Index
1.
INTRODUCTION ....................................................................................... 1-1
2.
TESTING OF MATERIALS ....................................................................... 2-1
2.1 TENSILE TESTING ..................................................................................... 2-1
2.1.1 Tensile strength .............................................................................. 2-1
2.2 LOAD EXTENSION DIAGRAMS .................................................................... 2-3
2.3 DUCTILITY ............................................................................................... 2-4
2.4 PROOF STRESS ........................................................................................ 2-5
2.5 STIFFNESS............................................................................................... 2-5
2.6 TENSILE TESTING OF PLASTICS ................................................................. 2-6
2.7 COMPRESSION TEST................................................................................. 2-6
2.8 HARDNESS TESTING ................................................................................. 2-6
2.8.1 Brinell test ....................................................................................... 2-6
2.8.2 Vickers test ..................................................................................... 2-6
2.8.3 Rockwell test................................................................................... 2-7
2.8.4 Hardness testing on aircraft ............................................................ 2-7
2.9 IMPACT TESTING ...................................................................................... 2-7
2.10 OTHER FORMS OF MATERIAL TESTING .................................................... 2-8
2.11 CREEP ................................................................................................. 2-8
2.12 CREEP IN METALS................................................................................. 2-8
2.12.1 Effect of stress and temperature on creep ................................... 2-8
2.12.2 The effect of grain size on creep ................................................. 2-9
2.12.3 Creep in plastics .......................................................................... 2-9
2.13 FATIGUE .............................................................................................. 2-9
2.14 FATIGUE TESTING ............................................................................... 2-10
2.15 S-N CURVES ...................................................................................... 2-10
2.16 CAUSES OF FATIGUE FAILURE ............................................................. 2-11
2.17 VIBRATION ......................................................................................... 2-12
2.18 FATIGUE METALLURGY ....................................................................... 2-12
2.19 FATIGUE PROMOTERS ......................................................................... 2-13
2.19.1 Design ....................................................................................... 2-13
2.19.2 Manufacture .............................................................................. 2-13
2.19.3 Environment .............................................................................. 2-13
2.20 FATIGUE PREVENTERS ........................................................................ 2-13
2.21 PRACTICAL DO'S AND DONT'S TO HELP PREVENT FATIGUE FAILURES ..... 2-14
2.22 FATIGUE MONITORING (MODERN TRENDS) ............................................ 2-14
2.22.1 Fatigue fuses ............................................................................. 2-14
2.23 INTELLIGENT SKINS DEVELOPMENT...................................................... 2-14
2.23.1 Structural Health Monitoring (SHM) ........................................... 2-15
2.24 COLD WORKING .................................................................................. 2-16
3.
FERROUS METALS ................................................................................. 3-1
3.1 PROCESSING THE RAW MATERIAL ............................................................. 3-1
3.2 NODULAR CAST IRON ............................................................................... 3-1
3.3 STEEL ..................................................................................................... 3-1
3.3.1 Carbon in steel................................................................................ 3-1
3.3.2 Metallurgical structure of steel......................................................... 3-2
3.3.3 The structure and properties of slowly cooled steels ....................... 3-2
3.3.4 The effect of cooling rates on carbon steels .................................... 3-2
3.4 HEAT TREATMENT OF CARBON STEELS ...................................................... 3-3
3.4.1 Problems associated with the hardening process............................ 3-3
3.4.2 Tempering....................................................................................... 3-4
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3.4.3 Annealing ........................................................................................ 3-4
3.4.4 Normalising ..................................................................................... 3-4
3.5 SURFACE HARDENING OF STEELS ............................................................. 3-4
3.5.1 Carburising ..................................................................................... 3-4
3.5.2 Nitriding .......................................................................................... 3-5
3.5.3 Areas to be left un-hardened ........................................................... 3-5
3.5.4 Flame / induction hardening ............................................................ 3-5
3.5.5 Other hard skin techniques ............................................................. 3-5
3.6 ALLOY STEELS ......................................................................................... 3-5
3.7 NICKEL STEELS ........................................................................................ 3-5
3.8 NICKEL ALLOYS ....................................................................................... 3-6
3.9 CHROME STEEL ....................................................................................... 3-6
3.10 NICKEL-CHROME STEELS AND ALLOYS................................................... 3-6
3.11 MANGANESE STEEL .............................................................................. 3-6
3.12 TUNGSTEN STEEL ................................................................................. 3-7
3.13 COBALT STEEL AND ALLOYS ................................................................. 3-7
3.14 VANADIUM STEEL ................................................................................. 3-7
3.15 MARAGING STEELS ............................................................................... 3-7
4.
NON FERROUS METALS ........................................................................ 4-1
4.1 PURE ALUMINIUM ..................................................................................... 4-1
4.2 ALUMINIUM ALLOYS ................................................................................. 4-1
4.3 ALUMINIUM-COPPER ALLOYS .................................................................... 4-1
4.4 OTHER AGE HARDENING ALUMINIUM ALLOYS ............................................. 4-1
4.5 CLAD MATERIALS ..................................................................................... 4-1
4.6 WORK / STRAIN HARDENING ALUMINIUM ALLOYS ........................................ 4-1
4.7 ALLOY IDENTIFICATION SYSTEMS .............................................................. 4-2
4.7.1 IADS (international alloy designation system) ................................. 4-2
4.8 MATERIAL CONDITION............................................................................... 4-2
4.9 BASIC TEMPER DESIGNATIONS .................................................................. 4-3
4.10 EXAMPLES OF IADS CODES ................................................................... 4-3
4.11 OTHER NUMBERING SYSTEMS................................................................ 4-6
4.12 MARKING OF ALUMINIUM ALLOY SHEETS ................................................ 4-6
4.13 CAST ALUMINIUM ALLOYS ..................................................................... 4-6
4.14 MAGNESIUM & IT'S ALLOYS ................................................................... 4-7
4.15 TITANIUM ............................................................................................. 4-7
4.16 CUTTING FLUIDS ................................................................................... 4-8
4.17 DRILLING TITANIUM ............................................................................... 4-8
4.17.1 Speed.......................................................................................... 4-8
4.17.2 Feed ............................................................................................ 4-8
4.17.3 Cutting fluid ................................................................................. 4-8
4.17.4 Centre drilling .............................................................................. 4-8
4.17.5 Safety precautions....................................................................... 4-9
5.
HEAT TREATMENT & APPLICATIONS OF NON FERROUS METALS... 5-1
5.1 SOLUTION HEAT TREATMENT .................................................................... 5-1
5.2 PRECIPITATION HEAT TREATMENT ............................................................. 5-1
5.3 ANNEALING HEAT TREATMENT .................................................................. 5-1
5.4 HEAT TREATMENT NOTES ......................................................................... 5-1
5.5 HEAT TREATMENT OF ALUMINIUM ALLOY RIVETS ........................................ 5-2
6.
METALLIC FORMING METHODS ............................................................ 6-1
6.1 CASTING............................................................................................... 6-1
6.1.1 Sand casting ................................................................................... 6-1
6.1.2 The sand ......................................................................................... 6-1
6.1.3 The mould ....................................................................................... 6-2
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6.1.4 Advantages / disadvantages of sand casting .................................. 6-2
6.1.5 Typical casting defects .................................................................... 6-2
6.1.6 Shell moulding ................................................................................ 6-2
6.1.7 Centrifugal casting .......................................................................... 6-3
6.1.8 Die casting ...................................................................................... 6-3
6.1.9 Investment casting .......................................................................... 6-3
6.2 FORGING ................................................................................................. 6-3
6.2.1 Drop stamping ................................................................................ 6-4
6.2.2 Hot pressing.................................................................................... 6-4
6.2.3 Upsetting ........................................................................................ 6-4
6.3 ROLLING ................................................................................................. 6-4
6.4 DRAWING ................................................................................................ 6-4
6.5 DEEP DRAWING / PRESSING ...................................................................... 6-5
6.6 PRESSING ............................................................................................... 6-5
6.7 STRETCH FORMING .................................................................................. 6-5
6.8 RUBBER PAD FORMING ............................................................................. 6-5
6.9 EXTRUSION.............................................................................................. 6-5
6.9.1 Impact extrusion ............................................................................. 6-5
6.10 SINTERING ........................................................................................... 6-5
6.11 SPINNING ............................................................................................. 6-6
6.12 CHEMICAL ‘MILLING’ ............................................................................. 6-6
6.13 ELECTRO CHEMICAL MACHINING ........................................................... 6-6
6.14 ELECTRO-DISCHARGE MACHINING E.D.M. ............................................... 6-6
6.15 CONVENTIONAL MACHINING .................................................................. 6-7
6.16 SUPERPLASTIC FORMING ...................................................................... 6-8
7.
COMPOSITES & NON METALLIC MATERIALS ...................................... 7-1
7.1 ARAMIDES ............................................................................................... 7-1
7.2 GLASS REINFORCED PLASTIC (GRP) .......................................................... 7-1
7.3 CARBON FIBRES, REINFORCED PLASTIC (CFRP) ......................................... 7-2
7.3.1 Recent developments ..................................................................... 7-2
7.3.2 General information ........................................................................ 7-3
7.4 PLASTIC SEALANTS & ADHESIVES ............................................................. 7-3
7.4.1 Plastics ........................................................................................... 7-3
7.4.2 Major plastic groups ........................................................................ 7-3
7.4.3 Primary advantages of the use of plastics ....................................... 7-4
7.4.4 Primary disadvantages of plastics ................................................... 7-4
7.4.5 Main uses for plastics .................................................................... 7-5
7.4.6 Some of the more common plastics & their application ................... 7-5
7.4.7 Thermo-setting plastics ................................................................... 7-7
7.4.8 Elastomer plastics (synthetic rubbers)............................................. 7-7
7.4.9 Plastic forms ................................................................................... 7-7
7.4.10 Plastic manufacturing processes ................................................. 7-8
7.4.11 Adhesives & sealants .................................................................. 7-8
7.4.12 The mechanics of bonding .......................................................... 7-8
7.4.13 Stresses on a bonded joint .......................................................... 7-9
7.4.14 Design of adhesive joints............................................................. 7-9
7.4.15 Advantages ................................................................................. 7-9
7.4.16 Disadvantages........................................................................... 7-10
7.4.17 Strength of adhesives ................................................................ 7-10
7.4.18 Types of adhesives ................................................................... 7-10
7.4.19 Adhesive forms.......................................................................... 7-11
7.4.20 Adhesives in use ....................................................................... 7-12
7.4.21 Surface preparation ................................................................... 7-12
7.4.22 Final assembly .......................................................................... 7-12
7.4.23 Typical (abbreviated) process.................................................... 7-12
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7.4.24 Safety ........................................................................................ 7-13
7.5 DEFECTS IN COMPOSITE COMPONENTS ................................................... 7-13
7.5.1 Causes of damage ........................................................................ 7-13
7.5.2 Types of damage .......................................................................... 7-13
7.5.3 Assessment of damage................................................................. 7-14
7.6 REPAIRS TO COMPOSITES ....................................................................... 7-14
7.6.1 Glass fibre composite repairs ........................................................ 7-14
7.6.2 Types of glass reinforcement ........................................................ 7-14
7.6.3 Resins........................................................................................... 7-15
7.6.4 Mixing ........................................................................................... 7-15
7.6.5 Pot life........................................................................................... 7-15
7.6.6 Curing ........................................................................................... 7-15
7.6.7 Gel coat ........................................................................................ 7-16
7.6.8 Fillers ............................................................................................ 7-16
7.6.9 Storage of GRP materials ............................................................. 7-16
7.6.10 Storing resin .............................................................................. 7-16
7.6.11 Storing hardener........................................................................ 7-16
7.6.12 Storing fabrics ........................................................................... 7-17
7.6.13 Safety precautions..................................................................... 7-17
7.6.14 Damage necessitating manufacturers liaison ............................ 7-17
7.6.15 Strength considerations of GRP repairs .................................... 7-18
7.6.16 Preparation for repair................................................................. 7-18
7.6.17 Surface Preparation .................................................................. 7-19
7.6.18 Techniques of laminating glass fibre.......................................... 7-19
7.6.19 Pre-wetting glass fibre ............................................................... 7-20
7.6.20 Repair to GRP skin .................................................................... 7-21
7.6.21 Repairs to multiple laminations .................................................. 7-22
7.6.22 Repair to sandwich structure ..................................................... 7-24
8.
CORROSION ............................................................................................ 8-1
8.1 CHEMISTRY & MECHANISMS...................................................................... 8-1
8.2 CAUSES .................................................................................................. 8-1
8.2.1 Direct chemical attack ..................................................................... 8-1
8.2.2 Electrochemical attack .................................................................... 8-1
8.3 TYPE AND SUSCEPTIBILITY ....................................................................... 8-2
8.3.1 Forms of corrosion .......................................................................... 8-2
8.3.2 Surface ........................................................................................... 8-2
8.3.3 Dissimilar metal corrosion ............................................................... 8-2
8.3.4 Intergranular corrosion .................................................................... 8-2
8.3.5 Exfoliation corrosion........................................................................ 8-2
8.3.6 Stress corrosion .............................................................................. 8-2
8.3.7 Fretting corrosion ............................................................................ 8-3
8.3.8 Crevice corrosion ............................................................................ 8-3
8.3.9 Filiform corrosion ............................................................................ 8-3
8.3.10 Microbiological contamination...................................................... 8-3
8.3.11 Factors affecting corrosion .......................................................... 8-3
8.3.12 Climatic ....................................................................................... 8-3
8.3.13 Size and type of metal ................................................................. 8-3
8.3.14 Corrosive agents ......................................................................... 8-4
8.4 CORROSION REMOVAL ............................................................................. 8-4
8.4.1 Cleaning and paint removal. ........................................................... 8-4
8.4.2 Corrosion of ferrous metals ............................................................. 8-4
8.4.3 High stressed steel components ..................................................... 8-5
8.4.4 Aluminium and aluminium alloys ..................................................... 8-5
8.4.5 Alclad .............................................................................................. 8-5
8.4.6 Typical painted corrosion treatment sequence ................................ 8-5
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8.4.7 Permanent anti-corrosion treatments .............................................. 8-6
8.4.8 Surface conversion coatings ........................................................... 8-6
8.4.9 Acid spillage.................................................................................... 8-6
8.4.10 Alkali spillage .............................................................................. 8-7
8.4.11 Mercury spillage .......................................................................... 8-7
8.4.12 Identification of metals ................................................................. 8-8
9.
AIRCRAFT FASTENERS.......................................................................... 9-1
9.1 SCREW THREAD NOMENCLATURE.............................................................. 9-1
9.2 THREAD FORMS ....................................................................................... 9-2
9.3 BOLTS ..................................................................................................... 9-2
9.3.1 British bolts ..................................................................................... 9-2
9.3.2 American bolts ................................................................................ 9-4
9.4 NUTS....................................................................................................... 9-8
9.5 STUDS ................................................................................................... 9-11
9.5.1 Fitting studs .................................................................................. 9-11
9.5.2 Stud removal................................................................................. 9-12
9.6 FRICTIONAL LOCKING DEVICES ............................................................... 9-13
9.7 POSITIVE LOCKING DEVICES.................................................................... 9-14
9.7.1 Split pins (Cotter pins in U.S.A.) .................................................... 9-14
9.7.2 Terry Pins ..................................................................................... 9-14
9.7.3 Tab Washers ................................................................................ 9-14
9.7.4 Locking Plates .............................................................................. 9-15
9.7.5 Taper Pins and Parallel Pins ......................................................... 9-15
9.7.6 Centre Popping and Peening (Burring) ......................................... 9-15
9.7.7 Wire Locking ................................................................................. 9-15
9.7.8 Methods of locking ........................................................................ 9-16
9.7.9 Wire locking principles .................................................................. 9-17
9.7.10 Turnbuckles and adjustable strut (control rods) ......................... 9-18
9.7.11 Locking, restraining & tell-tale wire on controls & switches ........ 9-19
9.8 MISCELLANEOUS FASTENERS ................................................................. 9-20
9.8.1 Hi lock and high / tigue fasteners .................................................. 9-20
9.8.2 Special purpose fasteners............................................................. 9-20
9.8.3 Turnlock fasteners (¼ turn fasteners)............................................ 9-24
9.9 SOLID RIVETS ........................................................................................ 9-26
9.9.1 Solid rivets (British) ....................................................................... 9-26
9.9.2 Solid rivets (American) .................................................................. 9-28
9.9.3 Heat treatment .............................................................................. 9-31
9.10 BLIND RIVETS ..................................................................................... 9-31
9.10.1 Friction lock rivet ....................................................................... 9-32
9.10.2 Mechanical lock rivets ............................................................... 9-32
9.10.3 Pull through rivets...................................................................... 9-33
9.10.4 Grip range ................................................................................. 9-34
9.10.5 Examples of blind rivets............................................................. 9-34
10.
PIPES & UNIONS ................................................................................ 10-1
10.1 PIPELINES ......................................................................................... 10-1
10.1.1 Pipeline connectors ................................................................... 10-1
10.1.2 Locking ...................................................................................... 10-2
10.1.3 Precautions ............................................................................... 10-2
10.2 HOSES AND HOSE ASSEMBLIES .......................................................... 10-3
10.2.1 Classification ............................................................................. 10-3
10.2.2 Construction .............................................................................. 10-3
10.2.3 Pre-installation checks............................................................... 10-5
10.2.4 Installation ................................................................................. 10-6
10.2.5 Hose assemblies with re-usable end fittings .............................. 10-6
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10.2.6 The end fittings .......................................................................... 10-8
10.2.7 The new hose .......................................................................... 10-11
10.2.8 Preparing the hose .................................................................. 10-12
10.2.9 Mating hoses & end fittings ..................................................... 10-13
10.2.10 Examination of locally made up assemblies ............................ 10-13
10.2.11 Testing locally made up assemblies ........................................ 10-14
10.2.12 Proof testing ............................................................................ 10-14
10.2.13 Installing hose assemblies ....................................................... 10-14
10.2.14 Protective sleeves ................................................................... 10-15
10.3 RIGID PIPES...................................................................................... 10-17
10.3.1 Manufacture of rigid pipes ....................................................... 10-17
10.3.2 Materials.................................................................................. 10-17
10.3.3 Introduction ............................................................................. 10-17
10.3.4 Bending tubes ......................................................................... 10-18
10.3.5 Unloading ................................................................................ 10-18
10.3.6 Pipe bending machine ............................................................. 10-18
10.3.7 Compression bending machines ............................................. 10-18
10.3.8 Preparing tube ends ................................................................ 10-19
10.3.9 Flaring operation ..................................................................... 10-19
10.3.10 Stainless Steel......................................................................... 10-20
10.4 IDENTIFICATION ................................................................................ 10-21
10.5 TESTING & LIFE ................................................................................ 10-21
10.5.1 Hoses ...................................................................................... 10-21
10.5.2 Rigid pipes .............................................................................. 10-21
10.6 UNIONS ............................................................................................ 10-22
10.7 FLARES & FLARELESS ...................................................................... 10-26
10.7.1 Flareless couplings.................................................................. 10-26
10.7.2 Fitting procedures.................................................................... 10-26
11.
11.1
11.2
11.3
12.
