuminating the surface, the delaminatedarea may appear to be a bubble, or an indentation in the surface. When viewedfrom the inside, a change of colour could indicate delamination because of achange in light reflection.

A visual inspection can also find several manufacturing defects such as resin-richor 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 orpercussion, test can be used. In some instances a properly designed miniaturehammer is used for the test while, in other procedures, a length of an appropriatehardwood, or a particular size of coin is employed to tap against the surface ofthe suspect area.

Variations in the tapping sound will provide clues as to the quality of the bond. Asharp solid sound indicates a good bond, whilst a dull thud indicates bondseparation. Care must be taken to make allowances for changes in materialthickness, fasteners and earlier repairs, all of which can give false indications.Whenever damage is found visually, then a percussion test should be donearound the area. In the majority of instances, if there is a hole, crack or otherdamage, 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-frequencyultrasonic beam through the structure and viewing the pattern of the resultingsound echo on an oscilloscope.

8.3.4 RADIOGRAPHY

Radiography can, sometimes, be used to detect cracks in the surface in additionto being able locate internal faults that cannot be visually detected. Radiographicprocedures may also be employed to detect water ingress within honeycomb corecells.

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8.4 ASSESSMENT OF DAMAGEOne of the greatest problems, caused by replacing aluminium alloy withcomposite structures (especially honeycomb sandwich), is the inspection fordamage. It is unfortunate that when a composite of any kind is struck, the majorityof the damage occurs internally and, often, there is little or no visible damageshowing at the surface

It is vital that ANY damage to a composite structure be thoroughly inspected, notonly for damage to its surfaces, but also (in a sandwich structure) for possibledamage to its core, which is usually softer than the skins. Damage that gives littleclue 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 maybe classified as negligible (or allowable), repairable by cover patch, repairable byinsertion or repairable by replacement.

These classifications may only be determined by reference to the appropriateaircraft SRM. Signs of secondary damage (i.e. damage occurring remote from theprimary damage) must not be overlooked. This is particularly important in thecase of impact damage where the shock may be transmitted through thestructure, to cause damage away from the point of impact. In some instancessecondary damage may be more serious than the primary damage.

Sometimes damage may be difficult to detect, due to the natural flexibility of thematerial 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, asit may indicate the presence of damage.

Where delamination is known, or suspected to exist, the area surrounding thevisible damage should be checked to determine the extent of the damage andintegrity of the laminations.

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9 BASIC COMPOSITE REPAIRS

WARNING: THE CHEMICALS, USED IN LAMINATING RESINS ANDCLEANING AGENTS, ARE HAZARDOUS SUBSTANCES AND EXTREMECARE IS CALLED FOR WHEN HANDLING THEM.

MOST RESINS ARE AN IRRITANT TO THE SKIN. MANY PEOPLE AREALLERGIC TO THE RESIN, AND REPEATED SKIN CONTACT CANCAUSE SERIOUS DAMAGE. IF SYMPTOMS OF AN ALLERGY APPEARWHEN THE RESIN IS USED, FURTHER CONTACT SHOULD BE AVOIDEDAND THE SYMPTOMS SHOULD SLOWLY FADE AWAY.

DIRECT SKIN CONTACT WITH THE RESIN SHOULD BE AVOIDED, ANDRUBBER OR PLASTIC GLOVES WORN WHEN THERE IS A POSSIBILITYOF THE HANDS BECOMING CONTAMINATED.

THE RESINS AND SOLVENTS, USED WITH SYNTHETIC FIBRES, AREALL POISONOUS. EVERY PRECAUTION SHOULD BE TAKEN TO KEEPTHEM 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, MUCHGLASS AND RESIN DUST WILL BE PRODUCED AND A RESPIRATORYMASK SHOULD BE WORN FOR PROTECTION. THE SAME DUST ISLIKELY TO CAUSE AN IRRITABLE SKIN RASH TO DEVELOP ON THEFOREARMS, ESPECIALLY WHEN GLASS FIBRE IS BEING HAND-SANDED.

BEFORE WASHING HANDS AND ARMS, AFTER WORKING WITH GFRP,IT IS ADVISABLE TO RINSE THEM IN COLD WATER. THE ARMSSHOULD BE WASHED IN SOAPY WATER AND THE OPERATOR SHOULDAVOID SCRATCHING, ESPECIALLY WHILE DUST IS LYING ON THESKIN.

Before commencing repairs on any composite material (whether it be a simple‘fibreglass’ skin or a complex honeycomb sandwich), the complete area of thedamage must be carefully surveyed. This must be done in accordance with theAMM, to ensure that ALL damage is discovered and assessed.

Any subsequent repair will depend on the type of damage, the extent of thatdamage and the importance (significance), to the safety of the aircraft, of thematerial being repaired. The AMM will provide either a repair scheme orcomponent replacement information.

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The strength of a glass fibre repair is dependent on the strength of the bond tothe original structure. Since the repair receives its working loads through thisbond, 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 duringthe repair process

Uniform Stress: - once again correct procedures during repair will ensurethat local stress concentrations are minimised.

9.1 REPAIR OF A SIMPLE COMPOSITE PANELIf a heavy object has been dropped onto a glass-fibre-reinforced epoxy panel, thedamage could consist of a small hole surrounded by damaged composite (refer toFig. 20).

Point of Impact

Damaged Area

Front of Panel

Rear of PanelDamage to Composite

Fig. 20

The damaged material is removed first (refer to Fig. 21), bearing in mind that thedamage may be small on the front, but may extend some distance at the rearside of the panel.

Undamaged Panel

Damaged Area RemovedFig. 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, eachon top of the previous piece. In this way, the hole is filled with successive layersand 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 CompositeFig. 22

9.2 REPAIR OF A SANDWICH PANELThese repairs are considered to be more difficult than composite panels, due totheir complexity, and require skilled personnel. In this example, the assumption isthat the dropped tool has broken the skin and damaged the core. As previouslystated, the first task is to remove the damaged material, usually with a router(refer to Fig. 23).

Plug of DamagedMaterial 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 bondedinto the hole (refer to Fig. 24).

HoneycombPlug

Core

Damage PluggedFig. 24

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

Completed Repair of Damaged AreaFig. 25

Note: The above examples are only an outline of the full repairs that may bedone, during aircraft maintenance.

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9.3 GLASS FIBRE REINFORCED COMPOSITE REPAIRSThese course notes deal exclusively with repairs to glass fibre reinforcedcomposites, including honeycomb-cored structures.

To summarise, glass fibre composites have two basic constituents, namely theglass fibre and the surrounding plastic matrix. The glass fibres reinforce theplastic matrix and carry most of the load. The matrix gives the composite itsrigidity and protects the fibres from attack by moisture or chemicals.

Glass fibres are generally woven into a fabric, which gives a regular orientation tothe fibres and allows them to be handled more easily.

To produce a glass fibre laminate, successive layers of the fabric are placed intoposition and impregnated with resin. The liquid resin solidifies within a few hoursand after post curing at elevated temperatures, forms a strong matrix around thefibres.

Using this technique, intricate shapes can easily be formed with the load carryingfilaments orientated in the best possible manner. It is also possible to reinforcethe laminate locally and to mould in load bearing fittings etc. into the laminate.

9.4 TYPES OF GLASS REINFORCEMENTAfter production of the basic glass fibres, they are collected together to form acollection of continuous, parallel fibres known as a roving.

Glass fibre cloth is made by weaving rovings together. Depending on thecloseness 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 andin 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 weftdirections and, as such, can take stresses in both directions. There are two maintypes 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 MATChopped 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 whichis made of woven fabric. As a result, it is of little use in aircraft construction.

9.4.4 RESINThe choice of resin for a particular application, is most important, because resinsare produced with the necessary properties to suit only certain requirements andare, 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 anexplosion may occur. The correct mixing procedure must be followed so thatthe 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 themixture begins to solidify (cure).

.Resin MixingIn any resin mix, the proportions are absolutely critical, since the cured strengthdepends on it. The proportions are normally specified by weight of the quantity ofresin required. An excess of hardener in the mixed resin is as damaging as adeficit. In both cases the cured resin will have an incomplete molecular structureand result in poor physical properties.

Scrupulous cleanliness is essential in the mixing process, which should beperformed 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 LIFEThe 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 ifthe temperature is too high, the resin will be comparatively thin and will drain outof the laminate before solidification occurs. Ambient temperature and humidityrequirements are specified by the resin manufacturer.

The length of time before a mix of activated resin begins to solidify is called ‘potlife’ and is dependent on the temperature and quantity of resin. Once the resinbecomes thick and stringy, the curing process is well on its way. Resin in thisstate should not be used ,since the cured strength properties will be seriouslydegraded.

To prevent waste, only sufficient resin should be mixed for the task in hand.

9.6 CURINGMost 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 hashardened, post-curing, at elevated temperature, is required for the resin to gainits full strength.

For repair purposes the heat is usually applied by means of an infra-red lamp orelectric heater. For components, which have been removed from the aircraft, anoven of suitable size may be used, to allow accurate control of temperature. If alarge enough oven is not available, then a hot-air ‘tent’ should be constructedaround the repair, and a thermometer used, to measure the average temperatureinside the tent.

Temperature may also be controlled by use of a temperature-indicating lacquer orpencil. These, when applied adjacent to a repair, will melt or change colour whena pre-determined temperature has been reached.

The times and temperatures, required to effect a cure, are specified in therelevant SRM. The maximum curing temperature must not be exceeded. A typicaltime and temperature would be 8 hours at 60ºC.

The use of pressure is normally specified for a repair whilst it is being cured. Thisassists 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 withchanges in climate and is only attacked by a few chemicals.

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9.7 GEL COATThe 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 ofthe exposed surface of the laminate. This prevents the outermost glass fibres ofthe laminate from becoming exposed to attack by moisture and sunlight. If the gelcoat 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 mayalso be used as paint and, after curing, it can polished to give a smooth, glossysurface.

9.8 STORAGE OF GFRP MATERIALSGFRP materials are expensive and, to ensure maximum shelf life, they should bestored in proper conditions.

9.8.1 STORING RESIN

Most laminating resins have a limited shelf life, which is specified by themanufacturer. In general, they should be stored in airtight tins at a cool

temperature (usually below 10ºC). The resin should be removed from storage atleast 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 whichtime it must be discarded. Resins, which have absorbed moisture, and becomecloudy, should normally be discarded, but they can sometimes be recovered byheating them to 120ºC, to evaporate the moisture. If the resin clears on cooling, itmay 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 airtightcontainers. Some hardeners may crystallise if they become cold. To liquefy thehardener it should be gently warmed and then allowed to cool at roomtemperature.

Note: The catalyst and accelerator, of a three-part laminating resin, should bestored separately to avoid inadvertent contact.

9.8.3 STORING FABRICS

Glass fabric should be stored in a warm, dry atmosphere, free from dust, oil orother contaminants. In order to preserve the fibre surface treatment it must no getdamp. 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 shouldbe stored in their original wrappings.

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9.9 PREPARATION FOR REPAIRWhen the damage has been assessed as repairable, preparatory steps may betaken which are common to most types of repair.

The gel coat should be removed by grinding, or by gently chiselling andpeeling it away, to determine whether the glass fibres are damaged. Signsof overstraining of the structure will show up as white cracks in thelaminations. If the rear of the structure is accessible, a strong light, shonethrough the laminates, will show up any damage (delamination or cracks)as a dark area. The affected area should be cut out and the damagetreated as a hole.

The damaged area should be cleaned and then cut back until soundmaterial is reached. No evidence of whitening or cracking must be allowedto remain.

Note: Before cutting out the damage, the area should be marked in someway, to determine its orientation for future reference (refer to Fig. 26).

Any control linkages, bearings etc. should be covered to keep out glass dustand surplus resin.

………………………………………………………………………………………………………

Sound Material

……

Repair Area…………………………………………………

……………………

* Orientation Mark

Orientation Marks on RepairFig. 26

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

It is possible to analyse a sample of material, removed from the damagedarea, 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 beseparated. The weight and direction of the fibre may now be determinedand related to the parent laminate by reference to the previously appliedorientation.

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Notes should be made to ensure that the repair will be to the samespecification as the original laminate (i.e. number, weight and direction ofeach layer). If the structure used a core material, the type and thicknessshould be noted. If the core is wood, the grain direction should be noted.

The patch edges may now be prepared according to the particular repairbeing followed (scarf or stepped)and any surface that will have fibre bondedto it must have a thorough preparation (see the following paragraph, entitled‘Surface Preparation’)

When preparing a chamfered (scarfed) edge, the sanding direction shouldbe towards the tip (refer to Fig. 27). The prepared edges should beexamined for any sign of delamination, which must be removed by furthersanding.

Sanding Direction

Direction of SandingFig. 27

Note: Some manufacturers specify that cut-outs should have radiusedcorners, while others permit square corners.

The inside of the structure should now be cleaned out and any loose piecesof 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 acetoneor another, approved fluid). Once cleaned, the area should not be touched withbare hands. All paint, gel coat etc. must be removed from the repair area. Thefollowing 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 glassand 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 isabraded.

Any dust must be removed with a clean cloth.

The newly exposed surface is thoroughly degreased, using a clean clothsaturated with the appropriate fluid (acetone).

The acetone must be allowed to evaporate from the surface. Careful use ofa hot air blower is recommended to drive off any traces of acetone that maybe trapped in the surface fibres.

Having cleaned the surface, the repair should commence as soon aspossible.

9.10 TECHNIQUES OF LAMINATING GLASS FIBREWhile the actual procedure will be detailed in the SRM, a typical list of thetechniques to be adopted would include:

Pre-shaped templates are used to cut out the required pieces of cloth for therepair.

The workshop temperature must be between 15ºC and 23ºC with a relativehumidity of not more than 65%.

The quantity of resin required should be estimated and mixed in the correctproportions 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 nopossibility of the container contaminating the contents (for this reason ‘unwaxedpaper cartons’ are recommended).

Note: If the resin is for structural repair work, a small sample (about 1cc) of mixedresin is cast in a container made from aluminium foil. The sample should belabelled and placed aside to cure for later inspection.

A coat of resin is brushed onto the prepared surface and the first layer of clothis placed on the resin. The cloth is stippled into the resin, ensuring that the clothweave 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 allowthe cloth to ‘wet out’ more quickly and help to prevent the cloth sticking on thebrush.

Note. Beware of using too much resin as this will result in a resin-rich and heavyrepair. 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 thecorrect area. This is done, by lifting the edge of the patch and removing theexcess with a sharp pair of scissors.

Each subsequent layer of cloth is then positioned and stippled into thepreceding layers (trimming as necessary) until the laminate is complete.

When laminating is complete, the repair must be allowed to cure without anyfurther disturbance.

9.11 PRE-WETTING GLASS FIBREThere are a few occasions (during aircraft structure repairs) when the use of pre-wetted cloth is expedient. The cloth is laminated on flat cellophane or plastic filmand as many as four layers may be laminated at once.

The pre-wetted cloth is then transferred to the job and stippled into place beforethe plastic film is then peeled off. During these occasions the following pointsmust be noted:

Care must be taken to ensure that the pre-wetted cloth produces a goodbond to the parent material.

The plastic backing film should be peeled off as the cloth is being laidbecause, with it in place, the laminations cannot assume a double curvatureor irregular shape.

It is important to ensure that no bubbles are trapped, though it is quitedifficult to detect bubbles in a multi-layer lamination.

The edges of each cloth layer must be staggered so that there is not anabrupt 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 ABSOLUTELYESSENTIAL WHEN WORKING WITH ADHESIVES AND SEALANTS.ALTHOUGH MANY OF THE ADHESIVES IN CURRENT USE ARESUPPLIED IN FILM FORM, SOME ARE LIQUIDS OR PASTES, FROMWHICH, TOXIC/FLAMMABLE VAPOURS ARE EMITTED, PRIOR TOCURING. MANY OF THE NECESSARY SURFACE PREPARATIONSOLVENTS ALSO GIVE OFF TOXIC/FLAMMABLE VAPOURS.

Adhesive bonding has been used on an ever-increasing scale and particularly inthe aerospace industry. Adhesives are used for constructional tasks varying fromaircraft fuselages, flight control surfaces, to propellers and helicopter rotor blades.

10.1 THE MECHANICS OF BONDINGThe actual adhesive bond may be achieved in two ways:

Mechanical: - here the adhesive penetrates into the surface and forms amechanical lock, by keying into the surface. It also forms re-entrants, wherethe adhesive penetrates behind parts of the structure, and becomes anintegral part of the component to be joined.

Chemical (Specific): - in this method of bonding, the adhesive is spread overthe 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 orpeel. Lap joints are the types more, generally favoured, as the strength of theadhesive bond is proportional to the area bonded,):

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Tensile.Where the two surfaces are pulled directly apa

Joint in Tension

Shear.Where the two surfaces tend to slide acrosseach other.

Joint in Shear

Cleavage. Where two edges are pulled apart.

Joint in Cleavage

Peel. Where one surface is stripped backfrom the other

Joint in Peel

Stresses on Bonded JointsFig. 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. Alsoproviding a smooth finish due to lack of rivet heads.

No local stress raisers, which are present with widely-pitched conventionalfasteners (Bolts, rivets etc.).

Can be used to join dissimilar materials and materials of awkward shapesand of different thickness, as rivetting and welding are not always possibleon very thin (or very thick) materials.

Although the strength per unit area, may be inferior to a mechanical orwelded joint, adhesive bonding takes place over a greater continuous areaand, therefore, gives comparable or increased strength, coupled withimproved stiffness.

Adhesive and sealants provide electrical insulation and prevent dissimilar-metal corrosion between different materials.

Leak-proof (fuel and gas) joints can be achieved.

The elastic properties of some adhesives, gives flexibility to the joint andmay 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 whereenvironmental 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 ADHESIVESThere exists, an enormous range of adhesives, and the correct type, for aparticular application will be specified in the relevant repair procedure.

Great care must be taken that only the correct type is used as, otherwise, acatastrophic failure may well occur, should an unsuitable adhesive be usedon a critical structure.

The two major groups of adhesives are: Flexible

Structural

10.2.1 FLEXIBLE ADHESIVESFlexible 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 beingepoxy or polyester)

10.2.2 STRUCTURAL ADHESIVES

Structural adhesives are primarily aimed at applications where high loads mustbe carried without excessive creep. They are, therefore, relatively rigid, butwithout being excessively hard or brittle

Note: Another group of adhesives is the two-polymer type, which has areasonably even balance of resin and elastomer, which results in a flexible, yetfairly 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 forcuring, whilst others can be cured by the addition of a catalyst or hardener.

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10.3 ADHESIVES IN USETo achieve optimum bonding, performance, and life in service, from adhesivesand sealants, it is absolutely crucial to follow carefully planned processes andprocedures and to pay the utmost attention to quality at every stage. In fact, themajor criticisms, levelled against the use of adhesives, are:

Absolute cleanliness at all stages is essential. Surface preparation of thecomponent is also crucial. To ensure consistent results on structuralcomponents, a purpose-built ‘clean room’ is required, in order to reducecontamination to a minimum.

Pressure and heat may be required. Sophisticated equipment is required toproduce pressure over the components in areas where adhesives areapplied. This will often entail vacuum bags, purpose-built ovens, orpressurised curing ovens (autoclaves).