SPRINGS............................................................................................. 11-1
TYPES IN USE ..................................................................................... 11-1
MATERIALS ........................................................................................ 11-1
APPLICATIONS ................................................................................... 11-1
BEARINGS .......................................................................................... 12-1
12.1 PURPOSE ........................................................................................... 12-1
12.2 CONSTRUCTION .................................................................................. 12-1
12.2.1 Ball bearings ............................................................................. 12-1
12.2.2 Roller bearings .......................................................................... 12-2
12.2.3 Maintenance of bearings ........................................................... 12-3
12.3 INSPECTION ....................................................................................... 12-4
12.3.1 General inspection procedure & faults ....................................... 12-4
12.4 STORAGE ........................................................................................... 12-4
12.5 TRANSMISSION ................................................................................... 12-5
12.5.1 Keys and keyways ..................................................................... 12-5
12.5.2 Splined & serrated drives .......................................................... 12-6
12.5.3 Master spline ............................................................................. 12-7
12.5.4 Examination .............................................................................. 12-7
12.5.5 Chains ....................................................................................... 12-7
12.5.6 Gear & gear trains – types & uses ........................................... 12-14
13.
CONTROL CABLES ........................................................................... 13-1
13.1 TYPES ............................................................................................... 13-1
13.2 FITTINGS ............................................................................................ 13-2
13.3 PULLEYS & BELL CRANKS ................................................................... 13-2
13.3.1 Pulleys ...................................................................................... 13-2
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13.3.2 Fairleads ................................................................................... 13-3
13.3.3 Screwjack .................................................................................. 13-3
13.3.4 Cable tensioning........................................................................ 13-4
13.3.5 Cable tension regulator ............................................................. 13-5
13.4 RIGID CONTROL CABLES ..................................................................... 13-6
13.4.1 Bowden controls ........................................................................ 13-7
13.4.2 Teleflex controls ...................................................................... 13-14
13.4.3 Teleflex control units & fittings ................................................. 13-17
13.4.4 Aircraft flexible control systems ............................................... 13-23
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1.
INTRODUCTION
In June 1990 at 17000ft the captain of a British Airways aircraft was suddenly half sucked out of the
front windscreen when it blew out. The windscreen departed because incorrect bolts were fitted when
the windscreen was replaced earlier. This module includes explanation of Bolt identification. It also
covers the types of material used in the construction of aircraft and the mechanisms.
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2.
TESTING OF MATERIALS
The mechanical properties of a material must be known by an engineer before he can incorporate that
material into a design. Mechanical property data is compiled from extensive material testing. Various
tests are used to determine the actual values of material properties under different loading applications
and test conditions.
2.1
TENSILE TESTING
Tensile testing is the most widely used mechanical test. It involves applying a steadily increasing load
to a test specimen, causing it to stretch until it eventually fractures. Accurate measurements are taken
of the load and extension, and the results are used to determine the strength of the material. To ensure
uniformity of test results, the test specimens used must conform to standard dimensions and finish as
laid down by the appropriate Standards Authority (British Standards). The cross section of the
specimen may be round or rectangular, but the relationship between the cross sectional area and a
specified "gauge length" of each specimen is constant. - The gauge length, is that portion of the parallel
part of the specimen which is to be used for measuring the subsequent extension during and/or after
the test.
2.1.1 TENSILE STRENGTH
Tensile strength in a material is obtained by measuring the maximum load which the test piece is able
to sustain, and dividing that value by the cross sectional area of the specimen. The applied load is
usually measured in Newtons (N) in Europe, or Tons (T) in the USA and the cross sectional area in
square millimetres (mm2) in Europe, or square inches (ins2) in USA. The value derived from this simple
calculation is called STRESS.
Stress
Load (N)
Original c.s.a. (mm2 )
The units of stress are N/mm2 (Europe). USA and old British Imperial units of stress are quoted in T/in2
(or T.S.I.).
Example 1
A steel rod 5 mm in diameter is loaded in tension with a force of 400 N. Calculate the tensile stress.
Stress
Load
Area
400
400

 20  37N / mm 2
2
2
r
 25
Exercise 1
Calculate the tensile stress in a steel rod with a rectangular section of 10 mm x 4 mm when it is
subjected to a load of 100 N.
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Exercise 2
Calculate the 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 the
area in square metres (m2) and load in MEGA-NEWTONS (N6). When using such units, the numerical
value is identical to that if the calculation had been made using mm2 and Newtons.
i.e.
A Stress of 1 N/mm2 = l MN/m2
Example 2
A structural member with a cross sectional area of 05m2 is subjected to a load of 10 MN. Calculate the
stress in the member in;
a. MN/m2 and b. N/mm2
a.
Stress
10
 20MN / m 2
05
Load
Area
b. 1N/mm2  1MN/m2
So Stress 20 N/mm2
As the load in the tensile test is increased from zero to a maximum value, the material extends. The
amount of extension produced by a given load, enables us to calculate the amount of strain produced.
Strain is calculated by measuring the extension and dividing by the original length of the material.
Strain 
Extension
Original Length
Note: Both measurements must be in the same units (usually mm).
Since strain is a ratio of two lengths, it has no units.
Strain 
Extension
1  15

 0  0575 (no units)
Original Length
20
Example 3
An aluminium test piece is marked with a 20 mm gauge length. It is strained in tension so that the
gauge length becomes 2115 mm. Calculate the strain.
Extension  21 15 - 20  1 15 mm
Exercise 3
A tie rod 1.5m long under a tensile load of 500 N extends by 12 mm. Calculate the strain.
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2.2
LOAD EXTENSION DIAGRAMS
If a gradually increasing tensile load is applied to a test piece and the load and extension continuously
measured, the results can be used to produce a Load / Extension graph. A number of different types
of graph will be obtained, depending on the material and its condition. The diagram below shows a
Load / Extension graph which typifies many metallic materials when stressed in tension.
The graph can be considered as comprising two major regions. Between points 0 and A the material is
elastic, i.e. when the load is removed the material will return to its original size and shape. In this
region, the extension is directly proportional to the applied load. This relationship is known as ‘Hooke's
Law’, which states :
Within the elastic region, elastic strain is directly proportional to the stress causing it.
Point A is the elastic limit. Between this point and point B the material continues to extend until the
maximum load is reached at point B. In this region the material is plastic. When the load is removed,
the material does not return to its original size and shape, but will retain some extension. After point B,
the cross sectional area reduces and begins to neck. The material continues to extend under reduced
load until it eventually fractures at C.
Aircraft structural designers interest in materials doesn't extend greatly beyond the elastic phase of
materials. Production engineers however are greatly interested in material properties beyond this
phase, since the forming capabilities of materials are dependent on their properties in the plastic
phase. The graph below shows the results obtained from a test on mild steel.
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The graph shows that considerable plastic extension occurs without any increase in load shortly after
the plastic limit is reached. The onset of increasing extension without a corresponding increase in load
at point `B' is known as the ‘yield point’, and if this level of stress is reached the metal is said to have
‘yielded’. This is a characteristic of mild steel and a few other relatively ductile materials.
If the load is further increased after passing the yield point, mild steel is capable of withstanding this
increase until the UTS is reached after which severe necking occurs and the material will fracture at a
reduced load. The unexpected ability of mild steel to accept more load after yielding is due to strain
hardening of the material. Work hardening of many materials is often carried out to increase the
strength.
Various types of load extension curves for other materials are shown below.
 ‘a’ represents a brittle material (glass)
 ‘b’ represents a material with some elasticity and limited plasticity (high carbon steel).
 ‘c’ represents a material with some elasticity and good plasticity (e.g. soft aluminium).
2.3
DUCTILITY
After fracture of a specimen following tensile testing an indication of material ductility is arrived at by
establishing the amount of plastic deformation which occurred. The two indicators of ductility are :
 Elongation
 Reduction in area (at the neck)
Elongation is the more reliable because it is easier to measure the extension of the gauge length than
the reduction in area. The standard measure of ductility is to establish the percentage elongation after
fracture.
Percentageelongation
Final Extension
 100
Original Gauge Length
Example 4
In a tensile test on a specimen with 56.5 mm gauge length, the length over the gauge marks at fracture
were 71.1 mm, What was the percentage elongation?
Elongation 
Final Extension
71  1 - 56  5
 100 
 100  25  8%
Gauge Length
56  5
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2.4
PROOF STRESS
Many materials do not exhibit a yield point, and so a substitute value must be employed. The value
chosen is the ‘Proof Stress’ which is defined as :
The tensile stress which is just sufficient to produce a non proportional elongation equal to a
specified percentage of the original gauge length.
Usually a value of 0.1% or 0.2% is used for Proof Stress, and the Proof Stress is then referred to as the
0.1% Proof Stress or the 0.2% Proof Stress respectively.
The Proof Stress may be obtained from the Load / Extension graph as follows :
If the 0.2% Proof Stress is required, 0.2% of the gauge length is marked on the extension axis. Then a
line parallel to the straight line portion of the graph is drawn until it intersects the non-linear portion of
the curve. The corresponding load is then read from the graph. Proof Stress is calculated by dividing
this load by the original cross sectional area.
0.1% Proof Stress will produce permanent set equivalent to one thousandth of the specimen's original
length.
0.2% Proof Stress will produce permanent set equivalent to one five hundredth of the original length.
2.5
STIFFNESS
Within the elastic range of a material, if we relate strain to the stress causing that extension we have a
measure of stiffness/rigidity or flexibility.
ie.
Stress
is a measureof stiffness
Strain
This value which is of great importance to designers is known as ‘the modulus of elasticity, young’s
modulus’, or the symbol E
Thus E = Stress divided by Strain and since strain has no units, the unit for `E' is the same as stress.
i.e. MN/m2 (EUROPE) or P.S.I. (USA)
The actual numerical value is usually large as it is a measure of the stress required to theoretically
doubled the length of a specimen if it did not break first. A typical value of E for steel is 30 x 106 P.S.i.
or 210,000 MN/m2
Relative stiffness values for some common materials using rubber as a datum :
 Wood
2000 x
 Aluminium
10,000 x
 Steel
30,000 x
 Diamond
171,000 x
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2.6
TENSILE TESTING OF PLASTICS
This is conducted in the same way as for metals, but the test piece is usually made from sheet
material. Although the basic load / extension curve for some plastics is somewhat similar to metal
curves, changes in test temperature or the rate of loading can have a major effect on the actual results.
Even though the material under test may be in the elastic range, the specimen may take some time to
return to its original size after the load is removed.
2.7
COMPRESSION TEST
Machines for compression testing are often the same as those used for tensile testing. The test
specimen is in the form of a short cylinder. The Load / Deflection graph in the elastic phase for ductile
materials is similar to that in the tensile test. The value of `E' is the same in compression as it is in
tension. Compression testing is seldom used as an acceptance test for metallic or plastic materials,
except for cast iron. Compression testing is generally restricted to building materials and research into
the properties of new materials.
2.8
HARDNESS TESTING
The hardness of materials is found by measuring their resistance to indentation. Various methods are
used, but the most common being Brinell, Vickers and Rockwell.
2.8.1 BRINELL TEST
In the Brinell Test, a hardened steel ball is forced into the surface of a prepared specimen using a
suitable force for a specified time. The diameter of the resulting indentation is then measured
accurately using a graduated microscope, the hardness number is determined by reference to a BHN
Brinell Hardness No. chart.
2.8.2 VICKERS TEST
The Vickers hardness test is similar but uses a square based diamond pyramid indenter. The
diagonals of indentation are accurately measured by a special microscope, and the Hardness Value
(HV) is again determined by reference to a chart.
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2.8.3 ROCKWELL TEST
The Rockwell Hardness Test also uses indentation as its basis, but two types of indenter are used, a
conical diamond indenter is employed for hard materials and a steel ball is used for soft materials. The
hardness number when using the steel ball is referred to as Rockwell B (e.g. RB 80 ) and the diamond
hardness number is known as Rockwell C (e.g. RC 65 ).
Note: Whereas Brinell and Vickers hardness values are based upon the area of indentation, the
Rockwell values are based upon the depth of the indentation.
No precise relationship exists between the various hardness numbers, but approximate relationships
have been compiled. Some comparative values between Brinell and Rockwell are shown below.
MATERIAL
BHN
HV
ROCKWELL
Aluminium alloy
100
100
B 57
Mild steel
130
130
B 73
Cutting tools
650
697
C 60
There is a good correlation between hardness and U.T.S. on some materials (e.g. steels )
2.8.4 HARDNESS TESTING ON AIRCRAFT
It is not normal to use Brinell, Rockwell or Vickers testing methods on aircraft in the hangar. There are,
however, portable hardness testers which may be used to test for material hardness on items such as
aircraft wheels, after an overheat condition. The overheat condition may cause the wheel material to
become soft or partially annealed.
2.9
IMPACT TESTING
The impact test is designed to determine the toughness of a material. The two types most commonly
used are the ‘charpy’ and ‘izod’ tests. Both use notched bar test pieces of standard dimensions, which
are struck by a fast moving, weighted pendulum. The energy which is absorbed by the test piece on
impact will give a measure of toughness. A brittle material will break easily and will absorb little energy,
so the swing of the pendulum will not be reduced significantly. However, a tough material will absorb
considerable energy, and greatly reduce the pendulum swing. Some materials show a tendency to
lose much of their toughness when the environmental temperature drops.
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Most materials show a drop in toughness with a reduction in temperature. However, some materials
(certain steels in particular) show a rapid drop in toughness as the temperature is progressively
reduced. This temperature range is called the Transition Zone, and parts designed for use at low
temperature should be operated above the material Transition Temperature. Nickel is one of the most
effective alloying elements for lowering the Transition Temperature of steels.
2.10
OTHER FORMS OF MATERIAL TESTING
Although some of the most important forms of material testing have been covered in this section,
several other forms of material testing are carried out. Not least important are Fatigue Testing and
Creep Testing. These are covered in a later section.
2.11
CREEP
Creep can be defined as the continuing deformation with the passage of time, in materials subjected to
prolonged stress. This deformation is plastic and occurs even though the acting stress may be well
below the yield stress of the material.
At temperatures below 0.4T (where T is the melting point of the material in kelvin) the creep rate is very
low, but at higher temperatures it becomes more rapid. For this reason, creep is commonly regarded
as being a high temperature phenomenon associated with superheated steam plant and gas turbine
technology. However, some of the soft, low melting point materials will creep significantly at, or little
above, ambient temperatures and some aircraft materials may creep when subjected to overheat
conditions.
2.12
CREEP IN METALS
When a metallic material is suitably stressed it undergoes immediate elastic deformation, which is then
followed by plastic strain which occurs in three stages:
 Primary Creep - which begins at a relatively rapid rate, but then decreases with time as strain
hardening sets in.
 Secondary Creep - in which the rate of strain is fairly uniform and at its lowest value.
 Tertiary Creep - in which the rate of strain increases rapidly, finally leading to rupture. This final
stage coincides with gross necking of the component prior to failure. The rate of creep is at a
maximum in this phase.
A typical creep curve for a long time, high temperature, creep test is shown below.
2.12.1 EFFECT OF STRESS AND TEMPERATURE ON CREEP
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Both stress and temperature have an effect on creep. At low temperature or very low stress, primary
creep may occur, but this falls to a negligible value in the secondary stage due to strain hardening of
the material. At higher stress and/or temperature, however the rate of secondary creep will increase
and lead to tertiary creep and inevitable failure.
It is clear from the foregoing that short time tensile tests do not give reliable information for the design
of structures which must carry static loads over long periods of time at elevated temperatures, and
strength data determined from long time creep tests (up to 10,000 hours) are therefore essential.
Although actual design data is based on the long time tests, short time creep tests are sometimes used
as acceptance tests.
2.12.2 THE EFFECT OF GRAIN SIZE ON CREEP
Since the creep mechanism is partly due to microscopic flow along the grain boundaries, creep
resistance is improved by increased grain size due to the reduced grain boundary region per unit
volume. It is mainly for this reason that some modern high performance turbine blades are being made
from directionally solidified or even single crystal castings.
2.12.3 CREEP IN PLASTICS
Plastics are also affected by creep and show similar though not identical behaviour to that described
for metals. Since most plastics possess lower thermal properties than metals the choice of plastic for
important applications, particularly at elevated temperature must take creep considerations into
account.
2.13
FATIGUE
An in-depth survey in recent years revealed that over 80 percent of failures of engineering components
were caused by fatigue. A characteristic of modern engineering is an increase in operating stresses,
temperatures and speeds. This particularly so in aerospace, and in many cases has made the fatigue
properties of materials more significant than their ordinary static strength properties.
Engineers became aware that alternating stresses of quite small amplitude could cause failure in a
component which was capable of safely carrying much greater steady loads. This phenomenon of
small, alternating loads causing failure was likened to a progressive weakening of the material, and
hence the name fatigue. Very few constructional members are immune from it especially those
operating in a dynamic environment.
Experience in the aircraft industry has shown us that the stress cycles that aircraft are subjected to may
be very complex with occasional high peaks due to gust loading of aircraft wings. For satisfactory
correlation with in service behaviour, full-size or large scale mock-ups must be tested in conditions as
close as possible to those existing in service.
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An experiment conducted back in 1861 found that a wrought iron girder which could safely sustain a
mass of 12 tons, broke if a mass of only 3 tons was raised and lowered on the girder some 3x106
times. It was also found that there was some mass, below 3 tons, which could be raised and lowered
on to the beam a colossal number (infinite) of times without causing any problem.
Some years later a German Engineer Wohler did work in this direction and eventually developed a
useful fatigue testing machine which bears his name and is still used in industry today. The machine
uses a test piece which is rotated in a chuck and a force is applied at the free end (at right angles to
the axis of rotation). The rotation thus produces a reversal of stress for every revolution of the test
piece.
2.14
FATIGUE TESTING
Various other types of fatigue testing are also used e.g. cyclic torsional, tension-compression etc.
Exhaustive fatigue testing with various materials has resulted in a better understanding of the fatigue
phenomenon and its implications from an engineering viewpoint.
2.15
S-N CURVES
One of the most useful end products from fatigue testing is something called an S-N curve which
shows graphically the relationship between the number of stress cycles (N) and the stress range (S).
It can be seen in the graph below that if the stress is reduced, the steel will endure a greater number of
stress cycles and a point is eventually reached where the curve becomes virtually horizontal, thus
indicating that the material will endure an infinite number of cycles. This limiting stress is called the
fatigue limit and for steels the fatigue limit is generally in the region of 40% to 60% of the value of the
static tensile strength (U.T.S.)
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Many non-ferrous metals show a different characteristic from steel as illustrated in the graph below.
In this case there is no fatigue limit as such and it can be seen that these materials will fail if subjected
to an appropriate number of stress reversals, even at very small stresses. When materials have no
fatigue limit an endurance limit together with a corresponding number of cycles is quoted instead. It
follows that components made from such materials must be designed with a specific life in mind and
removed from service at the appropriate time. The service fatigue life of complete airframes or
airframe members are typical examples of this philosophy.
Non-metallic materials are also liable to failure by fatigue. As in the case of metals the number of
stress cycles required to produce a fatigue failure increases as the maximum stress in the loading cycle
decreases. However, there is generally no fatigue limit for these materials and some form of
endurance limit must be applied.
The importance of fatigue strength can be illustrated by the fact that in a high cycle fatigue mode, a
mere 10% improvement in fatigue strength can bring about 100 times life improvement.
2.16
CAUSES OF FATIGUE FAILURE
As the fatigue characteristics of most materials are now known or can be ascertained, it would seem
reasonable to suppose that fatigue failure due to lack of suitable allowances in design should not
occur. Nevertheless fatigue cracking occurs frequently and even the most sophisticated engineering
product does not possess immunity from this mode of failure. Such failures are often due to
unforeseen factors in design, material, manufacturing, environmental conditions, or operating
conditions.
Two essential requirements for fatigue development in a material are :
 An applied stress fluctuation of sufficient magnitude (with or without an applied steady stress).
 A sufficient number of cycles of that fluctuating stress.
The stress fluctuations may be separated by considerable time intervals, e.g. aircraft cabin
pressurisation on each take off (say daily) or they may have a relatively short time interval e.g. the
aerodynamic buffeting/ vibration of a wing panel. The former would be called low cycle fatigue and the
latter be high cycle fatigue.
In practice the level of the fluctuating stress and the number of cycles to cause cracking of a given
material are affected by many other variables such as stress concentration, residual stresses,
corrosion, surface finish, material imperfections etc.
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2.17
VIBRATION
This has already been quoted as a cause of high cycle fatigue and because most dynamic structures
are subjected to vibration this is undoubtedly the most common origin. All objects have their own
natural frequency at which they will freely vibrate (i.e. resonant vibrations). Large, heavy flexible
components vibrate at low frequency. Small, light stiff components vibrate at high frequency.
Resonant frequencies are undesirable and in some cases could be disastrous, so it is important to
ensure that over the normal operating range, that critical components are not shaken at their natural
frequencies to create resonance. A simple everyday example of a resonance is the wheel wobble
present in a car at a certain road speed. The frequency of a component is governed by its mass and
stiffness. On certain critical parts it is sometimes necessary to do full scale fatigue tests to confirm
adequate fatigue life before putting the product into service.
2.18
FATIGUE METALLURGY
Under the action of fatigue stresses minute, local plastic deformation on an atomic scale takes place
along slip planes within the material grains. If the fatigue stresses are continued, micro cracks are
formed within the grains in the area of the highest local stress, (usually at or near the surface of the
material). The micro cracks join together and propagate across the grain boundaries but not along
them.
A fatigue fracture generally develops in three stages:
 Nucleation
 Propagation (crack growth)
 Ultimate (rapid) fracture.
The resultant fractured surface often
has a characteristic appearance of:
 An area on which a series of curved, parallel relatively smooth ridges are present and are centred
around the starting point of the crack. These ridges are sometimes called conchoidal lines or
beach marks or arrest lines.
 A rougher typically crystalline section which is the final rapid fracture when the cross section is
no longer capable of carrying its normal, steady load.
The arrest lines are normally formed when the loading is changed, or the loading is intermittent.
However, in addition to these characteristic and informative marks, there are similar but much finer
lines called striations which literally show the position of the crack front after each cycle. These
striations are obviously of great importance to metallurgists and failure investigators when attempting to
estimate the crack initiation and/or propagation life. The striations are often so fine and indistinct that
electron beam microscopes are required to count them.
In normal circumstances a great deal of energy is required to `weaken' the material sufficiently to
initiate a fatigue crack, and it is not surprising therefore to find that the nucleation phase takes a
relatively long time. However, once the initial crack is formed, the extremely high stress concentration
present at the crack front is sufficient to cause the crack to propagate relatively quickly, even gaining in
speed as the crack front not only increases in size, but it also reduces the component cross sectional
area. A point is eventually reached (known as the 'critical crack length') at which the remaining cross
section is sufficiently reduced to cause a gross overloading situation, and a sudden fracture finally
occurs.
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It is not unusual for the crack initiation phase to take 90% of the time to failure, with the propagation
phase only taking the remaining 10%. This is one of the major reasons for operators of equipment
being relatively unsuccessful in detecting fatigue cracks in components before a failure occurs.
2.19
FATIGUE PROMOTERS
 As fatigue cracks initiate at locations of highest stress and lowest local strength, the nucleation
site will be:
 dictated largely be geometry and the general stress distribution.
 at or near the surface.
 centred on surface defects/imperfections such as scratches, pits, inclusions, dislocations and the
like.
2.19.1 DESIGN
Apart from general stressing, the geometry of a component has a considerable influence on its
susceptibility to fatigue. A good designer will therefore minimise stress concentrations by:
 avoiding rapid changes in section and
 using generous blend radii or chamfers to eliminate sharp corners.
2.19.2 MANUFACTURE
Even if the designer specifies adequate blend radii, the actual product may still be prone to fatigue
failure if the manufacturing area fails to achieve this sometimes seemingly unimportant drawing
requirement. Several other manufacturing related causes of premature fatigue failure exist, the most
common of which, are
 Material defects e.g. inclusions, porosity, cold shuts, forging defects etc.
 Undetected cracks. e.g. heat treatment, shrinkage, bending, grinding, forming etc.
 Residual stresses. e.g. undue force used during assembly.
 Local damage. e.g. tooling marks, scores, scratches, dents, impact marks etc.
2.19.3 ENVIRONMENT
One of the most potent environmental promoters of fatigue is if the component is operating in a
corrosive medium. Although steel normally has a well defined fatigue limit on the S-N curve, if a fatigue
test is conducted in a corrosive environment, not only does the general fatigue strength drop
appreciably, but the curve then resembles the aluminium alloy curve i.e. the fatigue failure stress
continues to fall as the number of cycles increase. Other environmental effects such as fretting,
corrosion pitting, erosion or elevated temperatures will also adversely affect fatigue strength.
2.20
FATIGUE PREVENTERS
If a component is prone to fatigue failure in service several methods of improvement are available to us
:
 Quality.
Correct and eliminate any failure-related manufacturing or processing shortcomings.
 Material.
Select a material with a significantly better fatigue strength, or corrosion resistance or
corrosion protection if relevant.
 Geometry.
i. Increase the size (C.S.A.) to reduce the general stress level.
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ii. Modify the local geometry to reduce the change in section (large radius).
iii. Modify the geometry to change the vibration frequency.
iv. Introduce a damping feature to reduce the vibration amplitudes.
 Surface condition.
i. Improve the surface finish.
ii. Put a compressive stress in the skin e.g. shot peen.
2.21
PRACTICAL DO'S AND DONT'S TO HELP PREVENT FATIGUE FAILURES
DO’S
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/near contact of components which 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 correctly heat treated state.
2.22
FATIGUE MONITORING (MODERN TRENDS)
Use of fatigue meters to check overall stress levels on aircraft and to monitor the fatigue history of the
aircraft. Fatigue meters also allow us to check when the aircraft exceeds the design limits imposed on
it.
Use of strain gauges to monitor stress levels on specific aircraft structures. Strain gauges are thin foil
resistance elements bonded to aircraft structure. The resistance varies as the load varies.
2.22.1 FATIGUE FUSES
A metal fuse is bonded to the structure. The fuse is bonded and fails at different fatigue stresses. The
current flowing through the fuse varies and gives an indication of the stress level.
2.23
INTELLIGENT SKINS DEVELOPMENT
Modern developments in aircraft structures will allow the structures to be designed and built with a
variety of sensors and systems to be embedded into the 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.
The generic heading Smart Structures covers three area's:
 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 structure skin.
 Intelligent Skins
A skin embedded with fibre optic sensors.
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Smart Structures perceived benefits include:
 Self diagnostic to monitor structural integrity.
 Reduced life cycle costs.
 Reduced inspection costs.
 Potential weight savings/performance increases from increased knowledge of composite material
performance.
 Stealth characteristics improvement.
A fully monitored and self diagnostic system could:
 Assess structural integrity.
 Pin-point structural damage.
 Process flight history.
2.23.1 STRUCTURAL HEALTH MONITORING (SHM)
Composite laminates containing embedded fibre optic sensors can be used for SHM, including fatigue
monitoring and flight envelope exceedance monitoring. SHM advantages include:
 Covers 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
Detection and assessment of impact damage is possible using a mesh system of fibre optic sensors
which transmit less light when damaged. This system could be of use to both the pilot and ground
crew. These applications are both long and short term and may have inherent problems.
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2.24
COLD WORKING
Most fatigue failures occur whilst a material is subject to a tensile alternating stress. If the most fatigue
prone area, such as spar fastener holes, have a compression stress applied they are significantly more
resistance to fatigue failure. The fastener hole is checked for defects and surface finish improved by
reaming , a close fitting sleeve is fitted with the hole and a mandrel palled through it. The result is a
localised area which has a residual (compressive) stress and tensile force will, in this localised area,
result in neutral or at least significantly reduced level of tensile stress.
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3.
FERROUS METALS
Ferrous metals are metals that have iron as the main element. The most commonly used ferrous metal
is steel which use to manufacture bolts, bearings and engine components on aircraft.
3.1
PROCESSING THE RAW MATERIAL
Iron is one of the most common elements in the earth's crust. It comprises approximately 5%
compared with aluminium at 8%. It is never found naturally in its metallic state, but as iron ores which
contain in the range of 25% to 60% iron and are mined in open cast or open pit mines. Iron has a great
affinity (attraction) for oxygen.
3.2
NODULAR CAST IRON
This is a more modern development and is sometimes known as ‘Spheroidal Graphite Iron’. It is
produced by adding magnesium and nickel or magnesium, copper and silicon and is a tough, strong
hard-wearing material which can be used in applications where only wrought materials were used in
the past, a classic example being automobile crankshafts.
3.3
STEEL
Steel is essentially an alloy of iron and less than 2.5% carbon, usually with a few impurities. (In
practice, most steels do not have more than 1.5% CARBON)
Steel is produced by refining pig iron by removing excess carbon and other unwanted impurities. The
excess carbon is extracted by blowing oxygen or air through the molten metal, and/or adding iron
oxide. Slag containing other impurities are skimmed off. The most common furnace used for this
process was called the ‘Bessemer Convertor’, developed in 1856. It reduced the cost of steel to one
fifth of it's original cost. Bessemer converters were loaded with 20 - 50 tons of pig iron and air was
blown from the bottom for approximately 15 minutes.
3.3.1 CARBON IN STEEL
When carbon is alloyed with iron, the hardness and strength of the metal increases; for example, a
steel containing 0.4% carbon may be twice as strong as pure iron. The effect of varying amounts of
carbon is truly dramatic. If carbon is progressively added to pure iron the following occurs:
 Initially, the strength and hardness increases - (Steel containing 0.4% carbon has twice the
strength of pure iron.
 However, if 1% of carbon is added, the strength and hardness show a further increase but
ductility is reduced.
 If 1% to 1.5% of carbon is added, the hardness continues to increase, but there is no further
increase in strength and there is even less ductility. Steels containing such high amounts of
carbon are seldom used for anything except cutting implements e.g. razor blades & scissors
Typical amounts of carbon present in iron and steel are as follows:

mild steel (Low Carbon)
up to 0.25%
Low carbon steel is frequently used for wire, low tensile nuts, bolts, screws and structures.

Medium carbon steel
0.25% to 0.45%
Medium carbon steel is used for strong or dynamic applications such as shafts and gears.

High carbon steel
0.45% to 1.5%
High carbon steel is used for cutting tools e.g. knives, low quality drills and taps.

Cast iron
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Page 3-1
Cast iron is used in aero engine components.
3.3.2 METALLURGICAL STRUCTURE OF STEEL
The amount of carbon present in steel has a major effect on the mechanical properties. The form in
which the carbon is present is also important.
3.3.3 THE STRUCTURE AND PROPERTIES OF SLOWLY COOLED STEELS
Carbon can be present in these steels in the following forms:
 When the carbon is fully dissolved and therefore uniformly distributed in a solid solution, the
metallurgical structure is called ferrite. At room temperature only a very small amount of carbon
(0.006%) can be contained in solid solution, therefore this ferrite structure is almost pure iron. It
is not surprisingly - soft, weak and ductile.
 When 1 carbon atom chemically combines with 3 iron atoms the result is called cementite or iron
carbide. It is very hard and brittle.
 Cementite can be present either as free cementite or laminated with ferrite (in alternate layers) to
produce a metallurgical structure called pearlite. As pearlite is half cementite and half ferrite, it is
not surprising to find that pearlite combines the properties of ferrite and cementite. I.e. Whereas
ferrite was too soft and weak, and cementite was basically strong but too hard and brittle,
pearlite is strong without being brittle.
The amount of carbon necessary to produce a totally pearlite structure is 0.83%. Such a material is a
bit too hard for general structural use. If the carbon content exceeds this value, the excess carbon
forms carbon rich cementite areas along the grain boundaries, and this is known as free cementite.
Such high carbon steels as already stated are very hard, strong but very brittle.
Mild steel has a metallurgical structure comprising approximately one third Pearlite and two thirds
Ferrite. The accompanying diagrams show the structures and uses for Ferrite, ferrite and pearlite, all
pearlite, and pearlite plus free cementite.
3.3.4 THE EFFECT OF COOLING RATES ON CARBON STEELS
Previously we have considered the effect of carbon on the properties of a slowly cooled steel. If such
steels are rapidly cooled from relatively high temperature the metallurgical structure and properties can
be somewhat different.
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3.4
HEAT TREATMENT OF CARBON STEELS
If a straight carbon steel is progressively heated from cold, a steady rise in temperature occurs.
However, at approximately 700C there is a reduction in the rate of temperature rise even though the
heating is continued. Eventually, the temperature rise speeds up and the rate of rise is similar to that,
which occurred before the `hesitation'. This `hesitation' starts at 700C and finishes at up to 200C
higher (depending on the percentage of carbon present).
The start of this hesitation is known as the ‘lower critical point’ and the end is called the ‘upper critical
point’, and the phenomenon of the temperature response is due to a change in the crystalline structure
of the steel in between the two critical points. If carbon steel is heated just above its Upper Critical
Point the structure is called ‘Austenitic’. This structure is a solid solution of carbon in iron (i.e. all the
carbon is uniformly distributed throughout the iron). If the steel contains above 0.3% carbon, and it is
rapidly cooled (i.e. quenched) from above the Upper Critical Point it becomes hardened.
The more carbon present, the harder the steel will be after quenching. This rapid cooling causes a
change in the metallurgical structure and is called ‘Martensite’. Martensite is extremely hard but is not
suitable for most engineering purposes due to it being very brittle. For most applications it is necessary
to carry out a further heat treatment to reduce the brittleness of the steel, and this is called ‘tempering’.
To temper hardened carbon steel it is necessary to heat it to a suitable temperature below its Lower
Critical Point followed by cooling (usually quenching). The effect of this heat treatment is to slightly
reduce the hardness whilst at the same time greatly increasing the toughness. The actual tempering
temperature used depends on the requirements of strength, hardness and toughness. The higher the
tempering temperature, the lower will be strength and hardness, but the toughness will be greater. The
maximum tensile strength of hardened carbon steel is achievable when 0.83% carbon is present. If an
even greater amount of carbon is present, the hardness continues to increase but strength will
decrease.
3.4.1 PROBLEMS ASSOCIATED WITH THE HARDENING PROCESS
The effective hardening of carbon steels depends not only on the amount of carbon present but also on
very rapid cooling from high temperature. The cooling rate mainly depends on the cooling medium, the
size of tank, and the mass of the object to be cooled. Agitation in the cooling bath can also speed up
the cooling rate. In terms of cooling severity, brine is more effective than water, followed by oil and
finally air.
Carbon steels require an extremely rapid cooling phase, so brine or water are normally used, whereas
oil or air cooling is used on certain alloy steels. The rapid cooling rates involved in the hardening of
carbon steel cause enormous thermal stresses in the component and distortion is commonplace.
Cracking may also occur in some cases. To achieve relatively uniform cooling it is sometimes
necessary to immerse the object in a specific way because of its shape and mass.
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3.4.2 TEMPERING
As already stated, is carried out to improve the toughness of hardening steel whilst suffering only a
modest drop in strength. Accurate temperature measuring equipment is normally used in well
equipped facilities, but a simple method of using tempering colours can be used reasonably
successfully when such facilities are not available. When carbon steel is polished to a bright, clean
surface, and then slowly heated, a range of colours appear due to a thin oxide film forming during the
heating process. These colours are related fairly closely to temperatures. The higher temperature
achieved during the tempering process, the softer (and tougher) the material will become and viceversa.
COLOUR
TEMPERATURE
Straw
230/240c
Purple
270c
Blue
300c
Dark red
500C
3.4.3 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 treatment (rolling, forging, uneven cooling).
 To remove coarseness of grain.
Annealing is normally achieved on carbon steel by heating to just above the Upper Critical Limit
followed by very slow cooling. In practise the slow cooling rates are achieved by cooling in the furnace
or by immersing in a poor thermal conductor such as ashes. The end result is a stress free, fully
softened material suitable for major forming operations such as deep pressing, drawing, extruding etc.
3.4.4 NORMALISING
This process is similar to annealing except that the cooling is done in still air. The end result is a stress
free, soft material with uniform fine grain structure. Normalising is commonly used on actual
components after heavy machining operations or welding prior to the final hardening and tempering
processes.
3.5
SURFACE HARDENING OF STEELS
Unlike conventional through-hardening of steel, it is sometimes desirable to retain a relatively tough
(relatively soft) inner -core coupled with a very hard surface. This would typically be the case of a
component subjected to high dynamic stresses, which also had to resist surface wear. Some materials
can be case-hardened to achieve this aim. Several methods are used, depending on the parent
material and the particular application. Components suitable for this case-hardening treatment would
include gears (where the teeth need to be hardened), camshafts and crankshafts (bearing surfaces)
and cylinder barrels on piston engines.
3.5.1 CARBURISING
This is the most common method of case hardening low carbon steels and basically consists of heating
to approximately 900C with the component in intimate 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 of penetration is approximately 1 mm in 5-6 hours. Low carbon steels are particularly suited to this
type of treatment as it increases the carbon content and hence the hardness locally. Various methods
of carburising are used, the most common ones being:
MODULE 6 - Materials and Hardware
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 Pack Carburise. The object is sealed in a container containing a carbon rich (charcoal based)
powder and heated in a furnace.
 Gas Carburise. The object is placed in a basket in a furnace, through which is passed a suitable,
carbon rich gas (e.g. methane, propane).
 Liquid Carburise. The object is heated to a suitable temperature and then immersed in a hot salt
bath at 900C. The salts are usually based on sodium cyanide and the process is often called
cyanide hardening.
3.5.2 NITRIDING
This process involves the absorption of nitrogen (instead of carbon) in the surface of the steel.
Suitable "Nitralloy" steels are necessary for this process and usually containing 1% Aluminium, 1.5%
Chromium and 0.2% Moybdenum. A special furnace is used and ammonia gas is then circulated
through it. The furnace temperature of 500C converts the ammonia into a nitrogen rich gas and forms
hard iron-nitride in the surface of the steel. The case-depth achievable by this process is less than that
by pack carburising, but the major advantage of nitriding is that no hardening or tempering is necessary
to achieve the final hardness, and no finish machining is required after nitriding. The relatively low
temperature used results in negligible distortion. Aircraft piston engine cylinder barrels are often
nitrided, as are some crankshaft bearing surfaces and the stems of some aero engine valves.
3.5.3 AREAS TO BE LEFT UN-HARDENED
If certain surfaces of a component are not to be case hardened it is necessary to protect them during
the carburising or nitriding processes to locally prevent the hardening agent from being absorbed.