Inspection of the bonded joint is difficult. Special inspection techniques andtest pieces are necessary to check the integrity of the bond. Prior topreparing 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 ifnecessary. This procedure is essential, to enable the final assembly ‘wet’lay-up to proceed without delay, and without the risk of generating swarf orof 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 contaminationuntil the bonding process is complete.

Surfaces to be bonded are normally thoroughly cleaned/degreased in asuitable solvent. This may be followed by a chemical etch or light blastingtreatment, followed by a water wash and subsequent drying.

10.3.2 FINAL ASSEMBLY

The adhesive is applied (usually within a specified time, otherwise re-processingmay be necessary), and the assembly suitably clamped, or put in a nylon vacuumbag, and heated in an autoclave. The curing process then takes place undercarefully controlled temperature and pressure conditions.

When cool, the component is inspected, visually for positioning and for asatisfactory spew line. The glue-line thickness is also checked, with a calibratedelectronic probe, and specimen test pieces are tested for shear and peelproperties.

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Following a satisfactory inspection, the component is finally given appropriatecorrosion protection (usually over-painting).

Note: After commencing the final (wet) lay-up, curing of the adhesive must becarried out within a specified time (usually 12 hours). If this period is exceeded bya few hours it is necessary to increase the temperature and pressure levelsduring 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. afull shift or day), it will be necessary to dismantle and not only re-commence thewet lay-up, but also to, possibly, repeat some of the preliminary surfacepreparation 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 COMPOUNDSCertain areas of all aircraft are sealed to withstand pressurisation, prevent fuel orfume leakage and to delay the onset of corrosion, by sealing against the weather.

Most sealant compounds, consist of two or more ingredients, that arecompounded to produce a desired combination of strength, flexibility andadherence. 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 applicationstraight from their packaging. The consistency of some of these compounds canbe altered to satisfy a particular application method. If, for example, thinning isrequired, then a thinner (recommended by the sealant manufacturer), is mixedwith the sealant.

10.4.2 TWO-PART SEALANTSTwo-part sealants are compounds requiring separate packaging, to preventcuring prior to application. The two parts are identified as the base sealingcompound and the accelerator. Two-part sealants are generally mixed, bycombining equal portions (by weight), of the base and accelerator compoundsand any deviation from the prescribed ratios will result in inferior sealing oradhesion.

Many common sealants/adhesives are produced in pre-measured kits, thatsimply require the mixing together of the whole quantities of the materialssupplied. These eliminate the need for balances and other weighing equipment.

The instructions must be followed but, in general, require the addition of theaccelerator to the base compound, followed by thorough mixing beforeapplication.

A working life is usually quoted, which applies after mixing, so the work must bethoroughly prepared prior to mixing.

Some materials may be kept, after mixing, for a limited time, by the use ofrefrigeration. 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. Forexample, at temperatures below 15C, curing is extremely slow. At temperaturesabove 21C, curing times are usually faster. For best results, a temperature ofaround 25C, 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 notexceed 50C at any time during the curing cycle. The heat can be applied, byusing 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 asheet 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) eachyear and, with care and good husbandry, this figure can be reduced. The morethat aircraft can be manufactured, operated and maintained with the short- andlong-term considerations of the effects of corrosion in mind, then the more thosemaintenance 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 aresubsequently formed) they are relatively stable. Corrosion is the tendency ofmetals to revert to the thermodynamically more stable, oxidised, state. Thisoccurs when they react with dry air to form metal oxides, or with acids and alkalisto form metallic salts. Some metals, such as gold and platinum, strongly resistcorrosion.

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 isattacked, suffers an Anodic change while the corrosive agent undergoes a

Cathodic change. The result is that material is lost from the Anode and gained bythe Cathode, forming an ionic bond.

11.1 CHEMICAL (OXIDATION) CORROSIONIn a strict chemical sense, oxidation occurs whenever a metal is converted to itsions. 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 combinationof 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 ofsome metals. With the exception of gold and platinum, all metals, in contact withair, form a very thin, visible oxide film.

Chemical corrosion can be caused by direct exposure, of the metal surface, tocaustic liquids or gaseous agents such as:

Spilled battery acids or battery fumes. Spilled acids are less of a problemnow that Nickel Cadmium batteries are in common use.

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Flux deposits from inadequately cleaned joints. Flux residues arehygroscopic (readily absorb moisture).

Entrapped caustic cleaning compounds. Caustic cleaning solutions shouldbe kept capped when not in use. Many corrosion-removal solutions are, infact, 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 furthercorrosive attack. The oxidation rate normally falls sharply as the film thicknessincreases (refer to Fig. 29), so that, at some time, there is virtually no furtherincrease in film thickness.

TemperatureConstant

OxideThickness

Time

Oxide Thickness over TimeFig. 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 mayreact chemically, or combine with, water to produce a film that is not imperviousto the passage of further oxygen through it. The oxide skin may also crack orflake 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 rateof oxidation of a metal (refer to Fig. 30). The actual curves are not as smooth asthose shown.

550C

525C

500C

450C

OxideThickness

Time

Effect of Temperature on OxidationFig. 30

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11.1.3 EFFECT OF ALLOYING

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

0% Al

Oxide + 3% AlThickness

+ 7% Al

Time

Effect of Alloying on OxidationFig. 31

The reason for this effect is that the oxide film, which forms, is rich in aluminiumoxide, and provides more protection than iron oxide. This process is also involvedwhen chromium is added to nickel to produce ‘stainless steel’, on which, thereaction 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 e

Zn ++

Zn

ZNSO4 2H+2e H2 Cu

--H2 SO4 2H+ + SO4

A Galvanic CellFig. 32

The zinc forms the anode of the cell, and is oxidised into ions that dissolve intothe acid. At the surface of the copper plate (the cathode), a balancing reactionoccurs. The electrons, formed in the anode, are conducted around the circuit andmeet with positively charged hydrogen ions at the cathode, to give off hydrogengas. The thermodynamic driving force of this cell is the difference in galvanicpotential between the two metals (zinc and copper). The metal of lower potential(the anode) in such a cell is oxidised or corroded.

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Similar electrochemical corrosion processes, with balancing anodic and cathodicreactions, occur in neutral (non-acidic) electrolytes such as water. The anodicreaction will involve oxidation (corrosion) of the metal with the lower galvanicpotential, but the cathodic reaction will, usually, be the reduction of oxygendissolved 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 corrosionoccurs in pure water (which has a high resistance), whilst quite severe corrosionoccurs in salt water which conducts electricity quite well. Adding variouschemicals to the electrolyte can change the resistance and, therefore, thereaction of the galvanic cell. Adding sodium chloride (salt) to the solution, lowersthe resistance of the circuit and, hence, increases the current. An acid, such ashydrochloric 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 metalsand alloys, can be measured and typical values found in solutions of seawater, orwater with 3.5% salt dissolved in it. Table 6 shows, in any combination of twometals, 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 electrolytepresent.

Table 6EXTRACT FROM THE GALVANIC SERIES

Extract from the Galvanic Series(Based on Hydrogen at 25°C (298 K))

Potential in Volts Material Anodic/Cathodic

-2.71 Sodium-2.38 Magnesium-1.66 Aluminium-1.63 Titanium-0.76 Zinc

-0.74 Chromium-0.44 Iron

-0.40 Cadmium-0.25 Nickel-0.14 Tin-0.13 Lead0 Hydrogen

+0.34 Copper+0.80 Silver+1.2 Platinum+1.43 Gold

Anodic

Cathodic

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Electrical Resistance: - As the corrosion products build up between twometals 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 evenhalting the reaction. Alternatively, the products can bridge any insulation, whichhas 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, thenthe rate of corrosion can be much faster (or slower), than would be obtained ifthey 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 theanode area, then the corrosion can be quite severe (refer to Fig. 33).

Aluminium Rivet Steel Rivet

Steel Sheet Aluminium Sheet

Effect of Anode and Cathode Areas on the Rate of CorrosionFig. 33

Single Metal Cells: - Corrosion can happen within alloys or metallic mixturesand can occur between metal grains and their grain boundaries, as well as inseveral other places. It can also occur if small metallic impurities are presentwithin a pure metal, even if the amount of impurity is merely a fraction of onepercent. The removal of impurities from metals, at the manufacturing stage, cangreatly improve their corrosion resistance.

Oxygen Concentration (Differential Aeration): - Corrosion can occur when thecomposition of the electrolyte varies at different parts of the contact area. Forexample, 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 acrevice perhaps), will be low in dissolved oxygen. The effect of this is to make themetal, close to the highly oxygenated part, a cathode and that in contact withlower 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 willalso have the effect of producing different potentials at these points. This canresult in severe corrosion in components such as radiators and heat-exchangers.

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11.3 TYPES OF CORROSIONThere 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 MetalIntergranular ExfoliationStress FrettingCrevice FiliformPitting Corrosion FatigueMicrobiological Hydrogen Embrittlement

11.3.1 SURFACE CORROSION

General roughening, etching or pitting of the metal surface, frequentlyaccompanied by a powdery deposit of corrosion products, may be caused bydirect chemical or electrochemical attack. Corrosion can spread under thesurface coating unnoticed, until the paint or plating is lifted off the surface by thecorrosion products or forms blisters.

Surface corrosion is a fairly uniform corrosion attack, which slowly reduces thecross-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 attackmay produce deposits which depend on the type of metal that is being attacked.

‘Pure’ aluminium, stainless steel and copper have more resistance to surfacecorrosion than aluminium alloy, magnesium alloy and non-stainless steels. Thistype of corrosion only becomes serious over a period of time and gives a warningof worse corrosion to follow.

11.3.2 DISSIMILAR METAL CORROSION

Galvanic action leads to one of the more common forms of corrosion, whichoccurs between two dissimilar metals in contact with each other and where thereis moisture present. It is caused by the difference in galvanic potential of the twometals where plating or jointing compound has been removed or omitted. Thistype of corrosion can occur, for example, where steel bolts, nuts, or studs are incontact with magnesium-rich alloys such as aircraft wheels.This may be taking place out of sight and may result in extensive pitting. It mayor 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 micro-galvanic cells at the grain boundaries in the metal (refer to Fig.34).

Corrosion progresses from the metal surface, in narrow pathways, along grainboundaries, often penetrating quite deeply and having a serious, mechanicalweakening effect. The amount of metal corroded is small, relative to the volumeof metal affected.

Indications of the damage may NOT be visible to the naked eye. Intergranularcorrosion may often be detected by ultrasonic, eddy current or radiographicinspection procedures.

Intergranular CorrosionFig. 34

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11.3.4 EXFOLIATION CORROSION

Exfoliation (or layer) corrosion, of certain wrought aluminium alloys, is a form ofintergranular corrosion in which the attack occurs in layers parallel to the surface.The wedging action, of the corrosion products, occupies a larger volume than thealloy, and will cause lifting of the metal surface, causing it to ‘exfoliate’. Thisoccurs at an early stage, when the corrosion is on, or just below, the surface.

Exfoliation corrosion often attacks 7000 series alloys (those with an appreciableamount of Zinc). When the corrosion occurs well below the surface, extensivedamage 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 ofcorrosion if they have been poorly heat-treated.

11.3.5 STRESS CORROSION

Stress corrosion cracking is a cracking process, caused by the combined actionof a sustained tensile stress and a corrosive environment. Only certaincombinations of alloys and environments result in stress corrosion cracking,although this type of failure may occur at stresses well below the yield strength ofthe alloys. Many of the high-strength structural alloys, used in aircraft, are proneto stress corrosion cracking in a wide range of environments and they areparticularly susceptible in marine environments.

In aircraft alloys, the principal stresses, causing this stress corrosion cracking, arenot the applied service loads, but the stresses developed within the metal duringmanufacture and during assembly. For example, internal stresses can arise fromquenching after heat-treatment, from ‘force fits’, from badly mating parts, or fromwelding procedures. Service stresses are only significant when they act in thesame 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 filmbreakdown). The incubation is followed by a period of ‘Slow Growth’ of the stressconcentrations and culminates in a short, ‘Rapid Crack-Growth’ rate.

In highly stressed parts (e.g. landing gear components), cracks may originatefrom a stress raiser such as a scratch or surface corrosion. This problem is

characteristic of aluminium, copper, stainless steels and high-strength alloy steelsand may occur along lines of cold working. Signs of stress corrosion are given byminute cracks radiating from areas of the greatest stress concentration. Likelyareas for this type of corrosion are U/C jacks, shock absorbers, bellcranks withpressed-in bushes, or other areas where parts are a force fit, highly stressed orhave 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 loadedsurfaces, one, or both, of which are metallic. The rubbing action destroys anynatural protective film and also removes particles of metal from the surface. In itsearly stages, the debris of this corrosion forms a black powder. These particlesform an abrasive compound, which aggravates the effect of the rubbing actionand the surface is continually removed to expose fresh metal to the corrosiveattack. This form of attack can eventually cause cracking and fatigue failure.

The most likely areas affected are gears, screw jacks, loose panels, splinedhydraulic pump drives and rivets (when they become loose). , It may be seriousenough to cause cracking and fatigue failure.

11.3.7 CREVICE CORROSION

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

Low Oxygen Concentration(becomes anodic)

HighOxygen

Concentration(becomes cathodic)

Crevice

Crevice CorrosionFig. 35

Severe localised corrosion occurs at narrow openings or gaps between metalcomponents, often due to flexing. Corrosive agents are able to penetrate into thejoint.

11.3.8 FILIFORM CORROSIONFiliform corrosion occurs beneath thin, protective coatings, on aluminium andsteel alloys, with the paint or coating often bulging or blistering. On aircraftstructures, the attack often starts at fasteners and extends as thread-like lines ofcorrosion under the paint. It may not be readily visible until it has become quitesevere. The damage tends to be very shallow and is not, usually, structurallydangerous.

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11.3.9 PITTING CORROSION

Pitting corrosion can occur on aircraft materials when the protective film, whetherapplied or natural, breaks down locally and this may also lead to intergranularcorrosion. The corrosion often stems from the screening effect of silt, scale orcorrosion deposits that reduce the oxygen concentration at local points on themetal surface, which establishes differential concentration cells.

Local rough spots, inclusions, contaminations and lack of homogeneity in thealloy or metal are also possible causes of pitting. In size and depth, the pits arewidely 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 ofcorrosion. Pitting corrosion of an aluminium alloy component can be detected bythe appearance of white powder on the surface of the metal (refer to Fig. 36).

Pitting Corrosion of an Aluminium Alloy ComponentFig. 36

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11.3.10 CORROSION FATIGUE

This is similar to stress corrosion cracking, except that the applied loads arecyclic instead of static. Crack propagation is aided by the corrosion that occurs, atthe 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 micro-organisms which are not only able to produce corrosive substances (such ashydrogen sulphide, ammonia and inorganic acids), but can also act asdepolarisers or catalysts in corrosion reactions. Local depletion of oxygen andwater, 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 iscaused by contamination of fuel tanks (unless the fuel has an additive to protectagainst it). The growth of the fungi depends on several conditions, but a highambient temperature can drastically increase the rate of growth, and especiallyso when the temperature is above 30C with a high relative humidity. Thismicrobiological growth is sometimes called Cladosporium Resinæ.

Where fungal growth has formed, there is a probability that corrosion of the tankwill occur. The organisms, resembling a mucous, can cause problems with filtersand with the fuel contents gauge units. The roots of the fungus, penetrating theinternal 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 thetank, retaining water and preventing free flow to the water drain-valve. In integralfuel tanks, this can result in serious corrosion of the aircraft structure such thatpenetration of the bottom wing skin has been known to occur.

Spillage, of organic materials, from around galley and toilet areas, provides afurther source of microbial contamination.

There is evidence that such spillage can be more corrosive than its chemicalcomposition (acidity and chloride content) possibly due to fermentation by yeastand bacteria.

11.3.12 HYDROGEN EMBRITTLEMENT OF STEELS

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

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Hydrogen embrittlement shows itself in slow strain-rate mechanical tests and notby fast rate tests such as in impact testing. These steels can show a suddenfailure after many weeks of loading at well below their normal yield strength.

11.4 FACTORS AFFECTING CORROSIONMany factors will affect the cause, type, speed of attack, and seriousness ofmetal corrosion. Some are beyond the control of the aircraft designer ormaintenance engineer while some of them can be controlled.

11.4.1 CLIMATIC

The environmental conditions under which the aircraft is operated and maintainedcannot normally be controlled. The following factors will effect the rate at whichcorrosion 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. Hightemperatures will increase the rate of corrosion (all chemical reactionsoccur faster at higher temperatures).

The worst conditions would exist in a hot, wet, maritime environment.

11.4.2 SIZE AND TYPE OF METALSome metals corrode more easily than others. Magnesium corrodes readily,

whilst Titanium is extremely corrosion-resistant because it oxidises readily. Thickstructural sections are also more susceptible than thin sections, becausevariations in physical characteristics are greater. Such sections are also likely tohave 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 incorrosion, 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 PRODUCTSOne of the problems involved in corrosion control, is the recognition of corrosionproducts whenever they occur. The following brief descriptions are of typicalcorrosion 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. Theinitial oxide film, formed on freshly exposed steel, is very thin and invisible. In thepresence of water, or in a damp atmosphere, especially if sulphur dioxide(industrial atmosphere) or salt (marine environment) is present, thick layers ofhydrated oxide develop. These layers vary in colour from brown to black. Rustpromotes further corrosion by retaining salts and water. Mill scale (a type of oxideformed at high temperatures), also promotes rusting, by forming an electrolyticcell with the underlying steel. Heavy deposits of rust can be removed only byabrasive blasting or by immersion in rust-removing solutions.

Surface rust can develop on steel nuts, bolts and other fasteners and may notadversely affect the operational integrity of the equipment. Its appearance is anindication 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. Itmay vary from general etching of the surface, to the localised, intergranular-attack, characteristics of some strong alloys in certain states of heat-treatment.The corrosion products are white to grey and are powdery when dry. Superficialcorrosion can be removed by scouring, light abrasive blasting, or by chemicalmethods.

In general, pure aluminium sheet and ‘alclad’ surfaces have good corrosionresistance, except in marine environments. In these areas, aluminium and itsalloys need protection and high-strength aluminium alloys are always given asubstantial protective treatment.

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11.5.3 MAGNESIUM ALLOYSMagnesium corrosion products are white and voluminous, compared to the basemetal. When the failure of protective coatings on magnesium alloys occurs, thecorrosive attack tends to be severe in the exposed areas, and may penetratetotally through a magnesium structure in a very short time. Any corrosion, onmagnesium alloys, therefore requires prompt attention. In contrast to high-strength aluminium alloys, the strong magnesium alloys, used in aircraft, do notsuffer intergranular attack. Corrosion is readily visible on the surfaces ofMagnesium Alloys.

11.5.4 TITANIUM

Titanium is highly corrosion-resistant, but should be insulated from other metalsto avoid dissimilar metal corrosion of the adjacent material. Titanium alloys cansuffer stress corrosion at temperatures above 300C when in the presence of saltand fatigue cracks can develop more quickly in a saline atmosphere.

Cadmium can penetrate the surface of titanium alloys and embrittle them at alltemperatures above ambient (as can Lead, Tin and Zinc at temperatures higherthan approximately 120°C)). Embrittlement can occur if the cadmium is platedonto the titanium or if cadmium-plated steel parts (and cadmium-contaminatedspanners) are used with titanium. Great care must be taken to ensure that theseconditions never occur if at all possible.