Copper plating, nickel plating or a proprietary paste are generally used in such areas.
3.5.4 FLAME / INDUCTION HARDENING
Unlike carburising and nitriding, flame and induction hardening do not add a hardening agent into the
surface of a basically softer material. Instead, they are merely techniques for hardening the surface of
material by a `local heat treatment'. Steels suitable for these processes already contain sufficient
carbon (or other elements) to attain a high degree of hardness if heated and quenched, but only the
surface is locally heated by a flame or electrical induction coil and the heated surface is then
immediately quenched by water jets. The flame or induction coil is positioned so that it only heats the
area required to be hardened.
3.5.5 OTHER HARD SKIN TECHNIQUES
In addition to case hardening, there are other methods of producing hard surfaces on metals such as
by electro-plating, welding, bonding, metal spraying. All usually involve adding a harder surface to the
material.
3.6
ALLOY STEELS
The basic properties of pure metals and the basic reasons for alloying are covered elsewhere. Some
of the more common alloy steels and their major properties will now be covered.
3.7
NICKEL STEELS
Nickel is used extensively for alloying with steel as follows:
 In the range of 1% - 5% there is a marked improvement in strength (and hardness) without
lowering ductility. This high strength, tough steel is widely used for highly stressed parts.
 At about 25% nickel, the steel becomes highly corrosion resistant, heat resistant and
non-magnetic.
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 At 36% nickel, a unique steel is created call ‘invar’. This has the lowest coefficient of expansion
of any metal, 1/20th of steel. It is excellent for master gauges and instruments.
 Because of the effect of such amounts of nickel on the expansion properties of steel, a range of
nickel-steels marketed under the name of `nilo' are purpose made to trim the coefficient of
expansion to specific needs. These alloys are used in thermostats, spark plug electrodes etc.
3.8
NICKEL ALLOYS
When the amount of nickel present is predominant, the material becomes known as a Nickel alloy,
some of which are widely used in industry. One of the most important nickel based alloy groups is
Nimonic. These are a family of alloys containing 50% - 80% nickel with the balance being mainly
chrome and have excellent high temperature properties. They are used extensively in the hot section
of gas turbine engines. Nimonic alloys are also used in the Harrier hot air control ducting, mainly
because of it's extremely low coefficient of expansion at elevated temperatures (450C). Typical
Nimonic alloys are Nimonic 75 which is a 78% Nickel 20% Chromium, suitable for combustion tubes.
Nimonic 80, 90, 105 and 115 are similar nickel-chromium alloys used for turbine blades in the
temperature range 700 - 1,000C.
Other ranges of nickel based alloys come under the names of Inconel and Hastelloy which are
temperature resistant and corrosion resistant. Another common nickel alloy is Monel. This metal
which has 66% nickel and 33% copper, has excellent corrosion and chemical resistance, is tough,
ductile, reasonably strong (equivalent to mild steel) and is non-magnetic. It is used in lots of marine
applications, for surgical apparatus and for aircraft rivets. Monel does not respond to heat treatment,
but when alloyed with a small amount of aluminium (2% to 4%), it can be hardened to double its
strength. This version is known as ‘K-Monel’.
3.9
CHROME STEEL
When small amounts of chrome are added to steel, the strength and hardness increase, but there is
some loss of ductility. 1.5% chrome in a high carbon (1%) steel results in a very hard material which is
used extensively for instrument pivots and ball and roller bearings. Low chrome (1.5%-3%) steels are
used for high tensile fasteners and are also suitable for nitriding. Steels containing 12% or more
chrome are very corrosion resistant. Stainless steel comes into this category. One particular Stainless
Steel is designated 18/8 Stainless. It contains approximately 18% Chromium and 8% Nickel. These
Stainless steels are used extensively in engine parts, particularly for hot applications and for exhaust
areas where it's corrosion resistance is vital.
3.10
NICKEL-CHROME STEELS AND ALLOYS
This term is used when the amount of nickel present is greater than the chrome content. A wide range
of such steels exist, but the low nickel-chrome ones are suitable for through hardening or case
hardening. The nickel content is around 3-5% and the chrome ranges from 0.5 - 1.5%. Crankshafts
and con-rods are often made from this group. High nickel-chrome alloys (65-85% nickel 15-20%
chrome) have a high electrical resistance and are often used as heater elements.
3.11
MANGANESE STEEL
When small amounts of manganese are added to steel (up to 1.5%) the result is a steel which is strong
and hard (similar to nickel chrome steel). Such steel is often used for shafts and axles. If the
manganese content is in the range 1.5% to 10% the steel becomes extremely brittle and is of no
practical use. However, 12% manganese steel has very unusual properties and is extremely useful.
When this material is heated to approximately 1000C and water quenched, its structure becomes
austenitic and although it is only moderately hard, any attempt to cut it or abraid it results in the local
formation of hard martensite and it thus becomes highly resistant to cutting or abrasion. Because of
this peculiar property, it is used extensively for rock drills, stone crushers, rail lines at junctions etc..
Small amounts of manganese are used in steel production and in welding rods since it acts as a
purifying agent by reducing oxidation.
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3.12
TUNGSTEN STEEL
Tungsten is a very dense, hard element with the highest melting point of all metals. It is a very
important metal in High Speed Steels (H.S.S. a range of purpose designed cutting steels). Such steels
will retain considerable hardness and strength even when operating at dull-red heat. They can be used
to cut materials at twenty times the rate at which hard plain-carbon steels can achieve.
In addition to being used for a wide range of cutting tools, tungsten steels are also used for extrusion
dies.
A typical High Speed Steel would contain 14-18% tungsten, 4% chrome, and 0.6% carbon. Tungsten
is also used to form the incredibly hard cutting material Tungsten-Carbide which is widely used for
tipped tools. Tungsten carbide consists of tungsten powder bonded in a cobalt matrix.
3.13
COBALT STEEL AND ALLOYS
Cobalt is often included in High Speed Steels in addition to tungsten to improve still further the ability to
cut at high working temperatures. Cobalt is often used in high strength, permanent magnets, in some
of the Nimonic alloys and for high temperature components in gas turbine engines. Cobalt is also used
in a range of temperature resistant alloys called ‘Stellite’ (used in piston engine valves). As previously
stated, cobalt is also used as the matrix (bonding) material for tungsten-carbides.
3.14
VANADIUM STEEL
When added to steel, vanadium improves the strength without loss of ductility, but also greatly
improves its toughness and its resistance to fatigue. Valve springs usually include vanadium. Small
amounts of vanadium are often included in certain nickel-chrome steels.
3.15
MARAGING STEELS
Conventional very high tensile steels have a high carbon content and thus are very hard and difficult to
work and also tend to be somewhat brittle. To combat these shortcomings, maraging steels were
developed. These steels are over 50% stronger than normal high tensile steels and yet are very tough
and easy to machine. These properties are achieved by the almost total elimination of carbon and by
alloying with nickel, cobalt and molybdenum in such a way that it can be precipitation hardened.
Maraging steel can only be used for special, high stressed applications due to cost, which is about
three times that of conventional alloy steels. They are used for some airframe and engine components
and can be nitride hardened.
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4.
NON FERROUS METALS
Non ferrous metals are normally metals which are either aluminium or magnesium based. However,
titanium is a more modern `light' metal to gain engineering acceptance and increasing use..
4.1
PURE ALUMINIUM
Pure aluminium is extracted from the mineral rock bauxite. It is soft, weak, ductile and malleable. It is
approximately one third the weight of steel and approximately one third the stiffness of steel. Its
strength may be improved by cold work but it is still low strength material. It is highly corrosion
resistant due to the rapid formation of a thin, but very dense oxide film which limits further corrosion.
4.2
ALUMINIUM ALLOYS
To achieve medium / high strength properties, aluminium must be alloyed. The three most common
alloying elements in the wrought aluminium alloys are copper, magnesium and zinc. A common
element used when casting aluminium is silicon. Some aluminium alloys are heat treated to obtain
high strength Others which do not respond in this way are work hardened or strain hardened to
increase their strength.
4.3
ALUMINIUM-COPPER ALLOYS
When around 4% of copper is used, the material is often referred to as ‘Dural’ (Duralumin). When
suitably heat treated it is stronger than mild steel. Dural is used extensively in all parts of aircraft
construction, including the skin, frames and stringers. Dural is also a heat treatable alloy.
4.4
OTHER AGE HARDENING ALUMINIUM ALLOYS
Although the aluminium copper alloys (dural type) are the most common age hardening, high strength
materials, they are not unique. Aluminium when alloyed with 5% - 7% Zinc is also able to be agehardened and can be much stronger than dural. This is a more modern material than aluminium copper and is the highest strength aluminium alloy in general use. It is used in heavy loaded
applications such as Main Spars, Undercarriage and Mainplane Attachment brackets etc.. The three
basic types of heat treatment for dural are also used on the aluminium zinc alloys, i.e. Solution,
Precipitation and Annealing, although the actual temperatures will differ.
4.5
CLAD MATERIALS
Dural and similar strong aluminium alloys are not as resistant to corrosion as pure aluminium and for
external use such as skin, the high strength sheet has a thin layer of pure aluminium hot-rolled onto the
surfaces. These are then known as clad materials with commercial names such as Alclad, and
Pureclad. The thin coating is usually about 5% of the sheet thickness.
4.6
WORK / STRAIN HARDENING ALUMINIUM ALLOYS
Some aluminium alloys cannot have their strength improved by heat treatment and are for this reason
are sometimes referred to as "Non-heat treatable". However, when suitably alloyed, materials in this
group may develop reasonably high strength by controlled cold working. The most common of these
strain hardening materials is the aluminium-magnesium group. These are produced (mainly sheet or
plate) in a range of different hardness values or so-called ‘tempers’ by a suitable amount of working /
strain hardening. Sheets for example may be obtained in the annealed, ¼ hard, ½ hard, 3/4 hard or
fully hard state. The aluminium magnesium group are also good welding alloys, whereas the dural
group are not suitable for welding.
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4.7
ALLOY IDENTIFICATION SYSTEMS
4.7.1 IADS (INTERNATIONAL ALLOY DESIGNATION SYSTEM)
This is the most generally accepted modern identification system in use and is progressively replacing
previous varied numbering systems. Wrought aluminium alloys predominate in aircraft construction
and they are identified by a basic 4 digit numbering system e.g. lxxx, 2xxx etc. In this system `Pure'
aluminium and the aluminium alloys are basically identified by the first digit i.e.
1xxx Series
Pure (i.e. non-alloyed) aluminium.
2xxx Series
Aluminium-copper group.
5xxx Series
Aluminium-magnesium group.
7xxx Series
Aluminium-zinc group.
It is the Major alloying element which determines the series into which the particular alloy is listed).
The 3xxx, 4xxx, 6xxx and 8xxx series adopt the same principle, but these alloys are not in general use
in the aircraft industry and are therefore not listed in these notes.
The significance of the other 3 digits differs for non-alloyed aluminium and the aluminium alloys as
follows:
 Non-alloyed wrought aluminium. The-last two digits indicate the actual purity of the metal i.e.
these two digits give the numerical value of pure aluminium above a nominal datum level of
99%. The second digit relates to the control of impurities present. e.g.
1100 is 99% pure aluminium
1130 is 99.3% pure aluminium
1230 is 99.3% pure aluminium
1185 is 99.85% pure aluminium
1385 is 99.85% pure aluminium
 Alloyed wrought aluminium alloy. The third and fourth digits refer to the specific alloy,
whereas the second digit indicated a specification change e.g.
i.
2117 is an aluminium-copper alloy, was the 17th aluminium copper alloy catalogued and
has had one specification amendment.
ii. 7205 is an aluminium-zinc alloy, was the 5th aluminium-zinc, alloy catalogued and has had
two specification changes.
4.8
MATERIAL CONDITION
Apart from the basic material specification covering the chemical composition the actual condition of
the material, i.e. the hardness and heat treatment state are also relevant pieces of information. This
important data, is indicated as follows:
 Heat Treatable Alloys
A suffix T followed by a number (1-10) is added to the material specification to denote the heat
treatment condition of the material i.e.
2014 - T4 (T4 means it has been solution heat treated, followed by natural ageing). 2014 -T6 (T6
means it has been solution heat treated, followed by artificial ageing)
 Non-Heat Treatable Alloys
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 The suffixes used for these materials are `O' (annealed) or `H', followed by a number to indicate
the degree of work or strain hardening e.g. 5017-HXX The first digit after the H indicates
whether any heat treatment has been carried out. The second digit after the H indicates the
degree of strain hardening. (No's 1 to 8 are used, 1 being the least strain hardening and 8 being
the most).
4.9
HIX -
Strain hardened - no heat treatment.
H3X -
Strain hardened & partially annealed.
HX2 -
¼ Hard.
HX4 -
½ Hard.
HX6 -
3/4 Hard.
HX8 -
Fully hard.
BASIC TEMPER DESIGNATIONS
4.10
F
-
As fabricated (i.e. no treatment)
O
-
Annealed.
H
-
Strain hardened.
W
-
Solution heat treated (this is a temporary condition because the material ages
spontaneously).
T
-
Thermally treated to produce particular properties.
EXAMPLES OF IADS CODES
 Pure Aluminium example
1185
‘Pure’
Aluminium
Control over
impurities
99.85% ‘Pure’ Aluminium (1140
would be 99.40% pure).
 HEAT TREATABLE ALUMINIUM ALLOY EXAMPLE
2117 - T6
Aluminium-Copper alloy One revision of
(copper-prime alloying
specification
element) artificial
ageing
Specific alloy
Heat treatment
condition t6 is
solution and artificial
ageing
 NON HEAT TREATABLE ALUMINIUM ALLOY EXAMPLE
5234 – H18
Material State
Degree of strain
hardening (8 is
Max)
Aluminium Magnesium
Two revisions of
MODULE 6 - Materials and Hardware
Specific alloy
No heat treatment
Page 4-3
alloy Magnesium –
prime alloying element
specification
MODULE 6 - Materials and Hardware
Page 4-4
Alloy Identification Systems
Temper Designations
(added as suffix letters or digits to the alloy number)
Suffix letter F,H,O,T
indicates primary
treatment or condition
F
O
as
fabricated
annealed
cold
worked
H
Four Aluminium content or
main alloying
digit
elements
series
1\ \ \
99% minimum
2\ \ \
copper
3\ \ \
maganese
4\ \ \
silicon
5\ \ \
magnesium
6\ \ \ magnesium & silicon
7\ \ \
8\ \ \
First suffix digit
indicates secondary
treatment used to
influence properties
1
2
3
Cold
worked &
stabilised
T
Heat
treated
MODULE 6 - Materials and Hardware
Partial
solution
plus
natural
ageing
2
¼ Hard
4
½ Hard
6
¾ Hard
8
Hard
9
Extra hard
2
Annealed (cast
products only)
3
Solution plus cold
work
4
Solution plus
natural ageing
5
Artificially aged
only
6
Solution plus
artificial ageing
7
Solution plus
stabilising
8
Solution plus cold
work plus artificial
ageing
9
Solution plus
artificial ageing
plus cold work
zinc
others
cold
worked
only
Cold
worked &
partially
annealed
1
Second suffix digit
for condition
(H only indicates
residual hardening)
Page 4-5
4.11
OTHER NUMBERING SYSTEMS
Apart from the IADS system, many other exist worldwide but the British systems are basically confined
to three basic ones for light alloys.
 A British Standard for general engineering use - BS 1470 - 1475. In this series the prefix N is
used to denote non-heat treatable aluminium alloys and prefix H for the heat treatable alloys.
 British Standards for aerospace use - BS X LXX. (The "L" series) e.g. BS 3 L72 indicates the 3rd
amendment to the basic L 72 spec. LM - indicates a cast material. The wrought materials are
commonly abbreviated to L 71, L72, L 73 etc.
Examples of some of these aircraft BS codes are:
i. L72 ALCLAD
Solution Treated - Naturally aged
ii. L73
Solution Treated - Artificially aged
ALCLAD
Note: Both of these are very old codes (but still used).
iii. L159 DURAL
Solution Treated - Artificially aged
iv. L163 ALCLAD
Solution Treated - Naturally aged
 D.T.D. Specifications - these are material identification numbers issued by the Directorate of
Technical Development (a Ministry Department) for specialised applications. i.e. when
widespread use is not anticipated. If such a material finally becomes commonly used, a British
Standards specification is compiled and issued.
4.12
MARKING OF ALUMINIUM ALLOY SHEETS
Sheet material for aero use is marked with the specification number every few inches usually in a blue
ink e.g. ‘7075 - T6’. Some sheets may also have alternate lines of red numbers, which indicate that
heat treatment is needed before assembly. These red numbers then disappear when the necessary
heat treatment is done. The basic type of aluminium alloy can sometimes be determined by dipping a
sample in a 10% caustic soda solution. The dural types turn black due to the copper content, whereas
the others tend to stay bright. Clad dural (Alclad, Pureclad etc) is readily identified by this check if the
edge of the metal is treated with the caustic solution.
4.13
CAST ALUMINIUM ALLOYS
These are not used extensively on airframes mainly due to:
 lack of strength
 poor fatigue strength, and
 lack of ductility when compared to the wrought alloys.
The lack of ductility is particularly relevant as the very nature of an airframe structure requires the
ability to flex considerably without cracking. Although their use is obviously limited on airframes, cast
aluminium alloys are used extensively on engines, where there is a need to produce complex cored
shapes such as crankcases, drive casings, cylinder heads etc. No other method than casting would be
viable for such items. The stresses can be kept to a modest level on these parts by producing robust
casings of adequate stiffness.
Very few non-heat treatable cast alloys are used on aero 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 containing
approximately 2½% copper, l½% magnesium, 1½% nickel, l% iron. (A derivative of this material is also
used in wrought form for the skin of the supersonic Concord aircraft due to the high metal temperatures
encountered).
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Aluminium alloy castings often contain silicon as it creates high fluidity and thus is good for producing
complex shapes. It also reduces the coefficient of linear expansion, so is often included in piston
castings.
4.14
MAGNESIUM & IT'S ALLOYS
This is the lightest structural metal at only two thirds the weight of aluminium. Magnesium is found in
ores, but is mainly produced from sea water. Initially an extraction of magnesium chloride is made, and
an electrolytic process finally produces the metallic magnesium. Although extremely light, pure
magnesium is also weak and has very poor corrosion resistance. However, when suitably alloyed it
has a higher strength to weight ratio than the aluminium alloys.
Zinc and aluminium additions increase its strength and manganese improves its corrosion resistance.
Zirconium refines the grain and thereby increases its toughness and fatigue resistance. However, in
spite of its excellent strength to weight ratio magnesium alloys are not used extensively on airframes,
mainly due to inferior corrosion resistance, ductility, and greater fire hazard than the aluminium alloys.
Non structural, internal items such as seats, cupboards, selected panels etc. may be made from
magnesium alloy sheet. Magnesium alloy is easy to cast and castings are used for some engine
casings particularly where corrosion is unlikely e.g. oil washed parts. Magnesium alloy heat treatments
are similar to those for aluminium alloys, i.e. annealing, solution and ageing precipitation. If overheated,
magnesium will ignite and burn rapidly. Thick sections will not readily ignite due to magnesium's high
thermal conductivity, but magnesium dust and fine chips or swarf will ignite easily. Powder or inert gas
extinguishers should be used - not liquid or foam which may not extinguish the fire and may cause
minor explosions.
4.15
TITANIUM
This is a much newer material than the more common aluminium and magnesium groups. Although
heavier than either of these two materials, titanium is only approximately half the weight of steel. When
alloyed it is capable of much greater strength and temperature resistance than the aluminium alloys
and is as strong as many alloy steels. The fatigue strength is also better than many steels.
Unfortunately it is considerably more expensive to produce than the conventional light alloys. Titanium
is very flexible, (approx. twice that of steel), and has a low coefficient of linear expansion (50% less
than aluminium and 25% less than steel). It is non-magnetic and also has low thermal conductivity.
One of the main attractions of titanium and its alloys is its excellent corrosion resistance, which is equal
to or better than 18/8 stainless steel below 500C.
One important aspect of titanium which can cause problems if used in the wrong application i.e. fire.
Although difficult to ignite, once started a titanium fire is difficult to contain because the melting point of
titanium is about 200C greater than steel. So, molten titanium can actually penetrate steel parts such
as fuel pipes etc. Titanium fires usually start through high speed rubbing. The low thermal conductivity
of titanium prevents the rapid dissipation of heat which progressively builds up locally until ignition
finally occurs. Some of the more common alloying elements are aluminium, vanadium, molybdenum
and chromium. One common aerospace alloy which is as strong as some high tensile steels contains
6% aluminium and 4% vanadium. Commercially pure titanium is ‘non-heat treatable’ (It can be
annealed, but its strength / hardness cannot be improved by H.T.). When suitably alloyed, titanium
based materials are heat treatable. The strengthening is immediate i.e. it is not an age hardening
material.
Titanium is used extensively in aerospace gas turbines, but its use is limited on subsonic civil airframes
to fasteners, and high temperature areas such as engine bays, heat shields, hot bulkheads, air ducts
etc. High speed military airframes use considerably more titanium (up to 20% of mass) due to the
higher temperatures encountered.
In appearances titanium is similar to 18/8 stainless steel. Two practical methods of identification apart
from weight are:
 spark test - a light touch of a grinding wheel will produce a brilliant white trace, ending in a
brilliant white burst.
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 moisten the titanium and draw a line on a piece of glass - this will leave a dark line similar to a
pencil mark.
Additional Notes
The machining characteristics of the various titanium alloys differ considerably. This note is, therefore,
confined to general machining conditions, an should be related to information supplied by the material
manufacturer and experience gained from the use of various machines and equipment with particular
materials.
Low thermal conductivity resulting in extremely high temperatures at work/tool interface; the reduction
of cutting edge temperature is one of the basic problems in machining titanium, but can be solved to
some extent by low cutting speeds, heavy feeds and copious supplies of cutting fluids.
High rates of feed are essential; intermittent feeding or dwelling produces rapid work hardening and
may result in early breakdown of the tool.
4.16
CUTTING FLUIDS
High quality soluble fluids, used in the diluted form recommended by the manufacturers, or chlorinated
or sulphured oils, should be used in generous quantities for all machining operations. Titanium
materials are generally not susceptible to normal corrosion attack, but it has been established that
stress corrosion cracking can take place in some welded structures which are exposed to
trichloroethylene and other chlorinated hydro-carbons, the alloys most affected in practice being the
titanium-aluminium-tin family. When it is necessary to machine a welded titanium structure, or doubt
exists regarding the use of cutting fluids with a particular titanium alloy, the material manufacturer
should be consulted. Chlorinated solvents should be removed, after machining, by use of a solvent
such as methyl ketone.
4.17
DRILLING TITANIUM
Rigidity is essential when drilling titanium and titanium alloys. A high speed drill having a point angle of
105º to 120º, with a helix angle of 38º and a thickened web is recommended. It is important that a stub
(i.e. short) drill should be used. For holes of more than 6 mm (¼ inch) diameter, a 90º or ‘doubleangled’ point is better.
Drills must be precision ground and special care must be taken to ensure that the drill tip is completely
central, as any off-set of the tip will cause work hardening as a result of friction of the non-cutting edge.
4.17.1 SPEED
For satisfactory drill life, drill surface speeds should be within 3 to 13 metres (10 to 40 feet) per minute
are used, work hardening is likely to result.
4.17.2 FEED
A continuous feed of 0.05 to 0.1mm (0.002 to 0.005 inch) per revolution for holes below 6 mm.(0.25
inch) diameter, and of 0.1 to 0.2 mm (0.005 to 0.010 inch) per revolution for larger holes is
recommended. Positive power feed must be employed whenever possible.
4.17.3 CUTTING FLUID
Flood lubrication with a heavily chlorinated cutting oil of low viscosity helps to reduce frictional troubles.
4.17.4 CENTRE DRILLING
Centre drilling should always be used instead of centre punching, as the local work hardening caused
by centre punching will cause difficulty in starting the drill and will also tend to make the drill wander as
well as blunt the drill point.
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4.17.5 SAFETY PRECAUTIONS
Fine titanium swarf or powder, even when moist, is a possible fire risk, but is considerably less than
that involved in the machining of magnesium.
Dust particles, arising from polishing and grinding etc, are highly inflammable and must be disposed of
safely. Such dust may be kept totally immersed in water until it can be burnt under controlled
conditions.
It is essential that piles of fine titanium swarf, or dust, are not allowed to accumulate around machines
where they could be subsequently ignited.
When grinding, oils with a low flash point must be avoided.
Although the bulk material is considered safe, swarf produced under certain conditions, e.g. turning at
high speed with low feed, can cause fires. A fire can be dealt with by covering it gently with a mixture
of dry asbestos wool and chalk powder. No attempt should be made to put out a titanium fire with
water or with extinguisher not specified for titanium fires.
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5.
HEAT TREATMENT & APPLICATIONS OF NON FERROUS METALS
5.1
SOLUTION HEAT TREATMENT
Solution heat treatment is sometimes also called re-crystallisation H.T. This operation is to distribute
the copper uniformly throughout the aluminium (i.e. to create a solid solution). However, although the
aluminium can accommodate 5% or so of copper in solid solution at high temperature, this condition is
unstable at lower temperatures. So, after the alloy has cooled to room temperature, most of the copper
slowly comes out of solution and separates into local `islands' of copper aluminide. The gradual
formation of these islands causes an increase in hardness and strength and these properties reach
maximum values after approximately 4 days. This gradual hardening is termed ‘age-hardening’.
Although copper is the major alloying element, other elements are also present including magnesium
and manganese in the case of dural.
Typical Solution Heat Treatment conditions for dural are: Heat to approximately 500ºC (440º - 525ºC
depending on the precise alloy), soak (leave for a period of time) and quench in water. The soak time
will depend on the mass, but will generally be from 10 minutes to 1 hour, although heavy forging may
require several hours. The quench water will usually be cold, but sometimes hot water will be specified
especially for big forgings. Some working after solution H.T. is possible before it becomes too hard, but
should normally be completed within 2 hours. Some aluminium-copper rivets however, must be driven
within 1 hour and some within 10-20 minutes of the solution H.T. Note. This 'usable' time can be
extended by refrigeration (24 hours at 2ºC and upto 7 days at -5ºC).
5.2
PRECIPITATION HEAT TREATMENT
This is carried out on many aluminium-copper alloys after the solution H.T. and is primarily employed to
speed up the age hardening process. (i.e. will fully harden within hours instead of several days).
However, some of the aluminium-copper alloys require a precipitation H.T. to develop their full /
consistent properties. Precipitation H.T. conditions range from 8 to 24 hours at temperatures 120º 210ºC followed by air cooling. Excessive soak times result in ‘over-ageing' which will cause a reduction
in properties.
5.3
ANNEALING HEAT TREATMENT
This is to soften the material to enable it to be worked without cracking. Even in this condition, ageing
will gradually occur and 24 hours is the normal limit for working after annealing, although this can be
extended if the material is stored under refrigerated conditions to slow the ageing process. (-5ºC will
give approximately 2 days delay and -20ºC will give approximately 1 week). Typical annealing
procedure is to heat in oven at 400ºC for 20 minutes to 1 hour (depending on mass) followed by air
cooling. Dural type material must never be fitted to an aircraft in the annealed state since material
properties and corrosion resistance will be severely affected.
5.4
HEAT TREATMENT NOTES
The procedures for heat treating aluminium alloys are critical if correct properties are to be obtained.
 Uniform heating is absolutely essential and two methods are used, a muffle furnace or a salt
bath. The muffle furnace uses hot air which circulated around an inner chamber in which the
aluminium alloy is placed.
 The salt bath employs liquid mineral salts so the aluminium alloy can be submerged within the
heated fluid. The salts are solid at room temperature but become liquid when heated (usually
nitrate of soda or similar) and are heated electrically. Gradual heating of the bath is necessary
to avoid spitting/ spluttering. Another precaution when using a salt bath is to avoid any adjacent
flames or sparks as the salts are inflammable.
 Accurate thermostatic control is vital, as narrow tolerances on temperatures are specified
(Typically plus or minus 5ºC).
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 Quench tanks must be sited nearby the furnace or salt bath to avoid delay between removing
from the heating source and then quenching. Most quench tanks contain cold water but hot is
sometimes specified especially for heavy sections e.g. large forgings. Limits are also stipulated
for the permissible period between heating and quenching which is known as the lag-time
typically 10 seconds max. If these lag-times are exceeded, material properties or corrosion
resistance may be adversely affected. If the cooling rate during quenching is too slow this may
also affect the corrosion resistance.
 Thorough washing of the material is essential after salt bath heat treatment to remove any salt
residue.
 Whilst there is no limit to the number of times that heat treatment may be carried out on
aluminium copper alloys, if the material is clad with pure aluminium for corrosion resistance
(ALCLAD), a maximum of three is imposed. This is to limit the migration of copper from the
dural parent material into the pure aluminium cladding, which would significantly reduce its
corrosion resistance.
5.5
HEAT TREATMENT OF ALUMINIUM ALLOY RIVETS
The most common rivet materials are 1100, 2017, 2117, 2024 and 5056
 1100 rivets are ‘pure’ aluminium, soft, ductile and are used for riveting aluminium alloy sheets
where low rivet strength is acceptable. These would not be used for structural items and can be
driven at any time without any heat treatment.
 5056 rivets are also used without H.T. and are for riveting magnesium alloy sheets.
 2117 rivets are of moderately high strength and are suitable for riveting aluminium alloy sheets
for structural items. These rivets are heat treated by the rivet maker and require no further heat
treatment before driving. This is the most widely used rivet material. Such rivets are sometimes
known as ‘Field rivets’ because of their suitability for field use without the need to resort to heat
treatment before driving.
 2017 and 2024 rivets are high strength rivets for structural use. Because of their age hardening
properties these MUST be solution heat treated shortly before driving. 2017 rivets become too
hard for driving about 1 hour after Heat Treatment and 2024 rivets need to be driven with 10-20
minutes of Heat Treatment. If these rivets are stored in a refrigerator below 0ºC, they will remain
soft enough to be usable for several days. If the appropriate times are exceeded the rivets must
again be solution treated before use or continued storage in a refrigerator. Rivets requiring heat
treatment are heated either in tubular containers in a salt bath or in small wire baskets in a
muffle furnace.
Important note.
Rivets or other fittings should not be installed in the annealed condition. Solution heat treatment is vital
on such materials, if full strength and good corrosion resistance is to be achieved. Natural ageing or
precipitation hardening should follow, whichever is specified for the particular application.
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6.
METALLIC FORMING METHODS
There are four basic methods of converting raw material into the required manufactured shape whilst
also achieving the desired material structure. They are: casting, machining, deformation and various
forms of fabrication. i.e. the joining together of smaller pieces or particles of material to form a larger
object. Welding, adhesive bonding, mechanical fasteners or even powder metallurgy come under this
latter heading.
Casting exploits the fluidity of a liquid as it takes shape and solidifies in a mould. Machining processes
provide excellent precision but the process generates a large amount of waste material. Deformation
exploits the remarkable property of materials (mostly metals) to flow plastically in the solid state without
deterioration of their properties. Such processes result in a minimum of material waste. Fabrication
techniques enable complex shapes to be constructed from simpler particles or units.
6.1
CASTING
This involves the pouring of molten material into a shaped mould and allowing it to solidify to that
shape. It is an ancient process which enables complex shapes to be produced in a wide range of
materials in a single step. Cast components can range in size from the small teeth of a zip to large
casings of several metres in diameter. Ocean going ships propellers up to 10 metres in diameter are
produced this way. Modern casting techniques have resulted in:
 high quality i.e. minimum porosity and reasonably defect-free
 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 far the most widely used
are sand and metal.
6.1.1 SAND CASTING
The two basic types of sand casting are:
 Removable / reusable pattern (usually wood or metal)
 Disposable pattern (e.g. polystyrene patterns which vaporise when the metal is poured).
6.1.2 THE SAND
Although sand casting is simple in principle, there are many vital aspects of technique which are
necessary to produce good castings. For example, the sand must have:
 Adequate binding qualities (To achieve this a small percentage of clay is added).
 Suitable porosity characteristics (To permit the escape of gas/steam formed in the mould. There
are different requirements for different metals e.g. steel and aluminium).
 Correct grain size and sufficient strength (the sand is graded by means of a sieve and the
strength is controlled by the amount of bonding agent present).
 Suitable temperature resistance (i.e. the sand must withstand the molten metal temperature
without itself fusing/melting).
 Adequate hardness (the hardness may be checked by the resistance to indentation by a spring
loaded ball).
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 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).
6.1.3 THE MOULD
The mould design must also meet certain standards, some of which are:
 The top and bottom halves of the mould ‘cope’ and ‘drag respectively, must incorporate positive
alignment features.
 The pattern must be shaped such that withdrawal from the sand leaves a perfect impression.
Tapered faces are therefore better than perpendicular faces.
 Suitable feed channels must be provided for the molten metal to enter the mould. (These are
called the ‘sprue’ and the ‘runners’).
 Strategically placed reservoirs must be incorporated to ensure proper filling of the mould as the
metal shrinks and begins to solidify. (These are called `risers'. Typical steel shrinkage is 3%-4%
and aluminium shrinkage is 6%-7%).
 The incorporation of vents where necessary to permit the escape of gas and steam when the
molten metal contacts the sand.
 Local ‘chills’ are sometimes included in the mould to encourage more rapid local solidification of
the metal.
6.1.4 ADVANTAGES / DISADVANTAGES OF SAND CASTING
Sand casting is a simple process which does not require elaborate equipment and is economical for
small batches. It is also suitable for most metals. The major shortcomings are: the process is not very
rapid, it is not particularly accurate (due to lack of sand rigidity) and it is not suitable for thin-wall
sections.
6.1.5 TYPICAL CASTING DEFECTS
Casting defects vary to some extent depending on the casting process used, but the most common
ones are:
 Inclusions (e.g. sand or mould lining material sticking to the surface)
 Porosity (usually caused by gas/vapour which is unable to escape before solidification)
 Cold Shuts (when local areas of metal are not molecularly joined due to solidification occurring
too rapidly).
 Hot Tears (where the material is cracked by excessive tensile stresses resulting from thermal
contraction).
6.1.6 SHELL MOULDING
This is a process in which a thin shell is produced by bringing a mixture of sand and a thermosetting
resin into contact with a heated pattern. When a sufficiently thick shell has been produced, the shell is
finally cured, backed up by sand or steel shot in a moulding box. The subsequent casting process is
then the same as for normal sand casting.
The advantages of 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.
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6.1.7 CENTRIFUGAL CASTING
This technique involves the molten metal being poured into a rotating mould. The process is used for
the manufacture of hollow cylinders e.g. cylinder liners, bronze or white metal bearings etc. The
rotation can result in acceleration forces of up to 60g and this produces high quality dense castings
since all of the slag migrates to the bore (due to lower density than the metal) and can then be
machined out.
6.1.8 DIE CASTING
This process uses a permanent metal mould, which results in more accurate and better finished
castings than those produced in sand. Die casting can be sub divided into gravity or pressure
processes, depending on how the metal is fed into the mould.
 Gravity Die Casting - sometimes known as ‘Permanent Mould Casting’. This casting process is
virtually identical to sand casting except that the mould (die) is metal. A wide range of metals
can be cast and hollow castings are possible if a sand core is used. Fine grain structures are
produced due to the more rapid rate of cooling than that achieved in sand casting.
 Pressure Die Casting - as implied, molten metal is fed under high pressure (thousands of psi)
and held during solidification. Most die castings are in non ferrous materials (aluminium,
magnesium, zinc, copper and their alloys), because steels have too high a melting temperature
for the metal dies to accommodate. The dies are usually made from hard tool-steels and are
water cooled. Excellent detail, super finish, low porosity and thin sections can be achieved by
this process. Expensive equipment is necessary, but very high production rates are possible.
Automatic ejection occurs and on small components 100 units per minute is not uncommon.
Hollow castings cannot be made by die casting.
6.1.9 INVESTMENT CASTING
This is a very old method of casting which was used by the ancient Chinese, but it only became of
great industrial importance in the 1950's when gas turbine manufacturing began to take off. The
process was ideally suited to the production of complex shape turbine blades and nozzle guide vanes,
which often contained tortuous inner passages, very thin sections and had to be cast in exotic
materials. The basic process is as follows:
 A master die is made first from an easily worked metal such as brass.
 Hot wax is then injected into the die under pressure to produce a wax pattern.
 The wax pattern is then removed from the die and coated with a layer of investment material (a
ceramic slurry or paste) usually by dipping a number of times.
 When the investment coating is set it is then heated to allow the wax to run out, and molten metal
is then poured into the investment mould.
 When cool, the investment coating is then broken away from the cast metallic component.
For obvious reasons this investment casting process is often referred to as the ‘Lost Wax’ process. It is
a technique which is capable of producing precision castings with dimensional accuracy of less than
0.1 mm. Surface finish is also excellent, but the major advantage that the process offers is the ability to
produce accurate, complex shapes which would be impossible by machining.
6.2
FORGING
This is a squeezing / hammering technique which is intended to achieve large deformation / shaping of
the material. The process is usually carried out hot, i.e. above the re-crystallisation temperature, in
which case these large deformations can be attained without being accompanied by any massive
residual stresses. Sometimes a cold forging operation may be necessary but in this case the material
will be harder, stronger and pre-stressed i.e. still containing unrelieved internal stresses.
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Forging ranges from the simplest form of the hand operations conducted by the blacksmith, to the
massive mechanical, powered rams used for very large forgings. The forging hammer will often have a
relatively low strike rate, but sometimes high speed pneumatic hammers are used for High Energy
Rate Forming.
Forging not only shapes the metal but also reduces grain size and produces a directional control of
grain flow. Both of these are desirable features for many engineering applications, particularly for
highly stressed components, such as crankshafts and especially if they are subject to a mechanical
fatigue environment.
6.2.1 DROP STAMPING
Drop Stamping (drop forging) involves the use of shaped dies and a heavy drop-hammer which usually
falls under gravity. The piece of material to be forged is placed between the top and bottom dies and
the drop hammer is allowed to fall the necessary number of times for the contact faces of the dies to
come together. Flash gutters are provided to accommodate excess metal which squeezes between
the top and bottom dies. Connecting rods are typical components made by the drop stamping or drop
forging process.