11.5.5 COPPER ALLOYS

Copper and its alloys are relatively resistant to corrosion. Tarnishing has noserious consequences in most applications. Long-term exposure to industrial ormarine atmospheres gives rise to the formation of the blue-green patina (aerugoor verdigris) on copper surfaces, while brasses can suffer selective removal ofzinc (de-zincification). In aircraft construction, copper-based alloys are frequentlycadmium-plated, to prevent dissimilar metal corrosion.

11.5.6 CADMIUM AND ZINCCadmium and zinc are used as coatings, to protect the parts to which they areapplied. Both confer sacrificial protection on the underlying metal. Cadmium isnormally chosen for use in the aircraft industry, as it is more durable under severecorrosive conditions such as in marine and tropical environments. Both metalsproduce 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 usedfor its wear-resistance. Both metals protect steel only by excluding the corrosiveatmosphere. The degree of protection is proportional to the thickness of thecoating. Once the underlying steel is exposed (through loss of the coating, due toabrasion or other damage), then the coatings actually accelerate the rusting, dueto the fact that the steel is more anodic than the protective coating.

Chromium is also highly resistant to corrosion, whilst Nickel corrodes slowly inindustrial and marine atmospheres, to give a blue-green corrosion product.

11.6 CORROSION REMOVALGeneral 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 orpreservatives. This will aid in determining the extent of corrosive spread. Theselection 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 compoundsmay need solvent/emulsion-type cleaners.

General purpose, water-removable stripper is recommended for most paintstripping. Adequate ventilation should be provided and synthetic rubber surfacessuch as tyres, fabric and acrylics should be protected (remover will also softensealants). Rubber gloves, acid-repellent aprons and goggles, should be worn bypersonnel involved with paint removal operations. The following is the generalpaint 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-metallicscrapers may be used.

Remove the loosened paint and residual stripper by washing and scrubbingthe surface with water and a broom or brush. Water spray may assist, or theuse of steam cleaning equipment may be necessary.Note. Strippers can damage composite resins and plastics, so every effortshould 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 bedeposited. Some metal oxides protect the underlying base metal, but rustpromotes additional attack by attracting moisture and must be removed.

Rust shows on bolt heads, nuts or any un-protected hardware. It’s presence isnot immediately dangerous, but it will indicate a need for maintenance and willsuggest possible further corrosive attack on more critical areas. The mostpractical means of controlling the corrosion of steel is the complete removal ofcorrosion products by mechanical means.

Abrasive papers, power buffers, wire brushes and steel wool are all acceptablemethods of removing rust on lightly stressed areas. Residual rust usually remainsin pits and crevices. Some (dilute) phosphoric acid solutions may be used toneutralise oxidation and to convert active rust to phosphates, but they are notparticularly effective on installed components.

11.6.3 HIGH-STRESSED STEEL COMPONENTSCorrosion on these components may be dangerous and should be removed

carefully with mild abrasive papers or fine buffing compounds. Care should betaken not to overheat parts during corrosion removal. Protective finishes shouldbe re-applied immediately.

11.6.4 ALUMINIUM AND ALUMINIUM ALLOYSCorrosion attack, on aluminium surfaces, gives obvious indications, since the

products are white and voluminous. Even in its early stages, aluminium corrosionis 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 confusedwith the more serious forms of corrosion.

General surface attack penetrates slowly, but is speeded up in the presence ofdissolved salts. Considerable attack can take place before serious loss ofstrength 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 corrosionproducts 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 OFREDUX BONDED JOINTS. CERTAIN PAINT STRIPPERS WILL ATTACK ANDDEGRADE RESINS. USE ADEQUATE PERSONAL PROTECTIVE EQUIPMENTWHEN WORKING WITH CHEMICALS. USE ONLY THE APPROVED FLUIDSFOR REMOVING CORROSION PRODUCTS. INCORRECT COMPOUNDSWILL CAUSE SERIOUS DAMAGE TO METALS.

Obviously great care must be taken, not to remove too much of the protectivealuminium layer by mechanical methods, as the core alloy metal may beexposed, therefore, where heavy corrosion is found, on clad aluminium alloys, itmust be removed by chemical methods wherever possible.

Corrosion-free areas must be masked off and the appropriate remover (usually aphosphoric-acid based fluid) applied, normally with the use of a stiff bristledbrush, to the corroded surface, until all corrosion products have been removed.

Copious amounts of clean water should, next, be used to flood the area andremove all traces of the acid, then the surface should be dried thoroughly.

Note: A method of checking that the protective aluminium coating remains intactis by the application of one drop of diluted caustic soda to the cleaned area. If thealclad 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 aprotective coating (usually Alocrom 1200 or similar) is applied to the surface.

Further surface protection may be given by a coat of suitable primer, followed bythe approved top coat of paint.

11.6.6 MAGNESIUM ALLOYS

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

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11.6.7 ACID SPILLAGE

An acid spillage, on aircraft components, can cause severe damage. Acids willcorrode most metals used in the construction of aircraft. They will also destroywood and most other fabrics. Correct Health and Safety procedures must befollowed when working with such spillages.

Aircraft batteries, of the lead/acid type, give off acidic fumes and battery baysshould be well ventilated, while surfaces in the area should be treated with anti-acid paint. Vigilance is required of everyone working in the vicinity of batteries, todetect (as early as possible) the signs of acid spillage. The correct procedure tobe 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 tospread the acid.

If possible, flood the area with large quantities of clean water, taking carethat 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 ifacid has been cleaned up.

Dry the area completely and examine the area for signs of damaged paint orplated finish and signs of corrosion, especially where the paint may havebeen damaged.

Remove corrosion, repair damage and restore surface protection asappropriate.

11.6.8 ALKALI SPILLAGE

This is most likely to occur from the alternative Nickel-Cadmium (Ni-Cd) or Nickel-Iron (Ni-Fe) type of batteries, containing an electrolyte of Potassium Hydroxide(or Potassium Hydrate). The compartments of these batteries should also bepainted with anti-corrosive paint and adequate ventilation is as important as withthe lead/acid type of batteries. Proper Health and Safety procedures are, again,imperative.

Removal of the alkali spillage, and subsequent protective treatment, follows thesame basic steps as outlined in acid spillage, with the exception that the alkali isneutralised 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 TOTHE HEALTH AND SAFETY EXECUTIVE. ALL SAFETY PRECAUTIONSRELATING TO THE SAFE HANDLING OF MERCURY MUST BE STRICTLYFOLLOWED.

Mercury contamination is far more serious than any of the battery spillages andprompt 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 mercurycan, 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 andaccelerate powder propagation, resulting in a potentially catastrophic weakeningof the aircraft structure.

Signs of mercury attack on aluminium alloys are greyish powder, whiskerygrowths, or fuzzy deposits. If mercury corrosion is found, or suspected, then itmust be assumed that intergranular penetration has occurred and the structuralstrength is impaired. The metal in that area should be removed and the arearepaired in accordance with manufacturer’s instructions.

Ensure that toxic vapour precautions are observed at all times during thefollowing 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 allglobules have been removed and to check extent of corrosion damage.

Examine area for corrosion using a magnifier. Any parts found contaminatedshould be removed and replaced.

Note: Twist drills (which may be used to separate riveted panels, in an attemptto clean contaminated surfaces) must be discarded after use.

Further, periodic checks, using radiography, will be necessary on anyairframe that has suffered mercury contamination.

11.7 PERMANENT ANTI-CORROSION TREATMENTSThese are intended to remain intact throughout the life of the component, asdistinct from coatings, which may be renewed as a routine servicing operation.They give better adhesion for paint and most resist corrosive attack better thanthe 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 basemetal. Commonly called sacrificial protection, an example is found in thecadmium (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 toacid attack. If, however, the basic metal is exposed, it will corrode locallythrough electrolytic action. The attack may result in pitting corrosion of thebase metal or the corrosion may spread beneath the coating.

11.7.2 SPRAYED METAL COATINGSMost metal coatings can be applied by spraying, but only aluminium and zinc areused on aircraft. Aluminium, sprayed on steel, is frequently used for high-temperature 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 beendiscussed, 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 immediatesurface layer into a film of metal oxide, which has better corrosion resistance thanthe metal. Among those widely used on aircraft are:

Anodising of aluminium alloys, by an electrolytic process, which thickens thenatural, oxide film on the aluminium. The film is hard and inert.

Chromating of magnesium alloys, to produce a brown to black surface filmof 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 thephosphating of steel. There are numerous proprietary processes, each known byits trade name (e.g. Bonderising, Parkerising, or Walterising).

11.8 LOCATIONS OF CORROSION IN AIRCRAFTCertain locations in aircraft are more prone to corrosion than others. The rate ofdeterioration varies widely with aircraft design, build, operational use andenvironment. External surfaces are open to inspection and are usually protectedby paint. Magnesium and aluminium alloy surfaces are particularly susceptible tocorrosion along rivet lines, lap joints, fasteners, faying surfaces and whereprotective coatings have been damaged or neglected.

11.8.1 EXHAUST AREAS

Fairings, located in the path of the exhaust gases of gas turbine and pistonengines, are subject to highly corrosive influences. This is particularly so whereexhaust deposits may be trapped in fissures, crevices, seams or hinges. Suchdeposits are difficult to remove by ordinary cleaning methods.

During maintenance, the fairings in critical areas should be removed for cleaningand examination. All fairings, in other exhaust areas, should also be thoroughlycleaned and inspected. In some situations, a chemical barrier can be applied tocritical areas, to facilitate easier removal of deposits at a later date, and to reducethe 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 byrain. Heat-exchanger cores and cooling fins may also be vulnerable to corrosion.

Special attention should be given, particularly in a corrosive environment, toobstructions and crevices in the path of cooling air. These must be treated assoon 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, andrequire frequent cleaning and touching-up. Careful inspection should be given tocrevices, ribs and lower-skin surfaces, where debris can lodge. Landing gearassemblies should be examined, paying particular attention to magnesium alloywheels, 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 wateror with the water-dispersing fluids, used when re-lubricating.

11.8.4 BILGE AND WATER ENTRAPMENT AREASAlthough specifications call for drains wherever water is likely to collect, thesedrains can become blocked by debris, such as sealant or grease. Inspection ofthese drains must be frequent. Any areas beneath galleys and toilet/wash-roomsmust be very carefully inspected for corrosion, as these are usually the worstplaces in the whole airframe for severe corrosion. The protection in these areasmust 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, wherewater and dirt may collect and go unnoticed, because the moveable parts arenormally in the ‘closed’ position. If these items are left ‘open’, when the aircraft isparked, they may collect salt, from the atmosphere, or debris, which may beblowing about on the airfield. Thorough inspection of the components and theirassociated 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 thehinges 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 undamagedand well maintained. Following maintenance work, such as riveting and drilling, itis impossible to completely protect the skin to the original specification. Allmagnesium alloy skin areas must be thoroughly and regularly inspected, withspecial emphasis on edge locations, fasteners and paint finishes.

11.8.7 ALUMINIUM ALLOY SKINSThe most vulnerable skins are those which have been integrally machined,usually in main-plane structures. Due to the alloys and to the manufacturingprocesses used, they can be susceptible to intergranular and exfoliationcorrosion. Small bumps or raised areas under the paint sometimes indicateexfoliation of the actual metal. Treatment requires removal of all exfoliated metalfollowed 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-weldedskins and moisture, entering the seams, may set up electrolytic corrosion thateventually corrodes the spot-welds, or causes the skin to bulge. Generally, spot-welding is not considered good practice on aircraft structures.

Cavities, gaps, punctures or damaged places in honeycomb sandwich panelsshould be sealed to exclude water or dirt. Water should not be permitted toaccumulate in the structure adjacent to sandwich panels. Inspection ofhoneycomb sandwich panels and box structures is difficult and generally requiresthat the structure be dismantled.

11.8.9 ELECTRICAL EQUIPMENTSealing, venting and protective paint cannot wholly obviate the corrosion inbattery compartments. Spray, from electrolyte, spreads to adjacent cavities andcauses rapid attack on unprotected surfaces. Inspection should also be extendedto all vent systems associated with battery bays.

Circuit-breakers, contacts and switches are extremely sensitive to the effects ofcorrosion and need close inspection.

11.8.10 MISCELLANEOUS ITEMS

Loss of protective coatings, on carbon steel control cables can, over a period oftime, 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 thoroughinspection of the internal strands and, if any damage is found, the cable shouldbe rejected.

Cables should be carefully inspected, in the vicinity of bell-cranks, sheaves and inother places where the cables flex, as there is more chance of corrosion gettinginside the cables when the strands are moving around (or being moved by) theseitems.

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12 AIRCRAFT FASTENERS

WARNING: ONLY THE APPROVED METHODS OF FASTENING DEVICESMUST BE USED ON AIRCRAFT. SUBSTITUTION WITH INCORRECT PARTSCAN CAUSE FATAL FAILURES.

Fasteners, or fastening devices, are used to create secure joints between two ormore components. Types of fastening devices, used on aircraft, vary inaccordance with the materials, which require joining, and the importance of thejoined 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 orrepair, will also influence the choice of fasteners to be employed. Fasteners maybe metallic or non-metallic (or composites of both types). They may be flexible orrigid(or a combination of both) and may be used to form the three basiccategories of joints.

12.1 TEMPORARY JOINTSTemporary joints are used where the joint can be disassembled without damageand where, usually, the same fastener can be used to reassemble the joint. Boltsand nuts, circlips and quick-release fasteners are, typically, used in temporaryjoints.

12.2 PERMANENT JOINTSPermanent joints are those which are not intended to be disassembled on afrequent basis (if at all), and are joints where either the fastening medium or thejoined components will suffer damage in their separation. Adhesives, rivets andwelds are examples of uses of permanent joints.

12.3 FLEXIBLE JOINTSFlexible joints allow movement of the joined components relative to each other.Anti-vibration mounts, universal couplings and hinges are devices which may beemployed in flexible joints.

Whatever fasteners are used, to make a particular joint, it must be ensured thatonly the approved materials are utilised and that their legality is confirmed. Thiscan be done by reference to published Part Numbers, which are to be found inAircraft 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, fatalfailures in aircraft and their associated structures.

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12.4 SCREW THREADSThreaded fasteners allow parts to be fastened together with all of the strengththat unthreaded fasteners provide. However, unlike rivets and pins, threadedfasteners may be disassembled and reassembled an almost infinite number oftimes.

Due to the large range of different available fasteners, great care must be alwaysbe 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 seenin Table 7.

Most aircraft now use unified or metric threads but, however, some older aircraftuse 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 7COMMON INTERNATIONAL THREAD STANDARDS

International Standard Common Abbreviation

American National Coarse ANCAmerican National Fine ANFUnified Coarse UNCUnified Fine UNFBritish Association BABritish Standard Fine BSFISO Metric M

12.4.1 THE INCLINED PLANE AND THE HELIXThe 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 inclinationbetween the planes will cause a greater force to be exerted through a shorterdistance. Conversely, a larger angle will cause less force to be exerted through agreater distance (refer to Fig. 37).

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Transmission of Motion with an Inclined PlaneFig. 37

Whilst the wedge is, generally, used as a means of transmitting motion, it must beremembered that the action may be reversed and the wedge can be caused tomove when a force is applied to the inclined surfaces.

This is readily appreciated when the angle is large (and the larger the angle ofinclination becomes, then, the more readily is the motion reversed), but, nomatter how small the angle may be, the resultant of forces applied will still tend toproduce movement. Friction, between the surfaces, may, however, preventmovement from actually occurring.

When a continuous, inclined plane is cut around the outside (or the inside) of acylinder, then a spiral (also known as a ‘helix’) is produced (refer to Fig. 38). Thehelix angle is important in screw threads, because it dictates the number ofthreads, which can be cut, per axial linear increment (millimetres or inches) on, orin, the cylinder.

Helix Angle

Helix Angle of a Screw ThreadFig. 38

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In a similar manner to the previously mentioned wedges, a thread with a smallhelix angle (a fine thread), will exert a greater force than one with a larger helixangle (a coarse thread) for a given cylinder diameter.

Fine threads are, normally, associated with small and delicate instruments or inequipment, where secure holding power is often required of miniature-sizedfasteners. The greater ‘wedging action’ of fine threads also makes them muchmore dependable in situations where vibration (or a change of temperature) hasthe 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 temperaturechanges and vibration.

12.5 SCREW THREAD TERMINOLOGYIt 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, whichhas a definite plain portion on the shank, between its head and the beginning ofthe thread, and is used in conjunction with a nut, whereas a screw is threaded allthe way to the head.

Because there are so many variations in terminology, with the numerousmanufactures, the only safe way of describing a threaded (or any other) fasteneris to use the correct terminology, found in the relevant IPC, when orderingreplacement items.

When defining the length of bolts, reference is usually made to the length of theplain portion of the shank, of hexagonal-headed bolts (refer to Fig. 39), whilescrew lengths are designated differently, according to their type.

Designation of Fastener LengthsFig. 39

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Screw threads are usually formed with a ‘clockwise’ turning groove and arereferred to as ‘right-hand’ threads, but there are occasions where the thread isformed with the groove spiralling in an ‘anti-clockwise’ direction and, in thisinstance, 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 tobe put.

Screw Thread TerminologyFig. 40

The following terms are used to define the characteristics of a threaded item:

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

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

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

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

Multi-Start Thread: Consists of two or more separate, parallel threads cutinto the material carrying the thread. This method is used in order toachieve a quick-acting motion between two threaded items.

Runout: The part of the thread where the minor diameter increases until itequals the major diameter and merges with the plain portion of the shank.The runout cannot be used and any nut, rotated onto the runout, wouldbecome ‘thread-bound’.

12.5.1 SCREW THREAD FORMS

The form of a screw thread will depend upon the function for which it is to beused (refer to Fig. 41).

Where the thread is used to join components together (nuts, bolts, screws andstuds) then the conventional, truncated ‘V’-shaped threads, similar to the ISOMetric thread, will be found.

Turnbuckles and similar devices, (which are employed as adjusters of either thetension or of the distance between components), may also use‘V’-shapedthreads, while the Acme, Buttress and Square threads are utilised to transmitmovement or power (as may be seen in lathes, vices and Flap Jacks).

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Pitch

Pitch

60°

ISO Metric Thread

Pitch

45°

ButtressThread

Screw Thread FormsFig. 41

Square Thread

Pitch

29°

Acme Thread

Thread forms have developed over the years, from the early standardisation onthe BSW thread (with its rather coarse thread, which was prone to slackeningwhen subjected to vibration), to the modern, finer threads which are more suitablefor use on aerospace components and structures.

In an attempt to provide a common standard, Canada, the United States ofAmerica, and the United Kingdom adopted the Unified system of threads.

The International Standard Organisation (ISO), later, recommended that theUnified system be used internationally, in parallel with a system using Metric unitsof measurement, but with a similar form of thread profile and standards oftolerances

Unified Coarse (UNC) and Unified Fine (UNF) threads may be found wherevertheir use is appropriate, but special threads, such as UNS (for high-temperatureapplications) and UNJ (increased fatigue strength) have become more common.

Screw threads may be formed, by such processes as tapping, dieing, andmachine cutting or (where maximum fatigue resistance is required of a bolt), byrolling.

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12.5.2 OTHER THREAD FORMS

In the United States, a parallel but differing group of thread systems is used. Thefour main systems are ANC, ANF (also referred to as AF), UNC and UNF, withthe 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 alsoclassified by their class of fit, as can be seen in Table 8.