6.2.2 HOT PRESSING
This is similar in principle to drop-stamping but is actuated by one long, steady, squeezing operation as
opposed to a number of blows. This process tends to affect the whole structure of the component,
whereas some forging processes using multi, but light blows will mainly affect the material closest to
the surface.
6.2.3 UPSETTING
This process is sometimes called ‘Heading’ and usually involves locally heating of the end or ends of
the material immediately prior to forging. Poppet valves are formed in this way as well as forged bolts.
Sometimes this process is done cold in which case it is referred to as ‘Cold Heading’ and some rivet
heads are formed in this way.
6.3
ROLLING
This can be carried out hot or cold. When done hot it is capable of achieving major re-forming /
re-shaping and slabs can be reduced to plate or sheet and bars of circular or rectangular cross section
can be produced. Hot rolling can also produce structural shapes such as ‘H’ or ‘I’ section beams. If the
rolling is done cold, it is aimed at improved surface quality, better accuracy, and increased hardness /
strength. (Hot dilute sulphuric acid is used to remove the hot scale from steel prior to cold rolling). The
rolling process would also be used to produce sheet Aluminium alloys such as Dural and Alclad.
6.4
DRAWING
This is a purely tensile operation usually carried out hot. Wire, rod and tubing can be produced by this
process where the material is pulled through a shaped, hardened die. A ductile material is essential.
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6.5
DEEP DRAWING / PRESSING
This process uses a ram to deform a piece of sheet metal into a recessed die and is usually done hot.
6.6
PRESSING
This process uses male and female formers for shaping sheet material. The sheet is placed between
the formers which are then forced together by a powered ram. It is usually done hot except for the soft,
ductile materials.
6.7
STRETCH FORMING
This is a technique used for shaping sheet metal over a stretch-block or former. The sheet metal is
firmly gripped by clamps and the sheet is then stretched over the former by moving the clamps or the
former. The material is stretched beyond its elastic limit so that permanent deformation occurs. This
process is convenient for small batches (like the aircraft industry) and is particularly attractive since
only one former is needed. Note that local changes form concave/convex or vice versa cannot be
produced by this process.
6.8
RUBBER PAD FORMING
In principle this process uses a flexible, rubber pad attached to a hydraulic ram, which forces a piece of
sheet metal to conform to the shape of a forming block. Like stretch forming the process only uses one
former so it eliminates critical matching and alignment problems of conventional pressing and when
used for small batches (e.g. aircraft), low cost, easy to machine materials can be used for the forming
block. Rubber Bag (Hydroforming) forming uses the same principle, but incorporates a flexible
diaphragm and hydraulic pressure in place of the rubber pad.
6.9
EXTRUSION
This process forces hot metal through a shaped die to produce circular, rectangular, tubular, angle, half
round sections etc. The process is similar in some respects to drawing but extrusion forces metal from
a heated billet through hardened dies by compression whereas in drawing it is achieved by tension.
Malleability is therefore an essential property for the extrusion process.
Extrusion is normally restricted to aluminium alloys and copper alloys where extrusion temperatures of
400º/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 is usually a cold forming operation which is suitable to very soft and malleable materials e.g.
aluminium. The shaped component is formed by forcing a punch onto a ‘blank’ of material within a
shallow recess. The extruded shape results from the metal being forced to escape through the small
gap between the punch and the recess.
6.10
SINTERING
This process involves metal in powder form which is then heated to around 70% to 80% of its melting
temperature and is then squeezed to shape in a die. The process is often used to form components
made from materials with a very high melting temperature, such as tungsten. It also allows nonmetallic materials such as graphite and carbon to be incorporated into the mixture. The operation is
usually conducted in a controlled atmosphere (typically argon or nitrogen) to prevent oxidation. Under
the high pressures used, a metallurgical bond occurs (diffusion bonding) between the particles of
powder. The sintered end-product is typically around 10% to 20% porous and can then be
impregnated with graphite, high melting point grease etc. to provide excellent self lubricating properties
for plain bearings, bushes etc.
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Sintering can be used where the combined properties of materials are required e.g. copper/graphite for
electrical brushes i.e. copper to carry the current and graphite to act as a lubricant.
Tungsten carbide cutting tools can also be produced this way by incorporating tungsten carbide
particles within a cobalt matrix. Hot Isostatic Pressing uses a similar technique to sintering but uses
higher temperature and very much higher pressures to produce zero porosity. The technique is
sometimes used to heal micro-porosity in super-critical castings.)
6.11
SPINNING
This is an old process in which a piece of sheet metal can be formed to shape around a rotating former
which is mounted on a spindle of a lathe. The necessary force to deform the sheet metal is generated
by a long tool which is levered about a suitably positioned fulcrum. For thin gauge, soft metals the tool
can be manipulated by hand, for thicker gauge materials a hydraulic actuation is used on a purpose
built machine. Cones, flares, bowls and bell-mouth shapes are produced by spinning.
6.12
CHEMICAL ‘MILLING’
Sometimes referred to as chemical etching. It is a purely chemical process, not electro chemical.
Although simple in principle, chemical milling offers a method of producing complex patterns and
lightweight parts and incorporating integral ribs and stiffeners in sheet metal. Tapered sections can
also be easily formed. The unwanted material is literally eaten away by a suitable chemical. The
process is ideally suited to aluminium alloys. The chemical in this case is a hot alkaline solution, which
is usually caustic soda. It is a relatively slow process (typically a few ‘thou’ per minute) but its unique
advantages make it very attractive for airframe components. The areas which must not be eaten away
by the fluid are simply protected by a thin layer of plastic which can be brushed or sprayed on.
Although the chemically etched surface is not very rough, a drop in fatigue strength does result and in
critical applications restoration of fatigue strength is desirable and this is achieved by a light peening
operation using glass beads or steel shot.
6.13
ELECTRO CHEMICAL MACHINING
This process uses the principle of electro plating, but in reverse, i.e. instead of depositing a metal
coating on the workpiece by an electro-chemical process, metal is progressively removed from the
workpiece by reversing the polarity of the workpiece. In electro plating the workpiece is the cathode,
whereas in electro-chemical machining it becomes the anode and is thus, in effect, ‘de-plated’. The
process is ideal for metals which are difficult to machine and the finish achieved is good. High electric
current is required, (about 10,000 amps for 1 cubic inch per minute). Essential requirements for the
process are that the tool needs to be a good conductor (copper or brass) and must resist corrosion, the
electrolyte is often a salt solution.
6.14
ELECTRO-DISCHARGE MACHINING E.D.M.
This process is sometimes called spark machining or spark erosion because the technique involves the
removal of metal by an electrical spark which causes local particle melting between a suitable shape
electrode and the workpiece. The tool has no contact with the workpiece but it has to be a good
conductor so it is often made from brass, or copper. A suitable fluid usually kerosene, is fed under
pressure between the electrode and the workpiece to maintain a uniform electrical resistance, and also
to wash away the particles of eroded metal. The spark rate is around 10,000 per second and as the
greatest erosion occurs on the positive electrode, the workpiece is always made electrically positive.
The gap between the tool and the workpiece is critical and must be maintained throughout the
operation at around 0.001-0.003 inches (0.025-0.075 mm)
Any metal may be machined by this process but it is much slower than conventional machining. The
real advantage of EDM is that it is suitable on materials which are difficult to machine conventionally
and where it really excels is in its ability to produce high aspect ratio, very small holes of any cross
sectional shape. e.g. 0.010" dia. by 3 inches deep in very hard metals (0.025 mm x 7.5 cm).
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A novel variation of EDM is a technique sometimes referred to as ‘wire cutting’. This process uses a
moving fine piece of copper or nickel wire (0.002” - 0.010” dia) as the electrode. This fine wire is
positioned by, and fed over two pulleys and resembles a simple band-saw operation. The workpiece is
mounted on a table which can be moved in two axes, and if the table can be computer controlled, the
wire cutting process can cut accurate, complex shapes in metals (e.g. dovetails, firtrees etc.) which are
difficult to machine with conventional tools.
6.15
CONVENTIONAL MACHINING
The seven basic types are:
 Drilling / reaming,
 Turning,
 Milling,
 Sawing,
 Shaping / planing / slotting,
 Broaching
 Abrasive machining i.e. grinding.
These techniques have been well established for many years and most of the advances until relatively
recently have been confined to tooling improvements which have permitted higher material removal
rates. The early high carbon steel tools have been superseded by high speed steels (tungsten / cobalt
alloy steels), cemented carbides and ceramics.
So called Machining Centres have also been developed which are capable of automatic tool changes
and of doing difficult types of machining without transferring work to a different machine and re-setting
up. So a much more versatile machine tool has evolved. However, the biggest single machining
advance in modern times especially with regard to aircraft manufacture has been the introduction of
Numerically Controlled (NC) machines. NC milling in particular has revolutionised airframe
manufacture.
NC machines are machines in which motion is controlled by numbers either via punched tape or
magnetic tape. The tape instructions are based on the Binary System (or a variant) which is common
to most electronic computing devices. The primary advantage of NC machining is the ability to
accurately control the spindle, the tool or the workpiece movements in three directions (x, y and z axes)
independently or simultaneously. NC machines are capable of producing compound shapes and
contours and are specially suited to generation of integral spars, ribs, stiffeners in slabs or forgings.
NC machines usually incorporate a feed-back system which tells the control unit how much actual
movement is made, analysis is then done and final compensation eliminates any error, i.e. the motion
ceases when the input and feed-back signals agree. Electrical control of the machine servo motors
can control movements as small as 0.00002".
CNC machines (i.e. Computer Numerically Control) differ from NC machines only in that the electronic
control unit on the CNC machine is more sophisticated in such that it is adaptable to a wide variety of
software and can accommodate a diverse range of programs. Although the capital cost of NC / CNC
machines is high, the following advantages make such machines technically desirable and
economically viable where super light, complex, high-tech manufacture is concerned:
 Complex shapes with integral features are possible
 The number of jigs and fixtures is reduced.
 A-reduction in manufacturing time.
 Adaptable to short runs.
 Greater accuracy and consistency.
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 Program can be changed to accommodate modifications.
6.16
SUPERPLASTIC FORMING
Some Titanium alloys when heated become extremely ductile and can plastically deformed without
necking occurring. This superplasticity can be exploited in the forming process, an inert gas is used to
blow the material into the required shape.
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7.
COMPOSITES & NON METALLIC MATERIALS
Fibre reinforced composites are used in ac construction and consist of strong fibres such as glass or
carbon, set in a base (matrix) of plastic or epoxy resin, which mechanically and chemically protects the
fibres. These materials have had a dramatic effect on aircraft and engine construction in recent years
and are used on an ever increasing scale. Among the many different types of composite materials
available, the ones most commonly used in aerospace are:
 Glass reinforced plastic (GRP)
 Carbon fibre reinforced plastic (CFRP)
 Aramid (Kevlar)
7.1
ARAMIDES
This is one of the newer materials and is a trade name for a high strength to weight ratio fibre
developed by the Du-Pont Company. Kevlar is the brand name of one of a group of materials.
Another similar material is brand named ‘Nomex’. Kevlar is formed from a aromatic polyamide
(commonly known as a ‘Para Aramid’, Nomex is a ‘Meta-Aramid’), by extruding the polymer through a
die with very small holes. The resulting fibres are collected and joined into yarns which are
subsequently woven into suitable structural fabrics. It can be used in this form for ropes, cables, bullet
proof vests, torpedo netting etc. or made into a rigid composite material like GRP or CFRP by using
suitable adhesives. The stiffness of Kevlar lies roughly mid-way between GRP and CFRP, but its
tensile strength is comparable to carbon fibre. As it is about 17% lighter than CFRP, Kevlar thus has
the highest strength to weight ratio. This latter advantage, plus the fact that the cost of Kevlar is
partway between GRP and CFRP, means that it is being used on an ever increasing scale in the
aerospace business. Three peculiarities of Kevlar exist:
 Firstly, it has a significantly lower compressive strength than tensile strength. This is not the case
with GRP or CFRP.
 Secondly, it will slowly deteriorate if exposed to ultra violet (UV) light for prolonged periods. It is
therefore, necessary to use a pigmented paint or some other suitable barrier.
 Thirdly, it is difficult to cut a kevlar composite cleanly with conventional tools, high pressure water
jets are used.
7.2
GLASS REINFORCED PLASTIC (GRP)
Often referred to as Glass fibre or fibreglass, this material is comprised of glass fibres bonded together
by a suitable resin.
The ultimate tensile strength of undamaged very small diameter glass fibres (approximately 0.0lmm) is
extremely high, (greater than 2000 N/mm) although this figure is reduced significantly if the fibres are
slightly damaged. When moulded with resin, the resulting composite is of considerably lower strength.
Nevertheless, good GRP structures are stronger than mild steel and on a simple strength for weight
basis, can be comparable to high tensile steel if the fibre form and lay-up is near optimum. It is
however, considerably less stiff than steel or even aluminium. A graphic example of GRP flexibility is
the enormous deflection which takes place in the pole during a pole vault. As the glass fibres are
about a hundred times stronger than the resin, it is obviously necessary to get as much fibre packed
into the moulding as possible. Non structural items may be made from, or include a percentage of
chopped strand mat, ( i.e. glass fibres in a random, non woven state) but where considerable strength
is required, uni-directional glass cloth is used. To provide all round strength, sheets of uni-directional
cloth can be layed up at 90º to each other e.g. like the grain in plywood.
Sometimes such sheets are used as facings for an internal honeycomb of plastic impregnated paper,
to give a very efficient structure in terms of strength, stiffness and weight.
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The glass fibre sheet material can be supplied with cloth already impregnated with resin and partially
cured (‘Pre – Preg’) in which case it is necessary to keep the material in refrigerated storage. Resin
curing is usually done at temperature (range 120 - 170ºC) with the GRP component in its mould and
often under pressure, in an autoclave.
The main reasons for using GRP are:
 Where metal must not be used e.g. radar domes or other non-electrical conduction applications.
 Ease / cost of producing very complex shapes.
 Good strength / weight ratio.
 The ability to produce selected directional strength.
It's main disadvantage is that it lacks stiffness and as such is not suitable for applications subject to
high structural loading.
7.3
CARBON FIBRES, REINFORCED PLASTIC (CFRP)
Carbon Fibre, Reinforced Plastic (CFRP), is a composite material which was primarily developed to
retain (or improve on) the high strength to weight ratio characteristics exhibited by GRP but with very
much greater stiffness values (Young's Modulus `E').
Carbon fibres are very stiff and when formed into a composite the Young's Modulus value can be
higher than steel. CFRP is not only six times stiffer than GRP 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 in diameter and are produced by subjecting a fine thread of a suitable nylon
type plastic to a very high temperature to decompose the polymer, driving off all of the elements except
carbon. It is then stretched at white heat 2000-3000ºC to develop strength. Unfortunately, the process
is complex and very costly.
Nevertheless, where the high cost can be justified, CFRP can offer considerable weight savings over
conventional materials. CFRP components are generally made from ‘Pre-preg’ sheet (fibres
impregnated with resin and hardener which only require heat and pressure to cure). Some specialist
items are made by a laborious but ideal process called filament winding in which a carbon fibre string is
wound over a former in the shape of the workpiece whilst bonded with resin.
Because of CFRP's high stiffness modulus it is also used extensively to stiffen GRP or aluminium alloy
structures.
Aramid, a material known as Carbon-Carbon, where the resin is also graphised, is now being used for
rotors and stators on brake units. It offers a significant weight saving, as well as high efficiency as it
dissipates the heat generated very quickly. At present the life is limited due to cracking and wear.
7.3.1 RECENT DEVELOPMENTS
Fibre reinforced composites are being increasingly used in the fabrication of primary airframe
structures. It is, however unlikely that many all-composite aircraft will be built; a maximum composite
content of 40% is considered to be a reasonable estimate. Replacing 40% of an aluminium alloy
structure by CFRP would result in a 40% saving in total structural weight. CFRP is used on the wing,
tailplane and forward fuselage of the latest Harrier and for development work on the Jaguar wing and
engine bay doors. The use of composites in the manufacture of helicopter rotor blades has led to
significant increases in their life, in some cases, they have an unlimited life span (subject to damage).
The modern blade is highly complex and comprises CFRP, GRP, stainless steel, a honeycomb core
and foam filling.
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7.3.2 GENERAL INFORMATION
A sheet of fibre reinforced material is anistropic, that is, its properties depend on the direction of the
fibres. Random direction fibres would result in a much lower strength than uni-directional fibres parallel
to the applied load. However, the strength (and stiffness) of a uni-directional lay-up would be very low
with the applied load at 90º to the fibres, as this is primarily a test of the resin. Hence the usual practice
of placing alternate layers at 90º to each other.
Due to small variations in the size of the individual fibres and the final quality of the finished
component, which can be affected by careless handling, variations in cleanliness or lay-up, voids,
pressures, temperatures etc. there will inevitably be a greater scatter on final strength than on a
conventional metallic component. Due allowance on stress reserve factors is therefore essential.
Composites usually have good internal damping characteristics and are therefore less prone to
vibration resonances.
All three composites have very low elongation properties and therefore toughness. Aluminium alloy
has a typical elongation to fracture value of 11% whereas the composites range from 0.5% for CFRP to
3% got GRP.
The maximum operating temperature for GRP, CFRP and Kevlar composite depends to some extent
on the actual adhesive used but is generally in the range 220-250ºC. (Some composites such as
carbon fibre in a carbon matrix have very high permissible operating temperatures (around 3000ºC)
and are used for high energy braking applications and thermal barriers for space vehicles).
Composite materials consist of two or more different materials which are mechanically or
metallurgically bonded together, and each component material retains its identity, characteristic
structure and properties. The resulting composite material possesses physical properties (especially
stiffness and strength) which are unattainable with the individual constituents.
7.4
PLASTIC SEALANTS & ADHESIVES
7.4.1 PLASTICS
Plastics are based on carbon and hydrogen. The major raw material source is crude oil or coal. One
well known exception however, is cellulose, which comes from wood or from cotton plants. Although
the basic chemical elements of plastics are carbon and hydrogen, other elements which are present in
some plastics are oxygen, nitrogen, chlorine and sulphur. Plastics are made up of characteristic long
chain molecules i.e. they have a very high length/thickness ratio.
7.4.2 MAJOR PLASTIC GROUPS
There are three major groups of plastics, namely: Thermoplastics, Thermosetting plastics and
Elastomers.
 Thermoplastics have the following properties:
i.
they are solid at room temperature
ii.
they are soft (mouldable) on heating
iii.
they become hard again when cooled
iv. the characteristic softening on heating and hardening on cooling is repeatable.
 Thermosetting plastics have the following properties:
i.
they are soft or even liquid in their natural state
ii.
they become rigid when cured
iii.
they cannot be re-softened by heating once cured
iv. they are relatively hard and brittle.
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Note: Thermosetting plastics are generally stronger, have a lower ductility and lower impact
properties than the Thermoplastics
 Plastic Elastomers have considerable elastic properties. They will tolerate repeated elongation
and return to their original size and shape, similar to natural rubber.