Table 8CLASSES OF THREAD FITS

Class of Fit Type of Fit1 Loose2 Free3 Medium4 Close5 Tight

A Class 1 fit can be tightened, all the way down, by hand (such as with a wing-nut), whilst a Class4 or 5 fit requires a spanner throughout the tighteningoperation.

The Class 3 fit is the type mostly employed on aircraft, and would be typical of athread which is designed for use in a high-temperature environment and mayrequire the application of an anti-seize compound before installation.

By comparison, a fastener which is going to be subjected to the high tension orshear loads, associated with the securing of aircraft engine parts, would need tobe 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 itsidentification can be found in the IPC and supplied under a manufacturer’s partnumber. Whilst this is true and the manuals should always be used, there areother ways of identifying screw threads.

One method is to identify the screw by means of various marks, normally foundon the head of the screw. These marks may give a clue as to which type ofthread the screw has (AF, BSF, or Metric etc.). A measurement across the threadcrests, using a micrometer, would give the diameter of the screw in question.Finally, the identifying head markings would also give the material from which thescrew is made.

Two useful tools (refer to Fig. 42) may be used for different stages of threadmeasurement.

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 variousblades of the gauge against the screw thread until a match is achieved

Profile Gauge

55° 47½° PitchGauge

60°

0.25 - 2.5 mm

60°

Thread Profile Gauge and Screw Pitch GaugeFig. 42

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12.6 BOLTSThe bolts, used in the construction of aerospace components and structures,

have evolved into a bewildering range of materials, shapes and sizes, all of whichare dictated by the applications for which the items have been designed

Standards and systems have been established, to provide identification of themany different forms of threaded devices, in order to ensure that only the correctitems are installed in the relevant locations.

It is stressed here, that only the approved design materials may be used foraerospace components and, while a selection of some of the bolts are presentedin these course notes, by way of introduction, the relevant AMM, SRM and IPCwill be the sole authority for deciding the correct type of bolt that is to be used in aparticular application.

12.6.1 BRITISH BOLTS

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

AGS Aircraft General Standard AS Aircraft Standards Al. Al. Aluminium Alloy BA British Association BSF. British Standard Fine HTS. High Tensile steel HTSS. High Tensile Stainless Steel LTS. Low Tensile Steel SS Stainless Steel UNC. Unified National Coarse UNF. Unified National Fine.

12.6.2 IDENTIFICATION OF BS UNIFIED BOLTSBritish Standard Unified (BS Unified) bolts are identified by the use of an alpha-numeric code, which provides information relating to the type, material, surfacefinish, length, diameter and any other important characteristics of the threadeddevice

Table 9 shows a (very small) selection of aircraft standard bolts and screws witha (shortened) description of the type of device and the materials from which it ismade.

Reference to the table shows that the code A102 signifies a hexagonal-headedbolt which is made of high-tensile steel, while the code A175 represents a 100°countersunk-headed bolt, made from an aluminium alloy.

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Table 9Examples of Code Numbers for Unified Threads

Standard No. Description MaterialA102 Hex. Headed Bolt HTS.A104 Hex. Headed Bolt SSA111 Hex. Close Tolerance. Bolt HTSA112 Shear Bolt HTSA174 100º Countersunk. Head. Bolt SSA175 100º Countersunk. Head. Bolt Al AlA204 100º Countersunk. Head. Screw HTSA205 Pan Head. Screw HTS

Other methods of indicating that an item has a Unified thread are:

Three contiguous (touching) circles marked in a convenient position (machineitems).Note: Due to the difficulty in applying the identifying marks to individual items, it isplanned 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 theitems 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 thethread (cold forged items).

A ‘dog point’ (small protrusion) on the threaded shank end (usually applies toscrews).

Further numbers and letters are added to the identifying code, to provideinformation 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 signifiesincrements of tenths of an inch, so that a 5 would represent a bolt with a plainshank of 0.5 in, while the number 12 would signify the plain shank as being 1.2 inlong

Reference to Table 10, will show how the diameter of an item is designated bythe addition of another letter to the system, so that a bolt, with the code markingof A102 9 E, would signify a Unified-threaded, hexagon-headed bolt, made fromhigh-tensile steel, with a plain shank length of 0.9 in, and a diameter of ¼ in.

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Table 10EXAMPLES OF BS UNIFIED BOLT CODES

Code Diameter Code DiameterY 0-80 UNF J 3/8" UNF (UNJF)Z 2-64 UNF L 7/16" UNF (UNJF)A 4-40 UNC N 1/2" UNF (UNJF)B 6-32 UNC P 9/16" UNF (UNJF)C 8-32 UNC Q 5/8" UNF (UNJF)D 10-32 UNF UNJF) S 3/4" UNF (UNJF)E 1/4" UNF (UNJF) U 7/8" UNF (UNJF)G 5/16" UNF (UNJF) W 1" UNF (UNJF)

Note: In the earlier UK system (which may be encountered on older, or home-constructed, 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 theseitems also use a number to represent their nominal length and a letter code (ascan be seen in Table 11) to identify their diameter.

Other bolts of this era may have nicks at the corners of the head (High TensileSteel) or a raised ring on the bolt head (Cold Rolled) to assist differentiation oftheir particular designations.

Table 11EXAMPLES OF BA AND BSF BOLT AND SCREW CODES

Code Size Code Size

ABCEGJLN

6 BA4 BA2 BA

1/4” BSF5/16" BSF3/ 8" BSF

7/16" BSF1/2" BSF

P 9/16" BSFQ 5/8” BSFS 3/4" BSFU 7/8" BSFW 1" BSFX 12 BAY 10 BAZ 8 BA

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12.6.3 AMERICAN BOLTS

American aircraft bolts and nuts are threaded in the NC (American NationalCoarse), the NF (American National Fine), the UNC (Unified National Coarse),

and the UNF (Unified National Fine) thread series. The item is often coded to givethe diameter of the threaded portion and the number of threads per inch (tpi).

Aircraft bolts may be made from HTS, Corrosion-Resistant Steel or AluminiumAlloy. Head types may be hexagonal, clevis, eyebolt, internal wrenching andcountersunk (refer to Fig. 43) and head markings may be used to indicate otherfeatures such as close tolerance, aluminium alloy, CRS or other types of steel.

Examples of Aircraft BoltsFig 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 onetype 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 andTable 12 provides an indication of the various code numbers in use.

Table 12EXAMPLES OF AN STANDARD BOLTS (EARLY SERIES)

AN No. Type Material Process Thread ThreadSize Type

3 - 20 Bolt, hex. Steel Cadmium No. 10 to UNFHead Plated 1¼”

CRS NilAl. Al. Anodised

21 - 36 Bolt, Steel Cadmium No. 6 to UNFClevis Plated 1”

42 - 36 Bolt, Eye Steel Cadmium No. 10 to UNFPlated 9/16”

73 - 81 Bolt, hex. Steel Cadmium No. 10 to UNF orDrilled Plated ô” UNChead

173 - 186 Bolt, close Steel Cadmium No. 10 to UNF- tolerance Plated 1”

thread &head

Note: The later series uses a different number system

For identification purposes the AN number is used to indicate the type of bolt andits diameter. In addition a code is used to indicate the material, length andpresence of a split pin or locking wire hole as follows:

Diameter: The last figure, or last two figures, of the AN number indicatesthread diameter, 1 = No. 6, 2 = No.8, 3 = No.10, and 4 = ¼” withsubsequent numbers indicating the diameter in 1/16” increments.

Thus an AN4 is a hexagon headed bolt of ¼” diameter and an AN14 is ahexagon 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, ismeasured from under the head of the first full thread (refer to Fig. 44) and isquoted in 1/8” increments as a dash number.

The last figure of the dash number represents eighths and the first figureinches, so that an AN4 - 12 is a ¼” diameter hexagon headed bolt, 1¼”long.

DrilledShank

Steel CRS Steel, Close Tolerance CRS, Close Tolerance

Diameter

Aluminium Alloy Drilled Head, Drilled Head, Aluminium Alloy,(Except AN 73 -81) AN 73 -81 Close Tolerance

Head Markings for AN BoltsFig. 44

Length‘L’

Grip

Position of Drilled Hole: Bolts are normally supplied with a hole drilled inthe threaded part of the shank, but different arrangements may beobtained:

Drilled shank

Un-drilled shank

Drilled head only

Drilled head and shank

= normal coding e.g. AN24 - 15

= A added after dash No. e.g. AN24 - 15A

= H added before dash No.(replacing dash) A added e.g. AN25H15Aafter dash No.

= 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 UNCthread 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 purposestructural bolt used for applications involving tension or shear loads where a lightdrive fit is permissible.

Eye Bolt

Clevis Bolt

Special-to-Type BoltsFig. 45

Alloy steel bolts, smaller than 3/16” diameter, and aluminium alloy bolts smallerthan ¼” are not used on primary structure. Other bolts may be used as follows:

Close Tolerance Bolts: These bolts are machined more accurately than thestandard bolt. They may be hexagon headed (AN173 - AN186) or have a100º countersunk head (NAS80 - NAS86). They are used in applicationswhere 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 fabricatedfrom high-strength steel and are suitable for tensile or shear applications.The head is recessed to allow the insertion of a hexagonal key used forinstalling or removing the bolt. In Dural-type material, a heat-treated washermust 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 astandard screwdriver, or recessed, for a cross-pointed screwdriver. Thistype of bolt is used only for shear loads and never in tension. It is ofteninserted as a mechanical pin in a control system.

Eyebolt: The eye is designed for the attachment of cable shackles orturnbuckles and the bolt is used for tensile loads. The threaded end may bedrilled 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 mm-diameter bolt. The length is also shown in millimetres, so the bolt M6 -15 will be a6 mm- diameter bolt, which is 15 mm long. The basic terminology, for identifyingbolts of the Metric system, involves the nominal length, the grip length anddiameter (refer to Fig. 46).

Length

Grip

Diameter

Metric Bolt TerminologyFig. 46

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12.7 NUTSAerospace standard nuts are made in a variety of shapes and sizes. They can bemade of cadmium-plated carbon steel, stainless steel or anodised2024-Taluminium 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 orscrew to which they are attached, with the exception of high-tensile steel bolts,with which, mild steel nuts are used.

Selection of Typical NutsFig. 47

As they do not have any identifying marks or lettering, they are usually identifiedby their colour and their constructional features. Familiar types of nuts include:

Castle Nuts: which are used with drilled shank hexagon-headed bolts orstuds, eye-bolts and clevis bolts. They are fairly rugged and can withstand largetensile 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 insimilar applications, except that they are normally used for engine use only.

Plain Hexagon Nuts: are of rugged construction and suitable for large tensileloads. Since they require an auxiliary locking device, their use on aircraft islimited.

Light Hexagon Nuts: are a much lighter nut, used for miscellaneous lighttensile requirements.

Plain Check (or Lock) Nuts: are employed as locking devices for plain nuts, forthreaded rod ends and for other devices.

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Wing Nuts: are used where the desired tightness can be obtained merely withusing 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 threadsto ensure its tightness. The enemy of this system is vibration, which can causethe nut to slacken off, and in extreme cases, unwind itself completely from thebolt or screw.

In areas where this might occur, locking devices are used. These either increasethe frictional resistance between the threads, or take the form of positivesecurities that prevent any movement of the nut once they have been applied.

Stiffnuts and anchor nuts (refer to Fig. 48) employ various means of increasingthe friction forces between the threaded devices and common types include:

Nyloc: This looks like a standard hexagonal nut, but has a plastic insert in thecounter-bored end. This insert is initially unthreaded and has an internal diameterslightly smaller than the nut thread, so that, as the nut is screwed on the bolt, theplastic insert is displaced and a high degree of friction is created. Another type ofplastic ‘stop’ nut is named the ‘Capnut’. This type is completely sealed and isused 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 similartype 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, whichform a circle slightly smaller than the bolt or stud diameter. As the nut is turned, afriction load is imparted onto the threaded device.

Philidas: This nut has a circular crown which is slotted horizontally in twoplaces The thread on the slotted part is slightly ‘out of phase’ with the rest of thethread, so that increased friction is achieved when the nut is turned.

Aerotight: Similar to the Philidas in appearance, except that the slots arevertical. Its locking method is also similar.

Lightweight: The locking section of this stiffnut is slightly oval in shape and socauses increased friction when the thread passes through it.Note: Metal hexagonal type stiffnuts may be re-used, provided they are not beingused in vital areas such as flying controls and they retain their friction effect. Arecognised rule for serviceability is that they are discarded when they can bescrewed all the way down, on a new bolt, using only the fingers.

Anchor nuts and Stripnuts: Anchor nuts are supplied with single or doubleattachment 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 itmay be an assembly, comprising stiffnut, cage and base plate.

Single attachment types are used in corners or where space is limited and havetwo adjacent fixing points. Double anchor nuts have a hole either side of thestiffnut. They are fitted to the structure by riveting.

Where a number of anchor nuts are required, to secure panels etc. a number ofstiffnuts may be fitted into metal strips for ease of securing. Stripnuts are usuallyof the floating variety.

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Stiffnuts and Anchor Nuts

Fig. 48

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12.8 SCREWSScrews are, probably, the most commonly used threaded fastener in aircraftconstruction. They differ from bolts in that they are generally made from lower-strength materials. They can be fastened by a variety of tools, includingscrewdrivers, spanners and Allen keys. Most screws are threaded along theircomplete 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 widerange of tasks. It is common sense that great care must be taken to replacescrews with the correct items, by using the markings on the screw, the IPC andany other systems in current use within the supply department, to protect againstincorrect screws being installed.

Another point, requiring care, is the difference in terminology between the Britishand 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 maindifference between aircraft bolts and machine screws, is that the threads of amachine screw usually run the length of the shank, whereas bolts usually have anunthreaded 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 goodholding properties on thin metal sheets.

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A Selection of Machine ScrewsFig. 49

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12.8.2 STRUCTURAL SCREWS

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

100°

Grip Grip

Length

Diameter Diameter

Structural Screw TerminologyFig. 50

12.8.3 SELF-TAPPING SCREWS

Self-tapping screws (refer to Fig. 51) have coarse threads and are used to holdthin 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 finerthan 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 HeadsFig. 51

12.9 STUDSStuds are metal rods that are threaded at both ends (refer to Fig. 52). In generalthey are used where it is not possible, or desirable for a bolt to be used. Likemany screw types of fastener, most studs are produced in a standard form, withvariants used for special purposes. For example, where a standard type isunsuitable, such as when being used in a soft metal, then a stepped stud (whichhas a greater holding power) would be used. A stepped stud would also be usedwhere 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 betweenits 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 preventthe nut from slackening will be negated.

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Typical StudsFig. 52

12.9.1 STANDARD STUDS

By far the most widely used stud is the standard (plain, or parallel) type, in whichthe diameter of the whole stud, along its length, is constant. Standard studs areclassified 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 thatrequire special consideration.

To meet special requirements, the various types of standard studs may also besupplied with non-standard lengths of plain portion and ‘metal’ end. A simplemethod 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. Thestud can then be screwed into or removed from the material. Breaking the cinchthen 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, isof paramount importance. The diameter of the plain portion of the stud is reducedto the minor diameter of the end threads, thus lightening the stud withoutimpairing 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’ endof the stud has to be housed in soft metal. The thread of the ‘metal’ end is onesize larger than that of the ‘nut’ end. For example, a ¼ inch BSF stepped studhas 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 thetapped stud-hole has to be re-drilled and tapped with a larger thread, due todamage.

12.9.4 SHOULDERED STUDSThis 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 betweenthe ‘metal’ end of the thread and the normal diameter plain portion. This shoulderseats firmly on the surface of the ‘metal’ and gives additional resistance tosideways stresses. The clearance hole in the second component, through whichthe ‘nut’ end and plain portion of the stud passes, must be machined at the innerend to give clearance to the stud shoulder.

12.10 THREAD INSERTSThread inserts are a means of providing a stronger anchorage, for bolts, screwsor studs, in the comparatively softer metal alloys(aluminium, magnesium,bronze), wood, plastics or composite materials. They may also be used when it isnecessary 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 designsof each type will vary according to the many manufacturers or to the environmentin which the fastener must operate.

12.10.1 WIRE THREAD INSERTS

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

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12.10.2 THIN WALL INSERTS

Thin wall inserts appear in a variety of designs, materials and surface finishesand consist of a thin tube, which is threaded internally and may, or may not, bethreaded externally. Similarly, special tools are required from the manufacturer toprepare the holes for the inserts and various methods are adopted to secure eachparticular type of thin wall insert into its hole. Thin Wall inserts include:

Key-Locked Inserts: Key-Iocked inserts are threaded both internally andexternally and, after being screwed into the prepared hole, are (as their nameimplies), locked into their holes by tiny wedges or keys. The keys are thenpressed (or hammered) into place between the insert and the wall of the hole.

Swaged Inserts: Swaged inserts are also threaded internally and externallyand are, again, screwed into the hole before a tool is used to deform (swage) theinsert 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 beinstalled, (after the insert) and yet another special tool is used to complete thelocking action of the lock-ring.

Bonded Inserts: Bonded inserts are, usually, only internally threaded (to holdthe bolt, screw, stud etc) and are secured in the prepared hole by the use ofadhesives.

Obviously, from this information, it can be seen that great care must be taken toensure that only the approved types of inserts are used in aerospace componentsand that the procedures for their installation and removal (laid out in the relevantManuals) are carefully followed.

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12.11 DOWELS AND PINSDowels 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 appropriatematerial which are fixed (often permanently) in one of the components of a jointsuch that the protruding shank of the dowel locates with a corresponding hole inthe item being attached, thus ensuring accurate assembly.Two examples of the use of dowels may be found where a Propeller Control Unitis attached to an engine casing and there is a requirement for absolute accuracyin the alignment of the oil tubes and, again, where the segments of an enginecompressor need to be joined with precision so that the rotating members do notfoul 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 toprovide 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 incompletecylindrical 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. Toremove a roll pin it must be driven from the hole with a correct-sized punch.

Roll PinFig. 53

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12.11.3 CLEVIS PINS

Clevis or flat-head pins (refer to Fig. 54) are used for hinge pins in some aircraftcontrol systems. They are made of cadmium-plated steel and have grip lengths in1/16-inch increments. When a clevis pin is installed, a plain washer is usuallyplaced over the end of the shank and a cotter (split) pin is inserted, through thepre-drilled hole in the clevis pin, to lock it in place.

Diameter

Length

Clevis PinFig. 54

12.11.4 TAPER PINS

Both the plain and threaded taper pins (refer to Fig. 55) have a taper of 1 in 48and are used in various locations during aircraft construction. They are designedto carry shear loads and are manufactured from high-tensile steel. The pins donot allow any loose motion or play and are used for joining tubes and attachingcollars to shafts.

The plain taper pin is forced into the hole, which is reamed to the specified sizewith a Taper Pin Reamer, and is held in place by friction alone. To ensuresecurity, it can also be wire locked in place, by passing the lock wire through thepre-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 endspeened, after being installed, to secure them in their holes.

The Threaded Pin is similar to the plain pin except that its small end is threadedto 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 muchlike a split pin, to prevent it loosening. These pins are sometimes referred to asbifurcated taper pins.

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

Plain Taper Pin

Threaded Taper Pin

Taper PinsFig. 55

12.12 LOCKING DEVICESThe problems associated with threaded devices, and the effects of vibration ontheir security, were discussed previously, when the use of stiffnuts and anchornuts was considered.

In addition to using methods which increase the friction between threads, thereare 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 someinstances (particularly with light alloy assemblies), spring washers are assembledwith plain facing washers between the spring washer and the component. This isdone to prevent damage to the surface finish when the spring washer iscompressed although, with steel assemblies, the plain washer is usually omitted.

It is good practice to renew spring washers during overhaul or repair. Thisprocedure is most essential in engines and engine components as well as whereunits have reciprocating parts; such as in compressors or pumps.

In normal circumstances, however, spring washers can be re-used if they haveretained their ‘springiness’ and ‘sharpness’. Types of spring washers include:

Single and Double Coil Washers: Manufactured from rectangular-sectionedsteel sheet and formed into a portion of a helix, the single and double coil are themost common types of spring washer to be found on aircraft components

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Single Coil Double Coil Crinkle Cup

Spring WashersFig. 56

Crinkle Washers: Crinkle washers are usually manufactured from eithercopper alloy or corrosion- resistant steel. They are often used in lightly loadedapplications such as instruments and electrical installations.

Cup Washers: Cup (or Belleville) washers are manufactured from spring steeland 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 ofthe nut to lose torque. Assembly should always be in accordance with themanufacturer’s instructions.

12.12.2 SHAKE-PROOF WASHERS

Flat washers of this type (refer to Fig. 57), are manufactured from steel orphosphor bronze and are used in place of spring washers. In somecircumstances conical shake-proof washers are used for locking countersunkscrews.

Either the internal or the external diameters can be serrated, the serration beingdesigned to bite into the component and nut to prevent rotation.

All shake-proof washers should be used only ONCE. It is rare for these washersto be specified in assemblies where an anti-corrosion treatment of thecomponents has been specified, as this could damage the treatment.

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Shake-proof WashersFig. 57

12.12.3 TAB WASHERS

Tab washers (refer to Fig. 58), are normally used on plain nuts. The washers aremanufactured from thin metallic sheet material and have two or more tabsprojecting from the external diameter. They can also be designed for locking twoor more nuts.

When the washer is installed, one tab is bent against the component or insertedinto a hole provided, whilst a second tab is bent against the flat (or flats), of thenut, after it has been torqued down correctly.Note: Multi-tab washers can be re-used until all tabs have been used once.

Tab WashersFig. 58

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12.12.4 LOCK PLATES

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

Lock PlateFig. 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 plateusually has 12 facets to allow for this adjustment. The plate is then placed overthe 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 andunwinding the nut.

Note: A Tab washer could be used to do the same task. The lock plate is usedwhere the nut is frequently removed - the plate can be used indefinitelyproviding 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 cadmium-plated carbon steel or from corrosion-resistant steel. Their primary purpose is tolock slotted and castellated nuts as well as for securing clevis pins. The nuts arelocked onto their bolts by passing the pin through the hole in the bolt and the nutcastellations. The legs of the pin are spread in one of two methods.

Whilst either of these methods will secure the nut to the bolt, differentairworthiness authorities prefer one method to the other.The pins are measured by diameter and length. It must be noted that the nutsmust never be over-torqued to get the holes into line. The nut must either bebacked-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.

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

Two Methods of Securing Split PinsFig. 60

12.13 LOCKING WIREWire-locking (or ‘Safetying’ as it is known in the USA), is the commonest form oflocking in use throughout the aircraft industry. The wire is usually made ofcorrosion-resistant steel or heat-resistant nickel alloy. Fine copper wire is alsoused for some special locking operations.

The wire is normally classified by its diameter in increments of ‘Standard WireGauge’ (SWG) or ‘American Wire Gauge’ (AWG). The most usual gauge used is22 SWG (or its American equivalent), although great care must be taken to checkthe 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 theappropriate 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 lockedwill be restrained from turning in a loosening direction.

There should be approximately eight turns to every 25.4 mm (1 in) length of wireand no length of more than 9.5 mm (3/8 in) should be left untwisted. The angle ofpull, or approach (refer to Fig. 61), should be not less than 45 to the rotationalaxis.

When the wire has been passed through the last hole, the wire must be pulledtight and the twisting continued for at least 12 mm - 13 mm (½ in). The wire isthen cut and the end doubled under, to prevent personnel getting ‘snagged’ orbadly cut.

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Wire-locking AnglesFig. 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 theirrespective 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 attemptto slacken off, will add tension to the wire.

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

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12.13.1USE OF LOCKING WIRE WITH TURNBUCKLES

As with any threaded fastener, turnbuckles must be locked to prevent them fromcoming loose and jeopardising the control runs they are connecting.

There are a number of different types of wire-locking used on turnbuckles and theAMM must be consulted to find which method is specified. Methods used includethe single wrap and single wrap spiral as well as the double wrap and doublewrap spiral.

The single wrap and single wrap spiral use a single strand of the appropriate wirethat passes through the hole in the centre of the turnbuckle, finishing up wrappedaround each end. The single wrap spiral also uses a single piece of wire that isspiralled 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 twosingle wraps, one in each direction. A double wrap spiral consists of two singlewrap 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 tabsand the wire are in complete alignment (refer to Fig. 63). Whenever possible, theclosed end of the wire should be in the tab and the twisted end at the componentto be locked, although the exact method may be found in the AMM.

Locking Wire and Locking TabsFig. 63

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12.13.3 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 certaincritical 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 abroken wire indicates that the switch or control has been operated. This methodis employed on systems where it is necessary to know when a system has beenoperated (such as in a Fire Protection system).

12.14 QUICK-RELEASE FASTENERSSpecial fasteners have been designed to hold fairings, cowlings and inspectionpanels in position and to allow their rapid removal and replacement duringservicing.

12.14.1 DZUS FASTENERS

Cowling and other inspection access doors will usually be found with Dzusfasteners, that can be locked and unlocked by a quarter turn of the stud (refer toFig. 64).

These fasteners consist of a hard spring-steel wire, which is riveted across anopening on a fixed part of the airframe. The stud is mounted onto the panel (orremovable part), using a metal grommet.

When the panel is closed, a quarter turn of the stud pulls the wire into the curvedslot of the stud, securing the panel to the airframe.

Panels (and cowlings) usually have a number of fasteners installed to ensure fullsecurity and, to indicate that all fasteners are correctly secured, the cowling willhave a series of lines marked (painted) on the surface.

When the studs are correctly fastened, then their screwdriver slots will be in-linewith the lines marked on the surface of the panels.

Some Dzus fasteners have a built-in receptacle, which guides the legs of the studonto the wire, to facilitate correct engagement.

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Dzus Fastener ConstructionFig. 64

12.14.2 ODDIE FASTENERS

Oddie fasteners (refer to Fig. 65) have a central stud, which is held in position inthe panel with a rubber washer or a coiled spring. A two-legged clip is fastened tothe fixed component (usually with rivets). The stud is bullet-shaped and has tworecesses 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 screwdriverslot is in line with marks on the panel. There should be a definite click as thefastener engages. A quarter turn of the stud will release it from the spring, andfree the panel.

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Oddie Fastener ConstructionFig. 65

12.14.3 CAMLOC FASTENERS

Camloc fasteners (refer to Fig. 66) consist of a spring-loaded stud assembly anda receptacle. The stud assembly is fastened to the removable panel whilst thereceptacle is fastened to the airframe.

To lock the fastener, the stud is pushed against its spring with a screwdriver andgiven a quarter of a turn clock-wise. As a result, the cross-pin, on the stud, ridesup 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 thecam. The fastener is unlocked by a quarter turn anti-clockwise when the studspring causes the stud to snap outwards.

Camloc FastenerFig. 66

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12.14.4 AIRLOC FASTENERS

Airloc fasteners (refer to Fig. 67) consist of a stud with a cross-pin in theremovable 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 dropsinto an indentation in the receptacle and holds the fastener locked.

Cross Pin

InstalledPin

StudsStud

Receptical

Airloc FastenerFig. 67

12.14.5 PIP-PINS

Quick-release ‘Pip-pins’ are used in assemblies where it is necessary to rapidlyremove or reposition components. They usually take the place of morepermanent bolts.

The ‘pip-pin’ quick-release fastener (refer to Fig. 68) operates on a push-pullprinciple. It consists of a hollow body containing a spring-loaded plunger. Whenthe pin is pushed into a hole, two steel locking balls, held in the shank of the pin,

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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 areforced to protrude from the shank, as the spring around the plunger expands, andso lock the pin in position.

A pip-pin is removed by a simple pull on the ring. This action aligns the groove inthe 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 beseparated at regular intervals and also require a hinging action. An example ofwhere pip-pins would be required is on engine cowlings. These have to beopened daily to allow for engine inspection, and are removed completely forengine changes.

Pip Pin Locking Balls

Pin Release Ring

Typical Pip-PinFig. 68

12.14.6 CIRCLIPS AND LOCKING RINGS

Circlips and locking rings (refer to Fig. 69) are manufactured from spring sheetmetal or spring steel wire, They may also be specially designed for a particularpurpose. Hardened and tempered to give either and ‘inward’ or ‘outward’ spring,they can be used for locking several parts together, locating components withinbores 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 ringsusually have one bent end that is inserted into a radial hole, drilled through thecomponent, which matches an inner or outer ring.

All circlips are subject to some damage at times and it will usually be arequirement, after they have been removed, to inspect them thoroughly. Any thatshow damage or corrosion should be discarded, although it is usual practice todiscard the wire type circlips whenever they are removed

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Internal External Internal External

Spring Sheet Plate

Internal External

Spring Wire Type

Squeeze legs together,or expand, to remove therings

Circlips and Locking RingsFig. 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 cutinto the respective components, to receive the key), is used to transmit the drivingforce from one part to the other.

There are different types of keys and keyways, and these will be covered ingreater depth in the section on transmissions.

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12.14.8 PEENING

Peening (refer to Fig. 70) is a method of preventing a threaded device (bolt, nutor screw), becoming loose by distorting the end of the thread, after installing thedevice. The distortion is normally achieved (using a centre punch) by striking thethread of the bolt or screw where it emerges from the threaded device, thusjamming and effectively locking the threaded device and preventing it fromloosening.

When using a nut and bolt combination, then one and a half threads of the boltmust protrude from the nut in order to create an effective peening.

The disadvantage of peening (and the distortion of the thread) means that, oncethe joint is dismantled, then the threaded device is useless and can only bediscarded.

Peened (Burred) Metal Peened (Burred)into Slot of Fastener

Fastener

Peened (Burred) FastenersFig 70

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12.15 GLUE/ADHESIVE BONDED JOINTSAs 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, ormore, materials together. The materials can be any of the large variety of fabricsfound in the aerospace industry (metal, paper, plastic, rubber or wood).

Some advantages of using adhesives, to make joints, are that the materials beingjoined may or may not be similar and the joints can be made proof against theleakage of gases and liquids.

Adhesives are normally good electrical insulators, which can greatly reducedissimilar corrosion on metal joints, and are not, normally, affected bytemperature changes.

Joining with adhesives not only saves the weight (and costs) associated withthreaded fasteners (and rivets), but also eliminates the need to make holes in thestructure, for those fasteners, which avoids the possibility of potential stressraisers.

The absence of fasteners in an aircraft’s skin results in a smoother airflow aroundthe aircraft, and thus contributes to its aerodynamic efficiency.

Adhesive bonded joints also provide greater stiffening to the structure, comparedto that achieved with mechanical fastenings.

There are, however, some disadvantages in that the surfaces, of the items to bestuck together (the adherends), must be free from grease, oil or dust, and thetype of adhesive must be suitable for the conditions or environment in which it isintended to be placed.

Fumes from adhesives can be narcotic, toxic and extremely flammable, so thatgreat care must be taken when applying adhesives. This entails working in well-ventilated conditions, wearing the appropriate personal protective equipment andobserving the relevant safety precautions to prevent (and, if necessary, fight) theoutbreak of fire.

12.15.1 LOCKING BY ADHESIVES

Applying Shellac, Araldite etc to DTD 900 specification, may be used to lockmany small components, particularly those in instruments, valves, switches etc.Adhesive is applied to the outside of the nut face and the protruding screwthread, or to the component and screw head, after tightening, and preventsmovement between relevant parts.

It is good practice, when using Araldite, to mix a separate sample under similarconditions, to check that it hardens within the specified time period.

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12.15.2 LOCTITE

Loctite is the trade name for a liquid sealant, used to lock metal threads. It is anapproved, proprietary material, which hardens in the screw threads afterassembly. Loctite is supplied in various grades to give a predetermined lockingstrength in a variety of applications from stud locking to retaining bearinghousings.

When using Loctite, it is advisable to degrease the parts to achieve maximumstrength. If the threads are not degreased, about 15% of the locking strength isnormally lost. Loctite should only be used when specified by the approveddrawings or instructions, and applied in accordance with the manufacturer’sdirections

12.15.3 SYNTHETIC RESIN ADHESIVES

Synthetic resin adhesives are used extensively for joints in wooden structures, toavoid the localised stresses and strains, which may be set up, following the useof mechanical methods of attachment.

Synthetic resin adhesives, used for gluing aircraft structural assemblies, mustcomply with the requirement prescribed in an acceptable specification

Synthetic resin adhesives usually consist of two separate parts, namely the resinand the hardener. The resin develops its adhesive properties only as a result of achemical 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 componentsfrom each batch. Where off-cuts are not available, tests should be done onrepresentative test pieces.

12.16 METAL-TO-METAL BONDED JOINTSMetal-to-metal joints involve the use of heat, to raise the temperature of themetals to a point where, either by the use of hammering, by the application ofpressure, or by a chemical reaction between the metals being joined, the metalsfuse 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 ormore pieces of metal (and applying a filler rod if required), making one continuouspiece.

Welded joints are normally considered to be part of an aircraft’s permanentstructure and they would not be dismantled during routine maintenance.

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Only a trained welder, authorised by the CAA, may weld component parts of aBritish-registered civil aircraft and that person is required to submit, to the CAA, aseries of test welds, for examination, every twelve months. It is, therefore, beyondthe 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 ata temperature considerably lower than the metals being joined. The filler metal isan 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 thesurfaces, being joined and, to allow this to happen, the joint surfaces must be freeof oil, grease, dust, and corrosion.

It is also necessary to use of an approved substance (a ‘flux’), which is applied tothe metals, to prevent the formation of potentially corrosive oxide films while themetals 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, thefillers melt at higher temperatures than soft solder and provide a much strongerjoint, which is also capable of operating at higher temperatures.

Silver Solder consists of an alloy of copper and silver (with a melting point almosttwice that of the soft solders) while Brazing uses a copper-zinc alloy with amelting point higher than that of Silver Solder.

The source of heat used for hard soldering is, usually, a direct flame and adifferent 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 closelycontrolled.

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 emerycloth, wire brush or filing.

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13 AIRCRAFT RIVETS

An aircraft, even though made of the best materials and strongest parts, would beof doubtful value unless those parts were firmly held together. Several methodsare used to hold parts together; welding or soldering, threaded fasteners andriveting being three of the main methods. The use of threaded fasteners, andsoldering, has been mentioned previously.

Rivets are an alternative method of fastening structure, a rivet being a metal pinon which a head is formed, during manufacture. The rivet is inserted into a pre-drilled hole and the plain end of its shank is deformed (‘set’ or ‘closed’) by the useof a hand- or power-tool.

Rivets create a joint at least as strong as the material that is being joined. Rivetsare normally strong in shear, but they should not be subjected to excessivetensile loads.

There are two main categories of rivet:

Solid rivets: which are ‘set’ using a riveting gun on the manufactured head anda reaction (bucking) bar on the remote side

Blind rivets: which may be installed where access is restricted to the shankend of the rivet.

13.1 SOLID RIVETSThere are a number of different types of rivet head, the most common being themushroom and round heads (refer to Fig. 71). Both of these rivets project abovethe 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 canbe used internally, when a flat head will make closing the rivet easier.

Length

Diameter Mushroom Pan Round

Types of Solid RivetsFig. 71

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The majority of aircraft rivets are manufactured from aluminium alloys. Rivets canalso be made from other materials such as steel, Monel metal, titanium or copper.

Material specifications for British and American rivets are not identical. Themanufacturer’s publications (AMM or CMM) will give details on which rivets canbe 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 countersinkangle, if applicable) and the material from which the rivet is made. This latterrequirement involves many different identifying marks and letters.

13.2 RIVET IDENTIFICATIONThe identification of solid rivets covers a multitude of marks and letters thatindicate not only the material, but also the heat treatment, (if any), that the rivethas gone through.

The American rivets are, usually, ‘natural’ (gold) or grey in colour and have headmarkings, whilst British rivets, generally, use a combination of colour andalpha/numeric codes.

13.2.1 SOLID RIVETS (BRITISH)

Standards for British Solid rivets are issued by the Society of British AerospaceSBAC (AS series) or the British Standards Institute (SP series). The standardsoverlap to a certain extent, with obsolete rivets, in the AS range, being replacedby 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 standardnumber identifies the head shape, material and finish. This is followed by a threeor four figure code, the first one or two figures indicating the shank diameter inthirty-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 ithas 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 varioustypes of AS rivets. Many of these rivets are obsolescent and have beensuperseded by rivets conforming to SP standards.

Table 15 gives details of material and identification information for SP rivets withthe standard numbers shown in Table 16. SP rivets are also available in metricsizes.

Note: The colour coding (of both British systems) of solid rivets is generally thesame 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, Monelrivets are natural and 5% magnesium rivets are green. This way of coding allowsmaterial types to be more easily identified.

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Table 13MATERIAL IDENTIFICATION OF AS RIVETS

Matl. Spec. Material Type Identification Marks FinishL37 Dural ‘D’ on shank end NaturalL58 Al. Alloy ‘X’ on shank end Dyed or Anodised

(5% Mg.) GreenL86 Hidiminium ‘S’ on shank end Dyed VioletDTD204 Monel ‘M’ on shank end Natural or Cadmium

Plated

Table 14TYPICAL SPECIFICATION NUMBERS OF AS RIVETS

Material Snap Mush 90º Csk 100º 120º 90º CloseSpec. Csk Csk Tol.L37 AS156 AS158 AS161 - AS164 AS2918L58 AS157 AS159 AS162 AS4716 AS165 -L86 AS2227 AS2228 AS229 - AS2230 AS3362DTD204 - - AS5462 - AS465 -

Table 15MATERIAL IDENTIFICATION OF SP RIVETS

Material. Material Type Identification Marks FinishSpec. (On shank end)L36 Aluminium ‘I’ Black AnodicL37 Dural ‘7’ NaturalL58 Al. Alloy ‘8’ Green Anodic

(5% Mg.)L86 Hidiminium ‘0’ VioletBS1109 Steel - CadmiumDTD204 Monel ‘M’ Natural or

Cadmium

Table 16TYPICAL SPECIFICATION NUMBERS OF SP RIVETS

Material Spec. Snap Head Mushroom Head 100º Csk HeadL36 SP77 - SP68L37 SP78 SP83 SP69L58 SP79 SP84 SP70L86 SP80 SP85 SP71BS1109 SP76 - SP86DTD204 SP81 - SP87

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13.2.4 SOLID RIVETS (AMERICAN)

These are generally used in normal construction and repair work. They areidentified 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) areRoundhead, Brazier head, 100º Countersunk head, Flat head and Universalhead.