7.4.3 PRIMARY ADVANTAGES OF THE USE OF PLASTICS
Plastics are being used on an ever increasing scale and are frequently replacing some of the more
conventional materials such as metals, wood and natural rubbers. Different plastics have properties
which make them a popular choice over conventional aircraft materials. Some of the more important
properties / characteristics of plastics which help to explain their popularity are:
 Lightness most plastics have specific gravities of 1.1 to 1.6 whereas the lightest structural metal
(magnesium) has a value of 1.75. The more common engineering materials such as aluminium
and steel have values of 2.7 and 7.8 respectively.
 Corrosion Resistance excellent. Plastics will tolerate hostile corrosion environments and many
of them resist acid attack.
 Low Thermal Conductivity this property makes many plastics ideal for thermal insulators.
 Electrical Resistance excellent. Consequently plastics are used in enormous quantities for
electrical insulation applications.
 Formability many plastics are easily formed into the finished product by casting moulding or
extrusion, often in a single operation.
 Surface Finish excellent surface finishes can be achieved in the basic forming operation, so
finishing operations are not necessary.
 Relatively Low Cost because although some of the materials may not be particularly cheap, the
lack of machining necessary and the high production rates possible keeps the costs down.
 Light Transmission some plastics are naturally clear whilst other are opaque. Consequently a
range of light transmission properties are possible. Optical properties can also be achieved with
some plastics.
 Vibration Damping many plastics are naturally resistant to fatigue. Because of the high value of
internal damping present, resonances will tend to be of relatively low amplitude.
7.4.4 PRIMARY DISADVANTAGES OF PLASTICS
Although plastics are extremely useful materials, some shortcomings inevitably exist, particularly when
compared to some metals. Their major deficiencies are:
 Lack Of Strength most plastics are much weaker than metals e.g. mild steel has approximately
six times the strength of nylon. (However, mild steel is six times the weight of nylon so on a
strength/weight ratio, they are comparable).
 Low Stiffness plastics have a very inferior value of Young’s Modulus compared with the
common metals.
 Low Impact Strength many plastics have poor impact strength, but there are a few exceptions
such as certain polycarbonates.
 Poor Dimensional Stability mainly due to high values of thermal coefficient of expansion.
 Poor High Temperature Capability metals are generally capable of retaining reasonable
strength at much higher temperatures than the plastics. The long term maximum operating
temperature for the better plastics is not usually above 250ºC. High temperature metals can
operate for long periods well in excess of 800ºC.
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 Moisture Absorption many types absorb moisture which can result in a significant loss of
strength in a humid environment.
 Ultra Violet Light some plastics deteriorate when exposed to U.V. light for long periods.
Increased brittleness and loss of strength can occur.
7.4.5 MAIN USES FOR PLASTICS
Plastics are particularly useful for applications which involve relatively low stress levels where lightness
is important, where low electrical or thermal conductivity is important.
7.4.6 SOME OF THE MORE COMMON PLASTICS & THEIR APPLICATION
Plastics are now used on an enormous scale and types available are too numerous and complex to
deal with in depth. However, some of the more common plastics used in engineering are:
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7.4.6.1 Thermoplastics
 Acetate widely used for tool handles, and electrical goods.
 Poly-Ethylene commonly known as polythene. Uses include flexible tubing, cable insulation and
packaging.
 Poly-Propylene stronger, harder and more rigid than polythene. Uses include high pressure air
piping.
 Poly-Vinyl-Chloride commonly known as PVC. Varying degrees of rigidity / flexibility are
achievable by varying the amount of plasticiser used. Rigid moulded sections or piping can be
produced or flexible electric cable insulation
 Polystyrene can be produced in rigid form, but is more familiar when in the expanded form,
when it is useful for thermal insulation, buoyancy or shock resistant packaging.
 Acrylic these are particularly useful where light transmission is necessary. Perspex and
Plexiglas belong to this family. They have excellent light transmission properties and are also
resistant to splintering. There is a tendency for some fine craze cracking to develop if exposed
for long periods to ultra violet light. These transparent plastics may be solid or laminated. When
laminated two or more layers are bonded together with a clear adhesive and in this form they are
more shatter resistant and ideally suited to pressurised aircraft windows.
An even stronger and more shatterproof transparent plastic can be achieved by stretching the
acrylic in both directions before final shaping. These improved properties result from the
stretching operation causing a preferential alignment of the long chain molecules. Extreme care
should be taken when handling acrylics as they are they are easily scratched. The acrylics are
supplied with a paper or rubberised film which should not be removed until required for use. If
dirty, they should be cleaned with cold water or soapy water. Care should also be taken when
using solvents in the vicinity of acrylics. Some solvents or their vapours may cause crazing of the
material. Reference to the appropriate Manuals or manufacturers specification sheets are
essential.
 Poly-Carbonates these have similar uses to the acrylics (Perspex etc) but are more temperature
resistant and also have superior impact strength. They are also more expensive.
 Nylon belongs to the polyamide family and is an extremely useful and versatile material. It is
strong, tough and also has low friction properties. It can be used as a fibre or produced as a
moulding. Popular uses include textiles, furnishings, ropes, tyre reinforcement, bushes, pulleys,
gears, lightweight mouldings such as brackets, handles etc.
 Poly-Tetra-Fluoro-Ethylene commonly known as ‘PTFE’, it is similar to nylon in appearance but
is denser, whiter and much more expensive. It has a wax-like surface and this characteristic
results in very low friction properties which makes it suitable for bushes and gears. It also has a
high temperature capability (over 300ºC) and is also extensively used as a non-stick coating e.g.
Teflon. PTFE tape is often used as a thread sealant for oxygen pipe threads, and backing rings
for hydraulic seals
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7.4.7 THERMO-SETTING PLASTICS
 Bakelite one of the earliest plastics, it is hard and fairly brittle. It is often used with a suitable
filler material (mica, or wood flour) and is widely used for various electrical mouldings and low
stressed handles.
 Polyester Resin can be extruded into fine filaments and woven into fabric (like nylon) or cast into
shape or used for bonding glass fibres into a rigid composite material. It is also useful as a heat
resistant lacquer.
 Epoxy-Resin is a very strong material which is extensively used as the bonding medium for a
wide range of composite materials. It is also an extremely versatile adhesive. Epoxy resins can
be cured by means of suitable catalyst (hardener) without recourse to heating.
7.4.8 ELASTOMER PLASTICS (SYNTHETIC RUBBERS)
 Buna ‘S’ relatively cheap material with a performance similar at natural rubber. It is often used
for tyres and tubes, but its poor resistance to fuels/oils/cleaning fluids makes it unsuitable for
seals.
 Buna ‘N’ also known as Nitrile, it has excellent resistance to fuels and oils, and it is used for oil
and fuel hoses, gaskets, and seals. It also has low ‘stiction’ properties when in contact with
metal and is therefore particularly suited to moving seal applications.
 Silicone Rubber has very good high and low temperature properties (-80ºC to + 200ºC). Is often
used for seals but is also used for potting of electrical circuits because of its ability to retain its
rubbery state even at low temperatures.
 Fluoro-Elastomers these have exceptional high temperature properties and can be used at
250ºC. They are also solvent resistant and are mainly used for high temperature seals. A
common name for this material is Viton. This material is expensive.
 Neoprene has very good tensile properties and excellent elastic recovery qualities. It is also
solvent resistant and therefore has a wide range of applications as fuel and hydraulic seals and
gaskets. However, because of its special elastic recovery properties it is ideally suited to
diaphragms and hydraulic seals (DTS 585)
 Poly-Sulphide Rubber although it possesses relatively poor physical properties it has
exceptionally high resistance to fuels and oils and is widely used for lining or sealing fuel tanks.
It is also used for lightly stressed seals and hoses which come into contact with fuels or oils.
These compounds are commonly known as PRC or Thiokol.
7.4.9 PLASTIC FORMS
Plastics can be obtained in various forms, such as liquids, solids, powders, pastes, fibres (filaments),
films, and adhesives, (Note. adhesives are covered in detail later.
Apart from their conventional use, plastics also have specialised applications such as Ablative
Coatings for rockets, missiles and spacecraft. These plastics provide short term protection from
intensive, high speed, frictional heating. The thermal de-composure of the plastics into porous carbon
char and gas consumes immense amounts of heat in this chemical transformation.
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7.4.10 PLASTIC MANUFACTURING PROCESSES
The most common manufacturing methods are as follows:
 Casting. Where the molten material is simply poured into a mould and allowed to set.
 Moulding. Where powder, liquid or paste is forced into a set of shaped dies.
 Extrusion. Where plastic is forced through a suitable shape die. Rod, sheet tube, angle sections
etc. are produced this way.
 Lay-Up. Where load carrying plastic fibres and an adhesive are layered in a mould or around a
former.
 Sandwich-Construction. Where plastic facings have sandwiched between them a honeycomb
or foam core. Very stiff but light structures are achieved by this method.
 Sketches of three common methods of plastic forming techniques are shown overleaf:

Compression Moulding
Note: Vacuum Forming uses a similar tooling but in this case the plastic is sucked into contact with the
shaped dye (often used to manufacture aircraft interior trim).
7.4.11 ADHESIVES & SEALANTS
Adhesive bonding is being used on an ever increasing scale and particularly in the aerospace industry.
Adhesives are used for tasks varying from aircraft control surfaces, fuselage construction, to helicopter
rotor blades.
7.4.12 THE MECHANICS OF BONDING
The actual adhesive bond may be achieved in two ways:
 Mechanical. The adhesive penetrates into the surface and forms a mechanical lock by keying
into the surface. It also forms re-entrants where the adhesive penetrates behind parts of the
structure and becomes an integral part of the component to be joined.
 Chemical (Specific). In this method of bonding, the adhesive is spread over the surfaces to be
joined and forms a chemical bond with the surface
In practice, most adhesives use both of these methods of bonding to form a joint.
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7.4.13 STRESSES ON A BONDED JOINT
Adhesive joints are liable to experience four main types of stress:
Tensile. Where the two surfaces are
pulled directly apart.
Shear. Where the two surfaces tend to
slide across each other.
Cleavage. Where two edges are pulled
apart.
Peel. Where one surface is stripped
back from the other
7.4.14 DESIGN OF ADHESIVE JOINTS
Joint stress is at a maximum when the adhesive is in shear. Adhesives should not be used if
significant stresses will be carried in tension or peel. As the strength of the adhesive bond is
proportional to the area bonded, lap joints are generally favoured.
7.4.15 ADVANTAGES
The major reasons for the widespread use of adhesives are as follows:
 No weakening of the component due to the presence of holes. Also providing a smooth finish
due to lack of rivet heads.
 No local stress raisers which are present with widely pitched conventional fasteners, (Bolts, rivets
etc).
 Can be used to join dissimilar materials and materials of awkward shapes and different
thicknesses. (Rivetting and welding is not always possible on very thin and very thick materials).
 Although the strength per unit area may be inferior to a mechanical or welded joint, adhesive
bonding takes place over a greater continuous area and therefore gives comparable or
increased strength coupled with improved stiffness.
 Adhesive / sealants provide electrical insulation and prevent galvanic corrosion between different
materials.
 Leak-proof (fuel and gas) joints can be achieved.
 The elastic properties of some adhesives gives flexibility to the joint and may help to damp out
vibrations.
 Heat sensitive materials can be joined.
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7.4.16 DISADVANTAGES
The major disadvantages associated with adhesive bonding are:
 Limited heat resistance. This restricts the process to applications where environmental
temperatures will not generally be above 200ºC.
 Poor electrical and thermal conductivity.
 High thermal expansion.
 Limited resistance to certain chemicals (i.e. some paint strippers).
 Integrity difficult to check by non-destructive means.
7.4.17 STRENGTH OF ADHESIVES
The three most important consideration are:
 Failing Stress i.e.
Failing Load
Glued Area
 Creep behaviour
 Durability i.e. its long life capability without serious deterioration.
7.4.18 TYPES OF ADHESIVES
Although there exists an enormous range of adhesives, the two major groups are Structural and
Flexible. Structural adhesives are primarily aimed at applications where high loads must be carried
without excessive creep. They are therefore relatively rigid, but without being excessively hard or
brittle. Flexible adhesives are used when some flexing slight relative movement of the joint is essential
and where high load carrying properties are not paramount. In general, structural adhesives are based
on resins, (the most common ones being epoxy or polyester) whereas flexible adhesives are based on
flexible plastics or elastomers. In practice, structural adhesives often contain a small proportion of
elastomer and the flexible adhesives contain some resin. Another group is the two-polymer type which
has a reasonably even balance of resin and elastomer which results in a flexible yet fairly strong
adhesive. Examples of some specific adhesives are as follows:
7.4.18.1 Thermo-setting Adhesives (Fibreglass Resins)
These rely on heat to make them set. The setting process causes a chemical change to occur within
the adhesive. This change is not reversible and so once set the adhesive will not re-soften if heat is
applied.
7.4.18.2 Thermo-plastic adhesives (Copydex)
These are based on synthetic materials such as polyamides, vinyl, acrylics, cellulose etc. and also on
natural materials such as resins, shellac and rubber. They do not form as strong a bond as thermosetting adhesives, but being more flexible, they are suitable for joining non-rigid materials.
7.4.18.3 Solvent Activated and Impact Adhesives (Evo-Stick)
This type contains a solvent which softens the adhesive for easy application. A bond is formed when
the solvent evaporates. In the case of impact adhesives, the adhesive is spread over both surfaces and
left to dry by evaporation. When dry, the two surfaces are brought together and they bond by intermolecular attraction.
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7.4.18.4 Epoxy-Resin (Araldite)
These adhesives are base on the reaction product of Acetone and Phenol and can cure at room
temperature by the action of a hardener, or by the application of heat. They will bond most surfaces,
and as no gas or vapours are released during curing, require little or no pressure to form the joint. The
bond is strong in tension or shear, but the cured adhesive is very brittle and will fail in cleavage and
has poor peel strength. To improve these characteristics, modified epoxies have been produced with
thermoplastics such as nylon incorporated into the resin. This improves the performance in peel and
also improves the flexibility and the wetting of non-porous surfaces such as metal or glass.
7.4.18.5 Phenolic Adhesives (Aerodux)
These are based on an aldehyde and phenol reaction. The by-product of this reaction is to give off
water and formaldehyde so high pressures are necessary to prevent the join from being forced apart.
Curing requires a temperature of up to 480ºC. This is the type of adhesive used in the production of
Plywood’s. The addition of thermo-plastic modifiers such as synthetic rubbers have extended the use
of this type of adhesive.
7.4.18.6 Redux
Probably the most common structural adhesive. It is widely used in the manufacture of aircraft,
particularly for the attachment of stringers to fuselage skin in aircraft such as the BAe 146 and BAe
125. The obvious advantage is that rivets are not used in a pressurised area. It is only practical to use
the Redux method in the manufacturing process and repairs to this type of structure would be done by
rivetting. The components to attached are placed in an autoclave with the adhesive in sheet form
between them. Heat and pressure are then used to form the bond.
7.4.18.7 Thiokol (P.R.C.)
One extremely useful flexible adhesive/sealant because of its great resistance to oils, fuels and other
solvents is Polysulphide rubber (common trade name Thiokol). Its high flexibility coupled with its
solvent resistance makes it an ideal adhesive/sealant for fuel tanks. It is also resistant to degradation
by light, oxygen and heat and is used extensively for sealing aircraft pressure cabins and windows. It
is normally available as a two part mix and may be applied by brush, spatula or by an applicator. The
two parts should not be mixed until ready for use and the complete curing process may take up to 48
hours. Warm air heating will speed up this process. The unused adhesive has a limited life in storage,
so the date should be checked before use.
7.4.18.8 Specialist Adhesives
One of the specialist adhesives is ‘Cyano-Acrylate’ or so-called Super-glue. This cannot become
activated by itself, but is catalysed by the presence of atmospheric moisture or to the presence of
oxygen in the air the curing process can occur, almost instantaneously.
Another specialist group of adhesives commonly used in engineering are anaerobic. These are liquid
when exposed to air, but will cure when confined to small spaces in the absence of air. Its main use is
for adhesive locking of threaded fasteners (Loctite) although it is sometimes used for securing bearings
in housings or on shafts.
7.4.19 ADHESIVE FORMS
Adhesives can be obtained in a variety of forms, the most common being liquid, paste or film. Other
forms are available however, such as special foaming types which are used to splice honeycomb
sections together. Some require heat for curing, whilst others can be cured by the addition of a
catalyst or hardener.
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7.4.20 ADHESIVES IN USE
To achieve optimum bonding, performance and life in service from adhesives & sealants, it is
absolutely crucial to follow carefully planned processes and procedures and to pay the utmost attention
to quality at every stage. In fact, the major criticism. levelled against the use of adhesives are:
 Absolute cleanliness at all stages is essential. Surface preparation of the component also
crucial. The need for cleanliness in this type of process is fairly obvious; to ensure consistent
results on structural components a purpose built ‘clean room’ will reduce contamination to a
minimum.
 Pressure and heat may be required. Sophisticated equipment is required to produce pressure
over the components in areas where adhesives are applied. This will often entail vacuum bags,
purpose built ovens, or pressurised curing ovens (autoclaves).
 Inspection of the bonded joint is difficult. Special inspection techniques and test pieces are
necessary to check the integrity of the bond. Prior to preparing the mating surfaces for ‘gluing’, it
is necessary to carry out a ‘dry’ lay-up i.e. a trial assembly of all related parts to check and adjust
the fit if necessary. This procedure is essential to enable the final assembly ‘wet’ lay-up to
proceed without delay and without the risk of generating swarf or contaminating specially
prepared surfaces.
7.4.21 SURFACE PREPARATION
 Grease, oil or other contaminants must be removed by suitable solvents.
 An optimum surface roughness must be produced.
 Once pre-treated, a surface must be protected from harmful contamination until the bonding
process is complete.
 Surfaces to be bonded are normally thoroughly cleaned/degreased in a suitable solvent. This
may be followed by a chemical etch/oxide or light blasting treatment followed by a water wash
and drying.
7.4.22 FINAL ASSEMBLY
The adhesive is then applied (usually within a specified time e.g. 12 hours otherwise re-processing may
be necessary) and the assembly suitably clamped or put in a nylon vacuum bag and heated in an
autoclave. The curing process then takes place under carefully controlled temperature and pressure
conditions.
When cool the component is inspected:
 Visually for positioning and for satisfactory spew line.
 Glue - line thickness checked with calibrated electronic probe.
 Specimen test pieces checked out for shear and peel properties.
Following a satisfactory inspection the component is finally given appropriate corrosion protection usually over-painting.
Note. After commencing the final (wet) lay-up curing of the adhesive must be carried out within a
specified time (usually 12 hours). If this period is exceeded by a few hours it is necessary to increase
the temperature and pressure levels during curing and to obtain official concession cover for this
discrepancy. If the permissible time between wet lay-up and curing is greatly exceeded (e.g. a full shift
or day), it will be necessary to dismantle and not only re-commence the wet lay-up but also to possibly
repeat some of the preliminary surface preparation treatments such as etching.
7.4.23 TYPICAL (ABBREVIATED) PROCESS
 Dry lay-up (i.e. dummy run)
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 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:
i. Positioning, uniform, continuous glue-line etc.
ii. Check glue-line thickness (electronic probe).
iii. Check test piece results (shear & peel).
 Carry out final post cure surface treatments. e.g. over-painting of primer, sealant or top coat of
solvent resistant paint)
7.4.24 SAFETY
Although many of the adhesives in current use are supplied in film form, some are in liquid or paste
from which toxic /inflammable vapours are emitted prior to curing. Many of the necessary surface
preparation solvents also give off toxic / inflammable vapours. Controlled ventilation, protective
clothing, and anti-fire/explosion practices are therefore essential.