AN Rivet Head Types

Plain Dimple Raised Dot 2 Raised Dashes Cross

A AD D DD B

AN Material Identification and Code LettersFig. 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 aremerely used to illustrate a typical coding system. The other most commonstandard for American rivets is the MS (Military Standards) system which, whilsthaving slight differences from the AN system, uses similar terminology todescribe 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 isfollowed by another figure expressing the diameter of the rivet shank in 32nds ofan 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 aUniversal head, aluminium alloy (2117-T) rivet, of 3/32 inch diameter with a shanklength 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 alsoused 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 anda 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 isaluminium alloy. Digits and letters identify the degree of temper condition, ofaluminium 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 wouldbe used for riveting lightweight, soft, aluminium structures, where strength is not afactor.

The 2117-T (AD) rivet is made from aluminium alloy and (as has previously beenmentioned) is known as the ‘field’ rivet. It is the most commonly used rivet, mainlybecause it is ready to use as received and needs no further heat-treatment. Italso has a high resistance to corrosion.

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The 2017-T (D) and 2024-T (DD) rivets are made from high strength heat-treatable aluminium alloys. They are used where more strength is required thanthat obtained from the ‘field’ rivet. The rivets need to be heat-treated and, if notrequired immediately, they should be refrigerated until needed. The 2017-T rivetshould be driven within 1 hour of removal from refrigeration (or following heat-treatment) and the 2024-T must be driven within 10-20 minutes.

The 5056 (B) rivet is used for riveting magnesium alloy structures, because of itsgalvanic compatibility with magnesium (to reduce the risk of corrosion).

Mild Steel rivets are used for riveting steel parts while Corrosion Resistant Steelrivets 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 rivetmanufactured from mild steel.

Monel (M) rivets are used for riveting nickel-steel alloys. They may also be usedas a substitute for CRS rivets when specified.

Copper (C) rivets are also available, but their use is limited on aircraft. They mayonly be used on copper alloys or non-metallic materials, such as leather.

Note: Most metals, including aircraft rivets, are subject to corrosion. This may bethe result of local climatic conditions or the fabrication process used. It can bereduced to a minimum by using the correct materials and by the use of protectivecoatings on the structure and the rivets. The use of dissimilar metals should beavoided where possible and, as previously stated, the rivet manufacturers usuallyapply a protective coating on the rivets, which may be either of a zinc chromate, ametal spray or an anodic film finish.

13.3 HEAT-TREATMENT/REFRIGERATION OF SOLID RIVETSThe action of closing a rivet, and the strength required on completion, dictateswhether any heat-treatment will be required prior to closing. As previouslydiscussed, some rivets, for non-structural applications, can be manufactured frompure aluminium. These are given no heat-treatment and are soft, both before andafter 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 closewhilst possessing adequate strength.

Where rivets of a stronger material are required, then ‘D’ and ‘DD’ rivets can beused. These are also made from aluminium alloys, but to different(AN)specifications. They are heat-treated, just prior to use, and either formed within ashort time period of time (in which they ‘age-harden’), or they are stored, in a

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refrigerator, at temperatures well below zero degrees Celsius (which retards theage-hardening process), until required for use. They are known as ‘icebox’ rivetsin the USA.

13.3.1 HEAT-TREATMENT.

Metal temper is important in the riveting process, especially with aluminium alloyrivets. These generally have the same heat-treating characteristics as sheetalloys and can be annealed and hardened in much the same manner. The rivetmust 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 andthen they harden with age.

The process of heat-treatment of rivets (normalising) may be achieved in eitheran electric, air furnace or in a salt bath. The temperature range, depending on thealloy, is in the region of 495ºC - 505ºC. For convenient handling, the rivets areheated 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, ifthe rivets are not to be set immediately, they may be refrigerated to delay theage-hardening process. The solution-treated rivets are stored at low temperature(below freezing) and, under these conditions, will remain soft enough for drivingfor up to 2 weeks. Any rivets not used in that period should be removed and re-heat treated.

It should be noted that refrigeration only delays age-hardening and that age-hardening will continue at a rapid rate as soon as the rivets are removed from therefrigerator.

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 ofan 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 thinsheet. It offers only a slight resistance to airflow and is often used on exteriorskins, especially on aft sections of fuselage and empennage. A modified brazierhead 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 sothat it can be installed into a countersunk or dimpled hole and so be flush with thematerial’s surface. The countersunk angle may vary from 78º to 120º (the 100ºrivet being the most common type). Countersunk rivets are used to fasten metalsheets which overlap others. They are also used on exterior surfaces of theaircraft, because they offer only a slight resistance to airflow and thereforeminimise turbulence.

Flathead: used on interior structures, where there is insufficient clearance touse a roundhead rivet.

Roundhead: used in the interior of the aircraft and has a deep rounded topsection. The head is large enough to strengthen the sheet around the hole and tooffer 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 RIVETSThere are many places in an aircraft where access to both sides of the structureis 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 asaircraft interior furnishings, flooring material, de-icer boots etc, the full strength ofsolid shank rivets may not be necessary. For use in such places, special rivetshave been designed which can be set from one side only.

These rivets are often lighter than solid rivets, yet amply strong enough for theirintended use. The rivets are produced by several manufacturers, and haveunique characteristics requiring special installation tools and procedures. Thesame, general, basic information, relating to their fabrication, composition, uses,selection, installation, inspection and removal procedures applies to most ofthem.

Hollow rivets that can be closed by pulling a mandrel through them are oftenknown as ‘blind’ rivets and these in turn can be described as MechanicallyExpanded 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, tothe lack of access to the both sides of the joint), they must, in the absence ofspecific instructions, be of the same material as the original solid rivet, and be ofequivalent shear strength. The shear strength, of the rivet, may be increased, byusing a form of ‘plug’ to fill the hollow shank of the rivet.

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13.4.1 FRICTION LOCK RIVETS

These are generally fabricated in two parts, consisting of a rivet head with ahollow shank and a stem that extends through the hollow shank. They may,typically, be of the ‘friction lock’ protruding head or countersunk head styles ofrivet (refer to Fig. 73). Several events occur in sequence when a pulling force isapplied 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 tofracture at the weakest point. The bottom end of the stem is retained in theshank, giving much greater shear strength than could be obtained from a hollowrivet.

Note: With this type of rivet, the stem is often designed to break above the rivethead, necessitating a further action, which entails cutting off the extra portion ofthe stem with snips (or a specialised pneumatic gun) and milling the exposedportion flush with the head. This type of rivet is going out of style because of theextra work involved with setting it.

Friction Lock RivetsFig. 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 frictionlock rivet previously described, except in the manner in which the mandrel isretained in the rivet.

This type of rivet has a positive mechanical locking collar, to resist the vibrationsthat 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, inmost cases, can be substituted rivet for rivet. Three operations are performedwhen the rivet is installed (generally using a pneumatic gun):

When pulling force is exerted on the stem, the stem is pulled in, forming theblind head and clamping the sheets of metal together.

At a pre-determined point, the inner anvil, incorporated in the gun, forces thelocking collar into position.

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

Mechanical Lock RivetsFig. 74

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13.4.3 HOLLOW/PULL-THROUGH RIVETS

When installed, the rivet mandrel is pulled through these rivets, leaving a hollowrivet of much lower strength than the self-plugging types.

Different types of these rivets are supplied, either complete with individualmandrels or as individual rivets, used with a re-usable steel mandrel, which isdrawn completely through the rivets. In some cases, the rivets may be pluggedwith sealing pins which, as previously stated, give them additional strength aswell as sealing them.

13.4.4 GRIP RANGE

Unlike a solid rivet, the part of a blind rivet, available to form a head, cannotalways be seen. It is, therefore, necessary to know the range of total materialthickness that a given rivet can fasten together.

This is known as the ‘Grip Range’ of the rivet and requires the use of a gauge tomeasure the material thickness (refer to Fig. 75), which is used in conjunctionwith a rivet data table.

1 3 5 73

Rivet Group to beUsed = 4

Grip Measuring GaugeFig. 75

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13.4.5 TUCKER ‘POP’ RIVETS

Tucker ‘Pop’ rivets (refer to Fig. 76) are supplied mounted on steel mandrels. Thehead of the mandrel is pulled into the rivet, expanding it, before the mandrelfractures at the waisted portion. This waisted portion may either be near to thehead of the rivet, or part way up the stem. In the first case the rivet will beclassified 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-pneumaticgun. ‘Pop’ rivets are less suitable for use on aircraft as they tend to loosen withvibration and then become increasingly difficult to remove, because of thelooseness and the presence of the steel mandrel. (They also tend to spin whenattempts are made to drill them out).

‘Pop’ RivetsFig. 76

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Break head rivets must not be used if the structure is not accessible to retrievethe mandrel heads. It is sometimes permitted for the mandrels of Break Stemrivets to be dipped in an adhesive, so that they will not vibrate loose afterinstallation. If Tucker ‘Pop’ rivets are to be used externally on aircraft, the headsmust be sealed to prevent the ingress of dirt and moisture. Cellulose MetallicFiller is often recommended for this purpose.

The rivets are manufactured in either aluminium alloy or cadmium-plated Monelmetal, with either dome heads or 100º and 120º countersunk heads. The AGSreference number consists of the AGS number identifying the material and headtype, a three figure size code and letters specifying Break Head or Break Stem. Inthe size code the first figure represents the diameter, in increments of 1/32 inchwhile 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 itsdiameter 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 totallyremoved 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 olderaircraft. To close the rivet, the stem is pulled through and, at a predeterminedload, the stem breaks proud of the manufactured head of the rivet, plugging therivet body. Whilst the stems can be milled off on alloy rivets, those manufacturedof stainless steel or titanium break flush with the rivet head. A flush finish isrequired 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 RivetsFig. 77

13.4.7 CHOBERT RIVETSChobert rivets (refer to Fig. 78) are also similar to Tucker ‘Pop’ rivets, but havea tapered bore. The head of the mandrel is re-usable, and is pulled fullythrough the rivet on forming. This gives an advantage of no loose articles afterthe riveting operation is completed. The mandrel is drawn through the rivetusing a special tool, which carries a number of rivets on the mandrel to allowrepetitive and faster riveting. The tool simply feeds the next rivet into placeafter the closure of the previous one.

Where additional shear strength or water-tightness is required, sealing pins orplugs of the same material are driven into the bore of the closed rivets

Chobert RivetsFig. 78

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13.4.8 CHERRY RIVETS

Cherry Rivets (refer to Fig. 79) consist of a range of fasteners including Cherry-Lok and Cherry-Max, which are manufactured in the USA. The primary differencebetween these and the rivets mentioned previously is that the mandrel is lockedin position, after closing, instead of depending on friction alone.

During the final stages of closing, a locking collar, located in a recess in the rivethead, is forced into a groove in the stem and prevents the stem from any furthermovement. This method means that, when closed, the rivets have a shear andbearing strength high enough to allow their use in place of solid-shank rivets.

Cherry RivetsFig. 79

13.5 MISCELLANEOUS FASTENERSThese fasteners are, basically, close-tolerance, metal pins that combine the bestfeatures of a rivet and bolt. They usually require access to both sides of the jointbut are extremely strong in shear, with a shear strength equal to a standard ANbolt 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 heat-treated steel) which has a thin, manufactured head at one end and a part-threaded shank at the other. The threaded end of the Hi-Lok fastener contains ahexagon-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 turnedonto the threaded shank by a box wrench or an ordinary spanner. An Allen Keyengages in the recess in the shank end, to prevent rotation of the pin whilst thecollar is being tightened and, when a pre-determined load is reached, thehexagonal section of the collar shears off, leaving the pin securely fastened in thehole. Because the collar breaks off at a designated pre-load, the use of torquewrenches 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 AfterAssembly

Collar Wrenching DeviceShears Off

Typical Collar

Collar Driving-HexPin Recess-Hex

Hi-Lok FastenerFig. 80

13.5.2 HI-TIGUE FASTENERSHi-Tigue fasteners (refer to Fig. 81) are similar to Hi-Loks, excepting that theypossess a bead at the bottom of the shank, adjacent to the threaded portion ofthe fastener. The bead exerts a radial load to the side of the hole which serves tostrengthen the area surrounding the fastener hole. This reduces the effect ofcyclic loads on the fastener which, in turn, will reduce the effect of the coldworking of the joint and minimise the likelihood of subsequent failure. Hi-Tiguefasteners are closed in exactly the same manner as the Hi-Lok types.

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Bead

Hi-Tigue FastenerFig. 81

13.5.3 HI-SHEAR FASTENERS

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

Hi-Shear Fastener

Fig. 82

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13.6 SPECIAL PURPOSE FASTENERSIn addition to the fasteners already described, other rivet-type fasteners are oftenused in the manufacture and repair of aircraft. While some of these are designedfor 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 wouldnormally 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 (whichmay be of a hexagonal or countersunk headed type), a hollow steel bolt and astainless steel sleeve.

The fastener is installed with either a pneumatic or a hand-operated tool, withwhich the bolt is rotated and the nut is held stationary. This action expands thesleeve over the tapered end of the nut and draws the fastened items together. Ata pre-determined torque, the bolt breaks off at a notch-weakened point, flush withthe head of the nut. A different tool is required for each of the two head forms andfor each particular diameter bolt.

Jo-Bolts and Installation SequenceFig. 83

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13.6.2 TUBULAR RIVETS.

Tubular rivets are used primarily to save weight when riveting through tubular orhollow 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 aremade to AGS drawing specifications in several materials. The drawing numberindicates the type of rivet and the following letter denotes the material. Thenumber after the letter denotes the dimensions of the rivet, but has no particularsignificance 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, hasa length of 1 inch, and has a wall thickness of 26 SWG. Table 17 shows theletters used to indicate different tubular rivet materials and the features by whichthe materials may be recognised.

Table 17IDENTIFICATION CODES FOR TUBULAR RIVETS

LetterMaterialIdentification

Identification Feature

Protective PhysicalTreatment Characteristic

A Aluminium (L54) Anodic film Dyed black

D Duralumin (L37) None Natural colour

H Mild steel (T26) Cadmium Magneticplated

J Nickel alloy Cadmium Only slightly(DTD268) or plated magneticMonel metal(DTD204A)

K Monel metal None Only slightly(DTD204A) magnetic

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13.6.3 RIVNUTS

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

RivnutsFig. 84

Installation is achieved by drilling a hole into the skin and a small notch made onthe 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), andthe key aligned with the notch. The puller handle is squeezed, closing the nut andgripping the skin. The tool is then unscrewed from the Rivnut, leaving a threadedhole 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, butto avoid confusion, only the American types are considered here.

These Rivnuts are available in six grip ranges, the minimum grip Rivnut having aplain head while the next size has a radial dash mark on the head. Eachsucceeding grip range is indicated by an additional radial mark on the head withthe largest size having five radial dash marks.

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Keyway

Rivnut InstallationFig. 85

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14 SPRINGS

The invention of the wheel, for transport (and many other applications), isconsidered to be one of the major advances of mankind, but another, less-

praised, technical, innovation followed the development of devices employed toalleviate the discomfort of travelling on unmade or rutted roads.

Using the fact that the elasticity, inherent in most materials, allows them to absorbenergy by distorting or deflecting when under load - and then, to return to theiroriginal shape after the load has moderated (or has been removed), - earlysprings consisted of flat (and curved) sections of wood (and, later, metal), towhich were attached the carriages of the respective eras.

The dawning of The Industrial Revolution led to the mechanisation of practicallyevery facet of civilised life, from the production of food and textiles to the miningand processing of minerals in order to provide many other materials and thevarious machines deemed necessary for sophisticated living conditions.

In addition there has followed huge advances in transport, time-keeping, world-wide communication and (inevitably) military capabilities, in all of which can befound 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 SPRINGSSprings 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 morecommon forms are included here for consideration.

14.2.1 FLAT SPRINGS

Flat springs, while they were a development of flat, rectangular-sectioned, stripsof metal, can actually be found in forms other than simply flat as, for instance, inthe shape of the springs which control the contact breaker points in the magnetoof 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 aircraftembodied leaf springs in their landing gear, this type of spring is more familiar inthe automobile and railway industries.

14.2.3 SPIRAL SPRINGS

Spiral springs may be found in the form of spirally wound flat springs (known asmotor or power springs) or as spirally wound wire, such as the hair springs ofmany types of instruments.

14.2.4 HELICAL COMPRESSION AND TENSION SPRINGS

These are the most commonly found springs, which superseded the leaf springwhen space and lightness of structure were the requirement. They are made in awide variety of materials and sizes and may be found in a seemingly endlessnumber of applications.

14.2.5 HELICAL TORSION SPRINGSWhile being similarly wound to the previous two types, these springs have

specially shaped ends, to permit a torque force to be applied, and transmitted, ina plane normal to the helix axis.

14.2.6 BELLEVILLE (CONED DISC) SPRINGSBelleville springs are, in fact, shaped like the Cup Washers, which were

previously discussed in the topic on Locking Devices. Belleville Springs arecapable 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 MANUFACTUREDThe materials, used for the manufacture of springs, cover a very wide range ofmetallic and non-metallic (plastic and elastomer) substances. These notes will,however, be confined mainly to the discussion of metallic types, with a smallconsideration 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 to18.42 mm (0.375 in to 0.725 in) certain steels are drawn into wires and cold-wound 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 thewinding processes. Typical types of carbon- and alloy-steel stock, used for themanufacture 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, assuch, would only be used in applications of low stress and not where fatigueloadings 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 foundwhere long fatigue life is not required

Music Wire: which is a carbon-steel stock of high quality and is suitable forsmall-sized, helical springs in applications involving high fatigue stresses

Chrome-Vanadium Steel Wire: which is a stock that has been used for piston-type aero-engine valve springs and is, therefore, suitable for high-temperatureand high-stress conditions

Chrome-Silicon Steel Wire: which, when used in valve springs, has a higherfatigue 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 inconditions where high corrosion-resistance is the requirement. This grade of wirewould also be utilised in applications where resistance to creep at elevatedtemperatures is desired. Some grades of Stainless-Steel wires can be made toaccept magnetism, where this characteristic is needed alongside the otherqualities.

14.3.2 STEELS USED FOR HOT-WOUND SPRINGSAbove the cross-sectional diameters, previously mentioned, it is considered

impractical to cold-wind and so, the larger diameter metals (bars or rods) are hot-wound and then also subjected to various stress-relieving processes.

Similar carbon- and alloy-steels to those already discussed are employed in themanufacture of hot-wound springs, with the necessary variations in their contentsof 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 beformed from oil-tempered steels (thin sections - clock-type springs) or fromannealed 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 electricalconductivity is required, and on nickel alloys where the ability to work at elevatedtemperatures is desirable, these alloys include:

Spring Brass: which is comparatively inexpensive, has good electricalconductivity, but is unsuitable for high-stress applications

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Nickel Silver (also called German Silver): which has better characteristics thanbrass and is, in fact, made from different percentages of copper, zinc and nickel

Phosphor Bronze: which has a minimum percentage of 90% copper contentand has, therefore, excellent electrical conductivity. It is suitable for applicationsof higher stress levels than those of brass

Silicon Bronze: which has similar characteristics to those of phosphor bronzebut 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%) impartinggreater hardness and other superior mechanical properties

High-Nickel Alloys: which are the types more commonly found in aero-engineapplications 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, infact, contain almost 50% iron. All of these non-ferrous alloys can be found in thecold-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 toform the anti-vibration mountings (Metalastic Bushes and Housings) such asthose found in aero-engine and auxiliary power unit (APU) installations

Others combine synthetic rubber strands, encased within a sheath of braidedcotton, nylon or similar materials. They are, usually, referred to as ‘ShockAbsorbers’ or ‘Shock Cords’ rather than ‘Springs’ and are more familiarly knownby the generic name of ‘Bungee Cords’. Bungee Cords may be encountered onmany light- and medium-sized aircraft while their use on heavier aircraft is notunknown.

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14.4 CHARACTERISTICS OF TYPICAL AEROSPACE SPRINGSIf 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 ofthe material and ignoring the effects of temperature and constant or repeatedloadings), 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 theload is doubled, then the deflection also doubles - a characteristic of which gooduse 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 incertain control and indicating functions of aircraft structures and components.

(a) (b)Helical Springs Belleville Springs

Load/Deflection Graphs for SpringsFig. 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. Lossof 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 hydraulicsystem. Low-pressure fluid lines can be manufactured from metal or plastic(pipes and tubes) or, alternatively, from various forms of rubber (hoses). High-pressure fluid line can be made from a variety of materials, including aluminiumalloy, stainless steel, copper, titanium and also reinforced flexible hoses.

Fluid lines are made of rigid, semi-rigid and flexible tubes, depending on theiruse. 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 wallthickness in comparison to rigid lines. A variety of end-fittings may be used tomake connections between semi-rigid tubes.

Flexible fluid lines are made from rubber or synthetic materials and are usuallycalled ‘hoses’. Depending on the pressure they are designed to carry, hoses mayhave reinforcing materials wrapped around them. Various types of end-fittings areused to attach hoses to each other and to other components.

15.1 RIGID PIPESThese are usually manufactured in a standardised combination of length, outsidediameter (OD), and wall thickness. The use of threads, cut into the pipe wall, andthe need for special end-fittings means that, apart from some components, thereare 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 toshape and are often flared for connectors. Sizing is also by length, OD and wallthickness.

Various methods are used to connect semi-rigid tubes both to each other and toother connectors. These will depend upon the use, location and pressure beingcarried 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 andnot compatible) on the end of the tube, which matches a cone of the same angleon 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 afluid-tight metal-to-metal seal. The end-fittings are produced in a wide variety oftypes, 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 topass fluids through structural portions (bulkheads) of an aircraft or of an enginepower-plant assembly.

In-line connectors may be either of the pipe-to-pipe or pipe-to-adapter type ofconnectors 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 ConnectorsFig. 87

Where it is necessary to have fuel, oil or other tubes passing through structuralbulkheads, it requires an end-fitting with a long body and provision for securing

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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 theextra machine thread at one end, for attachment, to the bulkhead, by anadditional, threaded, locking device.

Typical American (including ‘Bulkhead’) ConnectorsFig. 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 theseapplications the flare-less coupling is designed to provide leak-free attachmentswithout flares.

Although there is no need to flare the tube, in one of the methods used, it isnecessary to pre-set the coupling, prior to its installation (refer to Fig. 89). Pre-setting is the process of applying enough pressure to a sleeve (also called aferrule) 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

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the nut causes the ferrule to ‘bite’ into the tube. Depending on the size of the tubeand its material, between one and one and a half turns of the nut is enough toform the pre-set.

When complete, the tube can be inspected and, if satisfactory, attached directlyto the appropriate union or adapter.

Pre-set Flare-Less CouplingFig. 89

Two other methods of forming flare-less couplings involve the swaging of metalsleeves around the ends of the tubes, which are being connected and the joiningof tubes by brazing. Both methods require specialist skills, which are beyond thescope of these notes.

15.3 FLEXIBLE PIPES (HOSES)The need for flexibility in many areas of aircraft construction means it is oftennecessary to employ hoses, instead of semi-rigid tubing, for the transmission offluids and gases under pressure. Whilst a number of hoses were previouslymanufactured from rubber, most modern hose manufacturers use either Teflon orother 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 innerand outer tubes are made from synthetic rubber, with the inner having a braidedcotton reinforcement (refer to Fig. 90). These hoses are used on instrumentsystems, vacuum systems, autopilots and other low-pressure systems, usuallyoperating at pressures below 300 psi (2.07 x 10³ kN/m²).

Low-Pressure HoseFig. 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 astress-inducing twist in it. Other markings could include the hose manufacturer’scode 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 whilstsmaller diameter hoses will be able to withstand such pressures, larger sizes maybe restricted to lower pressures.

Typical construction of this type of hose could be a seamless inner liner madefrom different materials, a layer of cotton braid, a layer of stainless-steelreinforcement and an outer layer of tough, oil-resistant, rubber-impregnatedcotton.

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15.3.3 HIGH-PRESSURE HOSES

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

High-Pressure Hose AssemblyFig. 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 tubesand wire braids, so as to resist the high pressure twisting and vibrating loads, aswell as providing an electrical bond throughout the assembly.

Flexible hoses have their sizes identified by their inner bore diameter and theoverall length. With pre-assembled hoses, the overall length of the assembly,from the centres of the nipple extremities, regardless of the shape of the endfittings, is used for identification purposes (refer to Fig. 92).

Flexible hoses, used in engine bays and other high temperature areas, will oftenhave a metallic stainless braid as the outside layer, to make the hose fire-resistant.

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Effective EffectiveLength Length

Effective Length of Hose AssembliesFig. 92

15.4 UNIONS AND CONNECTORS

EffectiveLength

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 unionsand connectors that are available for rigid pipes and flexible hoses on aircraft.These may be of British, European or American manufacture with the differentstandards 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 standardcomprises 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 andadapters), sleeves, collars, and nipples.

The cones (flares) on AGS end-fittings (unions and adapters) have an includedangle 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, butit should be noted that the flares and other hardware for this standard have anincluded angle of 74º.

15.4.3 MILITARY STANDARD (MS)

This standard (as previously discussed) has replaced the standards from the ANsystem. Many AN part numbers have been incorporated into the MS system andnow appear with MS designations

Other specifications in current use with aircraft manufactured in the USA includeNational 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 MaterialsDivision of SAE specifications (AMS) are yet another set of standards to whichaerospace materials may be produced. The Society of Automotive Engineers hasa second standard - referred to as the Aeronautical Standard (AS) - which is forcomponents that do not qualify for an AMS standard.

All these specifications provide for a range of fasteners with Unified threads in theUNC, UNF and UNJF series and, whereas British aircraft fasteners aremanufactured in a selected range of Unified threads, American fasteners are insome instances supplied in both UNC and UNF threads.

From all this it can be seen that great care must be taken when matching upunion assemblies with these many different forms of thread.

15.5 QUICK-RELEASE COUPLINGSQuick-release couplings are required at various points in aircraft systems. Typicaluses are in fuel, oil, hydraulic and pneumatic systems. Their purpose is to savetime in the removal and replacement of components; to prevent the loss of fluidand to protect the fluid from contamination. The use of these couplings alsoreduces the maintenance cost for the system involved.

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

Typical Quick-Release CouplingFig. 93

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16 BEARINGS

Bearings are, broadly, classified by the type of rolling element used in theirconstruction. Ball bearings employ steel balls, which rotate in grooved raceways,whilst Roller bearings utilise cylindrical, tapered and spherical rollers running insuitably shaped raceways (refer to Fig. 94).

Although these notes give information on the uses of the various types of ball androller bearings, - together with general information on installation, maintenanceand inspection, - the Aircraft Maintenance Manual (AMM) should be the finalarbiter 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 engineand gearbox applications with rotational speeds in excess of approximately 100rpm. Most other bearings, on an aircraft or in an engine, are intended foroscillating or slow rotation conditions and do not have a cage. They are generallyshielded or sealed and pre-packed with grease, although some have externallubrication facilities.

Ball and Roller BearingsFig. 94

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

Radial Bearings

Angular-Contact Bearings

Thrust Bearings

Instrument Precision Bearings

16.1.1 RADIAL BEARINGSRadial bearings are the most common type of bearing and can be found in alltypes of transmission assemblies such as shafts, gears, control rods and endfittings. They are manufactured with either a single or double row of balls, rigid fornormal applications and self-aligning for positions where accurate alignmentcannot 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 inone direction only. The outer ring is recessed on one side to allow the ball andcage assembly to be installed, thus enabling more balls to be used and the cageto 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 angular-contact bearing may be used but, where they vary in direction, an opposed pair ofbearings may be used.

16.1.3 THRUST BEARINGS

Thrust bearings are designed for axial loading only. They will usually be found inuse together with roller or radial ball bearings. The balls are retained in a cageand run on flat or grooved washers. These bearings are adversely affected bycentrifugal 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 bothinstruments and communication equipment.

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16.2 ROLLER BEARINGSRoller bearings may be divided into three main types that define their use. Theyare:

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 ofthe 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 theirdiameter, although some rollers have a length greater than their diameter to caterfor special applications.

Roller bearings, which have a length much greater than their diameter, arenormally called needle roller bearings. These are designed for radial loads onlyand are best used in situations where the movement is oscillatory rather thanrotary, such as in universal joints and control rod ends.

16.2.2 SPHERICAL ROLLER BEARINGSSpherical 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 toaccept a small degree of misalignment. These bearings will accept high radialloads and moderate axial loads.

16.2.3 TAPERED ROLLER BEARINGS

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

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16.3 BEARING INTERNAL CLEARANCEDue to the heat, generated during operation, Radial Ball and Cylindrical Rollerbearings are manufactured with different amounts of internal clearance.

The bearings are produced in four grades (groups), and are usually marked insome 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 bearingsare 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 temperaturescould reduce the internal clearance and are not suitable as thrust bearings nor forhigh-speed situations.

16.3.2 NORMAL GROUP (‘TWO DOT’) BEARINGS

Normal Group bearings are used for most general applications, where only onering, of the bearing race, is an interference fit and where no appreciable amountof 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 bearingsand are employed where both race rings are interference fits, or where one ring isan interference fit, and some transfer of heat must be accepted. These bearingsare also used for high speed and in applications where axial loadings arepredominant.

16.3.4 GROUP 4 (‘FOUR DOT’) BEARINGS

Group 4 bearings have the greatest internal clearance and are found where bothrings are interference fits and where the transfer of heat reduces internalclearances.

Standard bearings are produced in all four groups while instrument precisionbearings are supplied only in the first three groups

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16.4 BEARING MAINTENANCEBall and roller bearings, if properly lubricated and installed, have a long life andrequire little attention. Bearing failures may have serious results and, to avoidsuch problems, Aircraft Maintenance Manuals and approved MaintenanceSchedules include full lubrication and inspection instructions, which MUST befollowed in order to limit the likelihood of bearing failure.

16.4.1 LUBRICATIONAs has already been stated, most bearings, used in airframe applications, areshielded (sealed) to prevent the entry of dirt or fluids, which could affect bearinglife. These cannot, normally, be re-greased and must be replaced if there aresigns of wear, loss of lubricant or brinelling. (brinelling is the indentation of thesurfaces of the bearing races).

In some places, where there is risk of loss of lubricant, a grease nipple will beprovided to permit recharging with fresh grease. Greasing should only be doneafter the nipple has been wiped clean of all dirt and, on completion, all excessgrease 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 beinspected regularly because, if corrosion or wear begins, the bearing willdeteriorate rapidly. Bearings are usually inspected without removing them fromthe component (in situ), as continued removal and installation of bearings cancause wear and damage.

Wheel bearings are inspected when the wheel is returned to the Wheel ServicingBay for maintenance. Other items might also be inspected when their majorassembly 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 ofinterference or fouling with (or from) adjacent components.

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17 TRANSMISSIONS

In mechanical engineering terms, transmissions consist of a series of connectedparts (or mechanisms) whereby a source of power can be applied to anothercomponent, 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 providethe power, which eventually results in the desired motion in a particular systemand 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, isachieved 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 afew situations where (lighter and less expensive) belts and pulleys are used totransmit movement/power in place of cables.

Nominally flat belts and pulleys use only friction to transmit the power from inputto output shafts. These are, unfortunately, prone to slippage so, to reduce theproblem, vee-section belts were devised and yet a further improvement has seenthe development of serrated or ‘toothed’ belts and pulleys, which use the principleof ‘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 inputand output shafts which are displaced by some distance.

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

Simple Belt and Pulley SystemFig. 95

Some uses of belt and pulley installations in aviation can include the driving ofpropellers on micro-light aircraft, which use high-revving engines. These enginesrotate about 6000 rpm whilst propellers are most efficient at around 2000 - 2500rpm. Therefore the drive from the crankshaft pulley, via a strong wide belt to thepropeller pulley, gives a step down ratio of about 2.5:1 on most of this type ofaircraft.

Another application of belt drives is on certain piston-engined helicopters, whichuse a belt to connect the output pulley on the end of the crankshaft to thetransmission and rotor. The tension pulleys, which bear onto the belt, keep it atthe correct tension for normal use.

When starting-up, the tension can be totally released, allowing the engine to bestarted without the load of the rotors and transmission. In an emergency thereleased tension allows the rotors to free-wheel (autorotate) and, thus, enables asafe landing.

There are a number of places inside piston engines where toothed belts, areused to drive camshafts and other accessories from the crankshaft.

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

Spur

Bevel

Key

Worm and Wheel

Basic Forms of GearsFig. 96

17.2.1 GEAR TRAINS AND GEAR RATIOS

A gear ‘train’ consists of two (or more) gear wheels, running in series, onseparate, parallel, shafts such that one gear transmits its drive to the other. Geartrains can change the direction of rotation and can also alter the speed of theoutput shaft. The speed of rotation is dependent on the ratio between the numberof teeth of the input gear to that of the output gear (the Gear Ratio).

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If, for example, the input gear has 25 teeth and the output gear has 75 teeth, thenthe 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, thenthe 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 rotationor 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 toaccommodate the desired effect and by the power which is to be transmittedthrough the gears.

17.2.2 SPUR GEARSThe teeth of Spur gears are ‘straight cut’, which means that the teeth are cutparallel with the axis of the shaft. Straight cut spur gears are comparatively easyto manufacture but are noisy in operation. Spur gears form the simplest of gear‘trains’.

17.2.3 HELICAL GEARSHelical gears are also used to transmit drive between parallel shafts. They aremore complex to manufacture and are quieter in operation than spur gears but(unlike spur gears), helical gears produce an axial load on their respectivebearings. Another advantage of helical gears however, is that there are moreteeth in mesh, to provide a larger contact area than straight cut gears, on wheelsof the same width. This means that helical gears can transmit more power thanstraight gears of the same axial width.

17.2.4 BEVEL GEARSBevel gears are, generally, used to transmit the drive between shafts which haveintersecting axes. The angle of intersection (and thus the drive) will vary withindividual applications. Bevel gears can be found in many places, an example ofwhich could be that, taken from the main drive shaft of an aircraft engine, .to drivean accessory gearbox.

17.2.5 WORM AND WHEEL GEARS

The worm and wheel gear set consists of a helically-cut, worm gear, on an inputshaft, driving a spur gear-mounted wheel, on an output shaft. The axes of the twoshafts cross at 90° and are in different planes. The main difference between thisconfiguration and the bevel gears is that the worm and wheel combination gives amuch larger ‘step-down’ between the ‘driver’ and ‘driven’ shaft speeds wherespace is limited, though frictional losses are higher with the worm and wheelarrangement.

This configuration can only be used to drive one way; i.e. the input and output arealways the same. This allows the input system to drive the output slowly and witha high mechanical advantage (higher torque), without any back loads being able

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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 orhydraulic motor, but in which the air loads, on the flaps, must not be allowed todrive 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 trainwhich is used to reduce the speed of an aircraft engine’s output shaft to a moreacceptable 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 ina smaller frontal area being necessary for the power unit and the subsequentreduction in aerodynamic drag.

It should also be made clear, that neither the number of teeth on the planetarygears, 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 theoverall ratio remains at 2:1.

SPIDER

Planetary (Epicyclic) Reduction Gear TrainFig. 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’, whilethe 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) ithas 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, externally-cut gears, for a similar reduction in output speed.

Drive Gear(Pinion)

Direction of Rotation

Driven Gear(Spur)

Spur and Pinion Reduction Gear TrainFig. 98

17.2.8 ACCESSORY UNIT DRIVES

Aircraft engines also employ multiple gear trains (refer to Fig. 99), in their internaland external gearboxes. These provide the drives for accessories such as fuel,hydraulic and oil pumps, electrical generators, engine speed indicators and manyother devices

Here it can be seen that ‘idler’ gears are added to reverse the rotation andpossibly 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 thegearbox, has bevel gears.

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Typical External Accessory GearboxFig. 99

17.2.9 MESHING PATTERNS

Because of the high power being transmitted by gears in certain situations andkeeping in mind that (using spur gears) only one tooth at a time can be subjectedto that power, then the point of contact between the teeth in mesh is veryimportant.

Helical gears may have as many as 5 teeth in contact at any one time, thereforepower will be spread across more teeth. The loads must be applied mid-waybetween the front and rear faces of the gear wheel. They must also be exertedbetween 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 thegearbox and are usually achieved with the use of appropriately sized shims.

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17.3 CHAINS AND SPROCKETSChains, for aircraft use, are generally the simple roller type that consist of outerand inner plates, rollers, bearing pins and bushes (refer to Fig.100).

Chains may be one of four standard sizes but, for most aircraft installations, themanufacturer dictates the size and type of chains used. They are obtained ascomplete, proof-loaded, units from manufacturers, and are identified by theirallocated part numbers in the relevant aircraft IPC. Chain links or attachmentsshould never be drilled and re-riveted. Where chains have bolts in place of rollersand 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 TerminologyFig. 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 befound 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. Mostinstallations use chains to generate and convert rotary motion at each end, butuse cables to connect the chains together over long distances.

After installation in the aircraft, the chains should be examined for freedom fromtwist. Particular attention must be paid in instances where the attachment is madeto threaded rods by means of screwed end connectors. Care should also be taketo ensure the chain is not pulled out of line by the chain wheel. The wheel shouldengage 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) andcan be employed to provide simple rotary-to-straight line motion or to change thedirection of straight line motion in one plane. A change of direction in two planescan be achieved by the use of a special ‘bi-planar block’.

Typical Arrangements of Chain AssembliesFig. 101

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17.4 MAINTENANCE INSPECTIONSChain assemblies should be inspected for serviceability at periods specified in theAMM. The continued smoothness of operation between chain and chain wheel orpulley should be checked for wear. This can be confirmed by trying to lift the linksoff the wheel teeth and checking the links for looseness. The chain should alsobe checked for damage, cleanliness, correct lubrication and freedom fromcorrosion.

If a chain is suspected of becoming elongated, it should be removed, cleaned andsubjected to a specified tensile load. Its length is then measured and thismeasurement is compared with its nominal length when it was new. Should thedifference indicate an extension of 2% or more, in any section of the chain, thenthe chain assembly will be required to be replaced.

One of the most common operations done on chain assemblies is that ofchecking the tension of the chain. Care must be taken not to twist the end fittingswhen 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 andAmerican Standards and are ‘preformed’ during manufacture. Preforming is aprocess in which each strand is formed into the shape that it will take up in thecompleted cable. This makes the cable more flexible, easier to splice and lessprone to kinking. Another advantage of preformed cables is that, in the event of awire breaking, it will lie flat within its strand, so that the cable should be less likelyto jam in its pulleys and fairleads.

Preformed cables are manufactured from either galvanised carbon steel orcorrosion-resistant steel, and are impregnated with friction-preventive lubricantduring manufacture. Non-preformed single strand cable may be found on someminor aircraft systems. Aircraft cables are usually classified by either theirminimum breaking load or nominal diameter.

It is very rare for a cable to be manufactured by an operator. They are normallyordered through the aircraft’s IPC, and the aircraft manufacturer supplies thecable fully formed with the necessary end-fittings and to the correct load factor.

18.1 TYPES OF CABLESThe 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 cabledesignated as 7 x 19, consists of 7 strands, each containing 19 wires. The twomost common forms of construction are the flexible and the extra-flexible types.

1x7 1x19 7x7 7x19Non-Flexible Non Flexible Flexible Extra Flexible

Common Forms of Aircraft Control CablesFig. 102

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18.2 CABLE SYSTEM COMPONENTSThere are many components associated with cable systems and a selection ispresented 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, themajority of modern cables have a ‘swaged splice’ end-fitting. Most end-fittings, oncontrol cables, are special-to-type and end-fittings such as fork, threaded (internaland external), and ball end-fittings (refer to Fig. 103) can be found in variouslocations. The nominal overall length of a cable will depend on the type of end-fitting which is being employed.

Overall Length

Aircraft Cable End-FittingsFig. 103

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18.2.2 TURNBUCKLES

Turnbuckles are devices which are attached (via internal or external threads) toappropriately designed end-fittings of aircraft cables and are used to join lengthsof 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 andunresponsive.

Turnbuckles are adjusted by the use of a ‘left-hand’ thread in one end of theturnbuckle, and a ‘right-hand’ thread in the other end (refer to Fig. 104). When thecentre part of the turnbuckle is rotated, its length will increase or decrease, andso it will adjust the cable tension.

The groove, around one end of the turnbuckle barrel, indicates the ‘left handthread’.

Once the correct tension has been obtained and confirmed (using a cabletensiometer), the turnbuckle is checked for‘safety’(sufficient threads areengaged 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 insertedinto the turnbuckle when the corresponding longitudinal grooves in the barrel andend fittings are aligned.

Spring Locating Clips

Groove

Spring-Locked TurnbuckleFig. 104

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18.2.3 CABLE TENSIONING DEVICES

Although the cable tension can be correctly adjusted on the ground, that settension may alter once the aircraft is in flight. This can be due to the largetemperature differentials involved- particularly with larger aircraft, which fly athigh altitudes and are capable of experiencing various climates in one flight - andthe consequences of an expanding, contracting and flexing airframe. Toovercome these problems a tension regulator is installed in some control runs.

As previously stated, engineers will use a tensiometer to set and check thetension of a cable. The tension regulator (refer to Fig. 105) is a device which hassprings, incorporated within the mechanism, to ensure that the cable tensionremains constant, regardless of the flexing and temperature changes of theairframe.

Cable Tension Regulators can be very dangerous, when disconnecting cableruns, so it is important to ensure that they are locked or ‘snubbed’, in accordancewith the AMM, before any work is done on the controls.

Tension Regulator and Cable TensiometerFig. 105

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18.2.4 CABLE FAIRLEADS

The cables of a control run must be supported otherwise they may foul theairframe structure. They are supported by fairleads (refer to Fig. 106), which areusually made from fibre. These fairleads should not be lubricated as this willcollect dirt and dust, which will cause extra wear on the cable and fairlead. Wherea change in direction of the cable is required, a pulley is normally used, due to itslow friction in comparison with fairleads. Guards are fitted to pulleys when the riskof the cable riding off the pulley is high.

The fairleads, already mentioned, simply allow the cable to pass through thebulkheads without chafing. If, however, the bulkhead is the divider between thepressure cabin and the outside air pressure, then the fairlead will be designed tobe an airtight seal, as well as a cable guide.

Cable FairleadsFig. 106

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18.2.5 PULLEYS

Cables that run from the flight deck to the control surfaces, require the ability tochange direction (possibly a number of times).

If the cable needs to change direction to another angle, the conventional methodof a pulley allows this change with little friction. The example of the elevator flyingcontrol run of a simple aircraft, (refer to Fig. 107), has pulleys that can change thedirection of the cable through a large range of angles.

A Simple Elevator Control RunFig. 107

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18.3 FLEXIBLE CONTROL SYSTEMSNormal aircraft cables are only capable of performing a pulling action, due to theirlack of rigidity, so, where a two-directional movement (push/pull) is required itwould be necessary either to employ the use of rods, with the attendant weightpenalty, 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 aflexible sleeve or conduit (refer to Fig. 108). The control is intended for pulloperation only, with the cable being returned, on release of the control lever, by areturn spring. The transmitting end of the cable is attached to the actuating leverwhilst, at the receiving end, the cable is secured to the component to beoperated.

The flexible cable is made up of several strands of stainless steel wire withnipples 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, coveredwith cotton braiding and a waterproof coating. For long runs, or runs not requiringflexibility, the Bowden cable is fed through rigid metal tubing, which can be bentover large radius curves if required.

Bowden CableFig. 108

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The control fittings are used at each end of the cable to transmit and actuate themovement required. These fittings are the hand levers and adjustable stops (referto 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 ifthe 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 tobe separated for maintenance. Junction boxes are also used, to permit eithermore than one input, to actuate a single operating lever, or one input to operate anumber of operating mechanisms.

Bowden Cable ControlsFig. 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), theTeleflex system allows a single, flexible cable to operate in both push and pullmode, without the need for a return spring.

Examples of the types of systems, operated by Teleflex controls, are engine andpropeller controls, trimming controls and fuel valves. Teleflex controls can also beused to transmit movement from one place to another, such as in a mechanicalFlap Position indicator or as interlocks between controls and throttles duringcontrol lock operation.

Like the Bowden system, described previously, the Teleflex system consists of aflexible transmitting cable operating inside a rigid or flexible metal conduit. Themain advantages are that it provides a more accurate and positive controlthroughout the range of movement and the controlled component can betemporarily 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 withgear teeth of the control units and the end-fittings (refer to Fig. 110).

Two Types of Teleflex Control CableFig .110

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The conduits operate in a similar manner to the Bowden system and are madefrom copper, aluminium or steel. The linings are of PTFE on most conduits exceptin high temperature areas like engine bays.

To operate the system, the cable and conduit are connected to control units ateach end of the control run and, in between, to other units and fittings, which areused to direct the run. In many locations, the cables are attached to lever-operated 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 ofend-fittings (refer to Fig. 111).

Examples of Teleflex Control Cable RunsFig. 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 theload.

Cables must provide a path for the flow of electrons from the source, through theload and back to the source with the minimum resistance. Additionally, two otherimportant 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 majorityof 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 importantto be able to distinguish between the different types of cable and the size of thecore. One of the main difficulties is the extensive use of nylon and terylenebraids 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 itslength. If the cable is too thin to be printed on, the code will be printed on a non-metallic sleeves positioned along the cable.

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There are many different types of wire used for special applications in aircraftelectrical systems, but the majority of the wiring is achieved with MIL-W-5086 orMIL-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 POLYVINYL CHLORIDECONDUCTOR INSULATION

EXTRUDED NYLONJACKET

MIL-G-5086 Copper WireFigure 1

Where large amounts of current must be carried for long distances, MIL-W-7072aluminium wire is often used. This wire is insulated with either "FluorinatedEthylene Propolene (FEP), nylon or fibreglass braid. Aluminium wire smaller thansix-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 thisfield may interfere with some aircraft instrumentation. For example, the light thatilluminates the compass card of a magnetic compass is powered with low-voltageDC. The field from this small voltage can deflect the compass. To minimise thisoccurrence, a two-conductor twisted wire is used to carry the current to and fromthis light. By using a twisted wire, the fields cancel each other out and thus donot interfere with the compass.

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

TINNED COPPER POLYVINYL CHLORIDECONDUCTOR INSULATION

TINNED COPPEREXTRUDED NYLON SHIELD

JACKET

Shielded WireFigure 2

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Antennas are connected to most of the radio receivers and transmitters with aspecial type of shielded wire called "Coaxial Cable". This consists of a centralconductor surrounded by an insulator and a second conductor. The spacing andconcentricity of the two conductors are critical for the most efficient transfer ofenergy 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.

OUTERSOLID INNER INSULATOR

CENTER INSULATOR JACKETCONDUCTOR

BRAID OUTERCONDUCTOR

Coaxial CableFigure 3

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19.3 DATA BUS CABLE

One special type of cable used exclusively for various digital electronic systems iscalled “Data Bus Cable”. Data bus cable typically consists of a twisted pair ofwires surrounded by electrical shielding and insulators. Digital systems operateon different frequencies, voltages and current levels. It is extremely important toensure that the correct cable is used for the system installed. The cable shouldnot be pinched or bent during installation and data bus cable lengths may also becritical. Refer to current manufacturer’s manuals for cable specifications.

Figure 4 shows an example of a data bus cable.

TINNED COPPERCONDUCTORS

DATA BUSCABLE “B”

DATA BUSCABLE “A”

ETFE TEFZEL®INSULATION ETFE TEFZEL®

JACKET

TINNED COPPERBRAID SHIELD

Data Bus CableFigure 4

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19.4 CONDUCTOR MATERIAL & INSULATION

The wires installed in an aircraft electrical system must be chosen on the basis oftheir ability to carry the required current without overheating and to carry it withoutproducing an excessive voltage drop. There are a number of factors to considerwhen 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 theconductor is made from copper, the individual strands of wire are typically platedto protect the copper from corrosion. Figure 5 shows two types of conductorfound in aircraft systems.

STRANDEDCONDUCTORS

SOLIDCONDUCTOR

Electrical Wire TypeFigure 5

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The number of strands that make up the wire and the type of insulation on thewire typically determine the flexibility of a conductor. The type of insulation isvery important; various insulations have different ratings for heat, abrasion andflexibility. The length and type of installation are factors established by theaircraft 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 diameterand is indicated by a unit called “Circular Mil”. One circular mil is equal to thecross-sectional area of a 1-mil (0.001-in) diameter wire, measured in thousandthsof an inch. To determine the size in circular mils of a wire, simply square thewire's diameter measured in thousandths of an inch. Figure 6 shows thisconcept.

1 mil2 0.001 IN1 CIRCULAR MIL

(1 cmil)

American Wire Gauge (AWG)Figure 6

In AWG only even numbers are used, small wires have higher numbers, typicallystarting 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 7shows a typical wire gauge tool.

NOTE: AWG FORELECTRICAL WIRE

IS 24 - 0000

Wire Gauge ToolFigure 7

19.6 WIRE RESISTANCE

Resistance is the opposition to current flow and is measured in Ohms (). Theresistance of a wire will increase with an increase of length, but will decrease withan increase of cross-sectional area.

19.7 CURRENT CARRYING CAPABILITY

A wire fitted to an aircraft system should be able to carry the required currentwithout overheating and burning. Also it must be able to carry the required

current without producing a voltage drop greater than that which is permissible forthe circuit.

Most aircraft wiring that is required to carry large amounts of current for longdistances, is generally made up of aluminium wire.

Tables 1 and 2 shows the characteristics of MIL-W-5086 copper wire and MIL-W-7072 aluminium wire.

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Wire Size Single Wire Bundled Wire Max WeightMax Amps Max Amps Resistance Pounds per(In free Air) (Conduit) Ohms/1,000ft 1,000ft

(20°C)AN-20 11 7.5 10.25 5.6AN-18 16 10 6.44 8.4AN-16 22 13 4.76 10.8AN-14 32 17 2.99 17.1AN-12 41 23 1.88 25AN-10 55 33 1.1 42.7AN-8 73 46 0.7 69.2AN-6 101 60 0.44 102.7AN-4 135 80 0.27 162.5AN-2 181 100 0.18 247.6AN-0 245 150 0.11 382AN-00 283 175 0.09 482AN-000 328 200 0.07 620AN-0000 380 225 0.06 770

MIL-W-5086Table 1

Wire Size Single Wire Bundled Wire Max WeightMax Amps Max Amps Resistance Pounds per(In free Air) (Conduit) Ohms/1,000ft 1,000ft

(20°C)

AL-6 83 50 0.64AL-4 108 66 0.43AL-2 152 90 0.27AL-0 202 123 0.17 166AL-00 235 145 0.13 204

AL-000 266 162 0.11 250AL-0000 303 190 0.09 303

MIL-W-7072Table 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 outthis 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 allowaluminium wire smaller (in size, larger in number), than 6-gauge to be used onaircraft.

19.8 VOLTAGE DROP

When we add any electrical equipment to an aircraft, we must be sure that thecurrent flowing in the wiring does not drop the voltage below a set level. Table 3shows an example of the allowable voltage drop for various systems usingvarious supply voltages.

Nominal Allowable Voltage Drop - VoltsSystem

VoltageContinuous Intermittent

Operation Operation14 0.5 128 1 2115 4 8200 7 14

Allowable Voltage DropTable 3

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19.9 WIRE IDENTIFICATION

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

22 GAUGE

26TH WIRE INTHE CIRCUIT

FLIGHTINSTRUMENTATION

4TH SEGMENT

WIRE

GROUND(A= 1ST SEGMENT)

Wire CodeFigure 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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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 externalprotection. This method is used when there is no great danger of mechanicaldamage (Chafing, Rubbing). This type of installation is easy to install andmaintain, and is lighter in weight.

Wires are grouped and tied together in bundles for the neatest and most efficientrouting. No one bundle should carry wires from circuits that would disable bothmain and back-up systems. The bundles should be routed so as not to interferewith any of the controls or moving components. They must be routed where theycannot be damaged by persons entering or leaving the aircraft or by baggage orcargo moving over them or resting on them.

Figure 9 shows an example of an Open Wire bundle fitted to an aircraft sidewall.

“P” CLIPS ATTACHINGBUNDLE TO AIRCRAFT

FRAME

WIREBUNDLE

½ INCH MAXIMUMWITH NORMAL HAND

PRESSURE

CABLEBUNDLE

“P” CLIP

Open Wire BundleFigure 9

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19.12 WIRE & CABLE CLAMPING

Electrical cables or wire bundles are secured to the aircraft structure by means ofmetal 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 stressapplied by the cable to the clamp is not in a direction that will tend to bend theclamp. When a clamp is mounted on a vertical member, the loop of the clampshould always be at the bottom. Correct methods for installing clamps is shownin Figure 10.

DANGEROUS ANGLES

SAFE ANGLES

Correct Methods of Installing Cable clampsFigure 10

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19.13 CONDUIT

Mechanical protection can be provided for the wire by routing the bundles througheither flexible or rigid conduit. The size of the conduit is normally an insidediameter 25% larger than the diameter of the wire bundle being encased. Figure

11 shows the two types of conduits.

MINIMUM BENDCABLE

CONDUIT

CONDUIT

CLAMP

METALLICCONDUIT

CLAMP

RADIUS

(FOUR TIMESINSIDE

DIAMETER)

ADAPTOR

INSIDE

DIAMETER

CLAMP

ADAPTOR

FLEXIBLE CONDUIT

BRACKET

CABLECLAMP

RIGID CONDUIT

Cable ConduitFigure 11

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

LINE REPLACEMENTUNIT (LRU) CABLE

CONDUIT

DRAINHOLE

PLUGCONNECTION

Conduit Drain HoleFigure 12

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19.14 CONNECTORS

Most of the electrical components in an aircraft are designed so that they may beserviced with a minimum amount of time needed for their removal andinstallation. The electrical wiring is usually connected through quick-releaseplugs. 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 areclamped tight to prevent mechanical strain on the cable being transmitted into theconnectors themselves.

The most commonly used connector is the Military Standard (MS), type. EachMS connector has an identification number on it, Figure 13 shows a connectorand identification number.

MILITARYSTANDARD

TYPENUMBER

CLASS

SIZE

INSERTARRANGEMENT

NUMBER

CONTACTSTYLE

INDEXSLOT INSERT

NUMBER

Connector Identification NumberFigure 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 thenumber 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" (femaleor male), arrangement. The final letter in the identification is one of the lastletters in the alphabet, "W", "X", "Y" or "Z". These letters indicate the rotation ofthe insert in the connector. It is possible to connect the wrong plug to areceptacle, so to prevent this, the inserts may be rotated in their relationship tothe index slot. This ensures only the correct plug may be inserted into thereceptacle.

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

MS 3100BULKHEAD

RECEPTACLE

MS 3108BULKHEAD PLUG

MS 3101CABLE RECEPTACLE

MS 3102BOX RECEPTACLE

MS Quick-Release ConnectorsFigure 14

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19.15 CRIMPING

Crimping is a method of firmly attaching a terminal end to an electrical conductorby pressure forming or reshaping a metal barrel, together with the conductor.The forming of a satisfactory crimped joint depends on the correct combination ofconductor, crimp barrel and tool.

When applied with the correctly matched tool, a joint would be established whichhas both good electrical and mechanical properties. Figure 15 show a crimpedterminal.

STRIPPED WIRECONDUCTOR

DIAMOND GRIPCRIMP FORINSULATIONSUPPORT

CRIMPI

NSULATIONWIREINSULATION

Pre-Insulated Crimped TerminalFigure 15

TERMINALRING

CROSS CRIMPFOR GRIPPING

WIRE STRANDS

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19.16 CRIMPING TOOLS

There are a number of types of crimping tool available, but the best ones have aratchet mechanism that will not allow them to open until they have crimped theterminal to the proper size. These tools, often referred to as "PrecisionTermination Tools (PTT), require periodical calibration checks. If a terminal isproperly 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-insulatedwire terminals.

CRIMPINGHEAD

CRIMPINGJAWS

CONDUCTORBEING CRIMPED

RATCHETMECHANISM

HANDLE

Heavy-duty Crimping ToolsFigure 16

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19.17 WIRE SPLICING

The splicing of electrical wires may be done if approved for a particularinstallation. 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 aplain metal tube that is covered with a plastic tube after the splice has beenmade.

The stripped wire is inserted into the end of the tube and then crimped with aterminal crimping tool. When splices are made in wires that are in a cablebundle, the spliced wires are placed on the outside of the bundle. If severalsplices are to be made in any cable bundle, the splices should be staggered toreduce the bundle diameter. Figure 17 shows various situations of splices in acable bundle.

DO NOT PUTCABLE LACING

ON TOP OFTHE SPLICES

DISTRIBUTE SPILCESIN A CBLE BUNDLE

EVENLY ON THE OUTSIDEOF THE BUNDLE

1 CM

MIN IM U M

3 - PHASEPOWER SUPPLY

2 CMMIN IM U M

3 - PHASEPOWER SUPPLY

METALTUBE

CAB L E SIZEAWG 8 OR

LAR GER

CAB LE SIZ EAWG 8 OR

LA R GER

PLASTICIN S ULATIO N

CABLE SPLICE CONSTRUCTION

Cable SplicesFigure 17

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19.18 BEND RADIUS

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

CONNECTOR

RADIUS AT LEASTSIX TIMES OUTER

DIAMETER

STRAIGHT STRAINRELIEF

Bend RadiusFigure 18

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

TERMINALBLOCK

RADIUSMINIMUM OF THREETIMES THE OUTER

DIAMETER OFCABLE

Bend radius (Supported)Figure 20

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