7.5
DEFECTS IN COMPOSITE COMPONENTS
7.5.1 CAUSES OF DAMAGE
Damage to composite structures may result from a number of causes such as:
 Impact caused by bird strikes and contact with obstructions on the ground.
 Erosion caused by rain, hail, dust etc.
 Fire
 Overload caused by heavy landings, flight through turbulent air and excessive ‘g’ loading.
 Lightning strikes and static discharge.
 Chafing against internal fittings such as pipes and cables.
7.5.2 TYPES OF DAMAGE
The types of damage which may affect fibreglass structures are:
 Cracks which may simply affect the outer lamination or may penetrate through the skin.
 Delamination which involves separation of the fibreglass layers and may affect single or multiple
layers.
 Blisters which usually indicate a breakdown in the bond within the outer laminations and may be
caused by:
i. Moisture penetration through a small hole
ii. Poor initial bonding
 Holes. These may range from small pits, affecting one or two outer layers, to holes which
completely penetrate the component. These may be caused by lightning strikes or static
discharge.
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7.5.3 ASSESSMENT OF DAMAGE
As with metal structures, the damage occurring to GRP or CFRP structures may be classified as
negligible (or allowable), repairable by cover patch, repairable by insertion or repairable by
replacement.
These classifications may only be determined by reference to the appropriate aircraft Structural
Repairs Manual. Signs of secondary damage (i.e. damage occurring remote from the primary damage)
must not be overlooked. This is particularly importance in the case of impact damage where the shock
may be transmitted through the structure to cause damage away from the point of impact. In some
instances secondary damage may be more serious than the primary damage.
Sometimes damage may be difficult to detect due to the natural flexibility of the material which may
cause it to spring back into shape. Any evidence of cracking, straining, crazing or scuffing of the gel
coat should be regarded with suspicion as they may indicate the presence of damage.
Where delamination is known, or suspected to exist, the area surrounding the visible damage should
be checked to determine the extent of the damage and integrity of the laminations. This can be done
by tapping the skin with a small metallic object such as the edge of a coin which will produce a live
resonant tone if the laminations are sound or a flat response if delamination has occurred.
7.6
REPAIRS TO COMPOSITES
7.6.1 GLASS FIBRE COMPOSITE REPAIRS
On this course we will be dealing exclusively with repairs to glass fibre structures including honeycomb
cored structures.
Glass fibre composites have two basic constituents, the glass fibre and the surrounding plastic matrix.
The glass fibres reinforce the plastic matrix and carry most of the load. The matrix gives the composite
its rigidity and protects the fibres from attack by moisture or chemicals.
Glass fibres are generally woven into a fabric which gives a regular orientation to the fibres and allows
them to be handled more easily.
To produce a glass fibre laminate, successive layers of the fabric are placed into position and
impregnated with resin. The liquid resin solidifies within a few hours and after post curing at elevated
temperatures, forms a strong matrix around the fibres.
Using this technique, intricate shapes can easily be formed with the load carrying filaments orientated
in the best possible manner. It is also possible to reinforce the laminate locally and to mould in load
bearing fittings etc. into the laminate.
7.6.2 TYPES OF GLASS REINFORCEMENT
After production of the basic glass fibres they are collected together to form a collection of continuous,
parallel fibres known as a roving.
Glass fibre cloth is made by weaving roving together. Depending on the closeness of the weaves and
the number of roving in each weave of the fabric, different weights of cloth may be produced.
There are two main types of glass cloth, bi-directional and uni-directional.
7.6.2.1 Bi-directional Cloth
A bi-directional cloth has the same number of roving in both wrap and weft directions and as such can
take stresses in both directions.
There are two main types of bi-directional cloth, plain weave and twill weave.
Plain weave, as shown in fig 5.1.1, is woven with an over one and under one configuration and is used
for most flat surfaces. Twill weave, fig 5.1.2, has a weave with an over one and two configuration. It
gives drapeability and is used where curved component shapes are required.
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7.6.2.2 Uni-Directional Cloth
A uni-directional glass cloth has the majority of the glass fibres lying parallel and in one direction with
only enough transverse fibres to hold the fabric together.
Roving may also be used either individually or grouped together to give a fully uni-directional
composite.
7.6.2.3 Chopped Strand Mat
Chopped strand mat has random short fibres lightly held together with a binder.
A laminate of this material is heavy and of low strength compared with one of woven fabric. As a
result, it is of little use I aircraft construction.
7.6.3 RESINS
The choice of resin for a particular application is most important because they are produced with the
necessary properties to suit certain requirements and are not suitable for universal application.
Most laminating resin comes in two liquid parts; a resin and a hardener. Once hardener is mixed with
the basic resin a chemical reaction begins and the mixture begins to solidify.
Some resins are supplied as a three part mix consisting of resin (adhesive), accelerator and catalyst. It
is vitally importance when mixing this type of resin that the accelerator is never mixed with a free
catalyst or an explosion may occur. The correct mixing procedure must be: resin and catalyst must be
mixed together before adding the accelerator.
7.6.4 MIXING
In any resin mix the proportions are absolutely critical since the cured strength depends on it. The
proportions are normally specified by weight of the quantity of resin required. An excess of hardener in
the mixed resin is as damaging as a deficit. In both cases the cured resin will have an incomplete
molecular structure and poor physical properties as a result.
Scrupulous cleanliness is essential in the mixing process which should be carried out 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.
7.6.5 POT LIFE
The temperature of the resin mix affects the rate at which the curing reaction occurs. If the
temperature is too low the resin will be too thick to work, whereas if the temperature is too high, the
resin will be comparatively thin and will drain out of the laminate before solidification occurs. Ambient
temperature and humidity requirements are specified by the resin manufacturer.
The length of time before a mix of activated resin begins to solidify is called ‘pot life’ and is dependent
on the temperature and quantity of resin. Once the resin becomes thick and stringy, the curing process
is well on its way. Resin in this state should not be used since the cured strength properties will be
seriously degraded.
To prevent waste, only sufficient resin should be mixed for the task in hand.
7.6.6 CURING
Most resins used in aircraft structures will cure at room temperature (about 20ºC) but may take several
days to reach a fully cured state. Once the resin has hardened, post curing at elevated temperature is
required for the resin to gain its full strength.
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For repair purposes the heat is usually applied by means of an infra-red lamp or electric heater. For
components which have been removed from the aircraft an oven of suitable size may be used. This
allows accurate control of temperature. If a large enough oven is not available, a hot air tent should be
constructed around the repair with a thermometer measuring the average temperature inside the tent.
Temperature may also be controlled by use of temperature indicating lacquer or pencil. These, when
applied adjacent to a repair will melt or change colour when a pre-determined temperature has been
reached.
The times and temperature required to effect a cure are specified in the relevant Aircraft Repair
Manual. The maximum curing temperature must not be exceeded. Typical time and temperature is
60ºC for 8 hours.
The use of pressure is normally specified for a repair whilst it is being cured. This assists in
maintaining the correct profile of the repair and improves the bond. Pressure may be applied by
clamps, weights or by a vacuum bag.
Once the resin has cured, it is absolutely neutral. It will not swell or shrink with changes in climate and
is only attacked by a few chemicals.
7.6.7 GEL COAT
The durability and appearance of a glass fibre moulding is dependent on its exposed surface. The
purpose of the gel coat is to provide a resin rich covering of the exposed surface of the laminate. This
prevents the outermost glass fibres of the laminate from becoming exposed to attack by moisture and
sunlight. If the gel coat is pigmented, a solid coloured surface is also given to the laminate.
Generally, the gel coat surface is incorporated in the moulding process but it may also be used as a
paint and after curing polished to give a smooth glossy surface.
7.6.8 FILLERS
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).
7.6.9 STORAGE OF GRP MATERIALS
GRP materials are expensive and to ensure maximum shelf life, they should be stored in proper
conditions.
7.6.10 STORING RESIN
Most laminating resins have a limited shelf life which is specified by the manufacturer. In general they
should be stored in airtight tins at a cool temperature (usually below 10ºC). It should be removed from
storage at least 24 hours before use to allow it to assume workshop temperature. Depending on the
type of resin, the shelf life may be up to 12 months after which it must be discarded. Resins which
have absorbed moisture and become cloudy should normally be discarded but they can sometimes be
recovered by heating them to 120ºC to evaporate the moisture. If the resin clears on cooling, it may be
used but if it remains cloudy, it must be rejected.
7.6.11 STORING HARDENER
Hardeners generally react with oxygen in the air and must be stored in airtight containers. Some
hardeners may crystallise if they become cold. To liquify the hardener it should be gently warmed and
then allowed to cool at room temperature.
Note. The catalyst and accelerator of a three-part laminating resin should be stored separately to
avoid inadvertent contact.
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7.6.12 STORING FABRICS
Glass fabric should be stored in a warm, dry atmosphere free from dust, oil or other contaminants. In
order to preserve the fibre surface treatment it must no get damp.
Before use it is recommended that the fabric is heated to 45ºC in an oven to drive off any moisture that
may be in the fabric.
Pre-preg fabrics should be stored in refrigerated conditions.
All fabrics should be stored in their original wrappings.
7.6.13 SAFETY PRECAUTIONS
The chemicals used in laminating resins and cleaning agents are hazardous substances and extreme
care is called for when handling them.
Most resins are irritant to the skin. Many people are allergic to the resin and repeated skin contact can
cause serious damage. If symptoms of an allergy appear when the resin is used, further contact
should be avoided and the symptoms should slowly fade away.
Direct skin contact with the resin should be avoided and rubber, or plastic gloves worn when there is a
possibility of the hands becoming contaminated. The resins and solvents used in glass fibre are all
poisonous so every precaution should be taken to keep them away from food. The face and especially
the eyes should also be protected from resin and its solvents.
If a rotary grinder is used on a glass fibre laminate, much glass and resin dust will be produced and a
respiratory mask should be worn for protection. The same dust is likely to cause an irritant skin rash to
develop on the forearms, especially when glass fibre is being hand sanded.
Before washing hands and arms after working with GRP it is advisable to rinse them in cold water. The
arms should be washed in soapy water and the operation should avoid scratching, especially while
dust is lying on the skin.
7.6.14 DAMAGE NECESSITATING MANUFACTURERS LIAISON
Damage occurring in certain areas of an aircraft will be beyond the scope of the organisation to effect
repairs. In such cases it may be necessary to liaise with the manufacturer.
The Slingsby T67 Firefly is used as an example where such liaison would be necessary.
7.6.14.1 Non-repairable areas
Repairs in these areas must be approved by the manufacturer. Non-repairable areas of the T67 are
shown in the diagram below.
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7.6.14.2 Extensive skin repair
If large areas of skin require repair, it will be difficult to reform the correct surface profile without proper,
rigid moulds. Also the structure may be weakened by the extensive removal and repair of load bearing
skin. Normally, in such cases, replacement moulded sections are available from the manufacturer.
7.6.14.3 Repairs involving glass fibre rovings
Generally these areas may be repaired by the manufacturer only. To determine the repairability and
exact method of construction, full details should be submitted to the manufacturer.
7.6.14.4 Fittings Requiring Jigging for Positional Location
Fittings that are torn from position may require special jigging to ensure they are correctly located
relative to neighbouring components.
7.6.15 STRENGTH CONSIDERATIONS OF GRP REPAIRS
The strength of a glass fibre repair is dependent on the strength of the bond to the original structure.
Since the repair receives its working loads through this bond, it is imperative that every effort is made
to ensure a sound connection. Some of the important considerations are:
 Correct Surface Preparation.
 Correct Bond Strength. This requires correct procedures to be used during the repair process.
 Uniform Stress. Once again correct procedures during repair will ensure that local stress
concentrations are minimised.
7.6.16 PREPARATION FOR REPAIR
When the damage has been assessed as repairable , preparatory steps may be taken which are
common to most types of repair.
 To determine whether the glass fibres are damage, the gel coat should be removed by grinding it
away or by gently chiselling and peeling it off. Signs of overstraining of the structure will show
up as white cracks in the laminations. If the rear of the structure is accessible, a strong light
shone through the laminates will show up any damage (delamination or cracks) as a dark area.
The affected area should be cut out and the damage treated as a hole.
 The damaged area should be cleaned and then cut back until sound material is reached. No
evidence of whitening or cracking must be allowed to remain.
 Note. Before cutting out the damage, the area should be marked in some way to determine its
orientation for future reference. Any control linkages, bearings etc. should be covered to keep
out glass dust and surplus resin.
 The type and number of glass cloth layers used in the damaged area must now be determined.
This may require the manufacturer to be consulted.
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 It is possible to analyse a sample of material removed from the damaged area by igniting one
corner of the sample with a match or cigarette lighter. This burns off of the resin and allows
individual fabric layers to be separated. The weight and fibre direction may now be determined
and related to the parent laminate by reference to the orientation marks applied in the above
paragraph.
Notes should be made to ensure that the repair will be the same as the original laminate (i.e.
number, weight and direction of each layer). If the structure used a core material, the type and
thickness should be noted. If the core is wood, the grain direction should be noted.
 The patch edges may now be prepared according to the particular repair being followed (scarf or
stepped). Any surface that will have fibre bonded to it must be prepared in accordance with
chapter ‘Surface Preparation’.
When preparing a chamfered (scarfed) edge, the sanding direction should be towards the tip.
The prepared edges should be examined for any sign of delamination which must be removed by
further sanding.
Note. Some manufacturers specify that cut-outs should have radiused corners, others permit
square corners.
 The inside of the structure should now be cleaned out and any loose pieces of glass fibre and
accumulations of dust removed.
7.6.17 SURFACE PREPARATION
The area on which is to be carried out must now be thoroughly degreased with acetone or methyl ethyl
keton (MEK). Once cleaned the area should not be touched with bare hands. All paint, gel coat etc.
must be removed from the repair area. The following procedure should then be adopted.
 The repair area should be thoroughly abraded using glass or garnet paper. The object of this
abrasion is to remove the top film of resin from the glass and slightly roughen the glass fabric so
that it becomes whiskery.
Note. Care must be taken to ensure that not too much of the glass fabric is abraded.
 Remove any dust with a clean cloth.
 Degrease the newly exposed surface to remove any traces of wax or grease. A clean cloth
saturated with clean acetone or MEK should be used to wipe the surface.
 The acetone must be allowed to evaporate from the surface. Careful use of a hot air blower is
recommended to drive off any traces of acetone that may be trapped in the surface fibres.
 Having cleaned the surface, the repair should commence as soon as possible.
7.6.18 TECHNIQUES OF LAMINATING GLASS FIBRE

Pre-shaped templates are used to cut out the required pieces of cloth for the repair.

The workshop temperature must be between 15ºC and 23ºC with a relative humidity of not more
than 65%.

The quantity of resin required should be estimated and mixed in the correct proportions of resin and
hardener according to the manufacturer’s instructions. The container in which the resin is mixed
must be clean and there must be no possibility of the container contaminating the contents. For
this reason ‘unwaxed paper cartons’ are recommended.
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If the resin is for structural repair work, a small sample (about 1cc) of mixed resin is now cast in a
container made from aluminium foil. The sample should be labelled and placed aside to cure for
later inspection.

A coat of resin is brushed onto the prepared surface and the first layer of cloth is placed on the
resin. The cloth is stippled into the resin ensuring that the cloth weave pattern is not disturbed and
that all the air bubbles are worked out.
The brush used for stippling should be slightly wet with resin which will allow the cloth to ‘wet out’
more quickly and help to prevent the cloth sticking on the brush.
Note. Beware of using too much resin as this will result in a resin rich and heavy repair. Ideally
there should be just enough resin in a laminate to wet out the cloth. The fibres when correctly
wetted out are almost invisible.

The edges of the cloth are trimmed to ensure that the repair only covers the correct area. This is
done by lifting the edge of the patch and removing the excess with a sharp pair of scissors.

Each subsequent layer of cloth is then positioned and stippled into the preceding layers (trimming
as necessary) until the laminate is complete.

When laminating is complete, the repair must be allowed to cure without any further disturbance.
7.6.19 PRE-WETTING GLASS FIBRE
There are a few occasions when carrying out repairs to aircraft that the use of pre-wetted cloth is
expedient. The cloth is laminated on flat cellophane or plastic film (up to four layers may be laminated
at once).
The pre-wetted cloth is then transferred to the job and stippled in place. The plastic film is then peeled
off.
The following points must be noted:
 Care must be taken to ensure that the pre-wetted cloth produces a good bond to the parent
material.
 The plastic backing film should be peeled off as the cloth is being laid because, with it in place,
the laminations cannot assume a double curvature or irregular shape.
 It is importance to ensure that no bubbles are trapped. It is difficult to detect bubbles in a multilayer lamination.
 The edges of each cloth layer must be staggered so that there is not an abrupt end to a number
of layers.
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7.6.20 REPAIR TO GRP SKIN
To ensure that the edges of the cloth are staggered and do not form an abrupt end to a number of
layers, two basic repair techniques are used, these are the stepped and scarf repairs. The techniques
to use will be determined by the manufacturer.
7.6.20.1 Temporary Repairs
Temporary repairs may be permitted in certain circumstances, e.g. when proper repair materials are
not available, but they should always be replaced by permanent repairs at the earliest opportunity. The
repairs may take the form of doped on fabric patches or bolted aluminium alloy plates.
Note. Rivets, other than those specially designed for laminated structural repairs, should not be used.
As the rivet shank swells during forming, the loads imposed on the skin may induce delamination.
7.6.20.2 Scratches, Pits and Dents
These are considered to be minor damage providing that they do not penetrate the glass cloth. They
may be repaired by filling with a mixture of resin and hardener which should be allowed to cure before
being sanded down to a smooth contour.
Note. Where doubt exists in the case of a dent, ensure that no delamination has occurred between the
layers of glass cloth.
7.6.20.3 Cracks
Cracks below a certain length (typically about 3”), including those which completely penetrate the
laminate, may be repaired by external patching. The ends of the crack should be stop-drilled with a
5/32” drill.
Note. When drilling GRP laminates, care must be taken not to cause delamination. Pressure on the
drill must be kept to a minimum.
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7.6.20.4 Small Holes
Holes not exceeding a given maximum diameter (typically 9.5mm or 0.375 inches) which pass
completely through the skin may be repaired as shown in the diagram below.
 One glass fibre ply should be cut from each side of the laminate in a circular area extending as
least 12mm from the edge of the damage. Great care must be taken to avoid cutting into the
underlying plies.
 Two discs of fibre glass cloth of identical size and weave as the original cloth should be prepared
as repair plies.
 The resin should be mixed and separated into two containers, one part being used to impregnate
the repair plies and the other to be mixed with chopped glass fibres to form the plug.
7.6.21 REPAIRS TO MULTIPLE LAMINATIONS
These repairs are carried out when a number of plies have been damaged or when a hole has been
made in the laminate which is greater in size than that which may be required by a small hole repair.
The method and type of repair will be governed by the structural strength and aerodynamic
requirements.
Note. All the following repairs must be regarded as typical examples only and the relevant aircraft
Structural Repair Manual must always be consulted for specific repairs.
7.6.21.1 Overlay Patch
This type of repair would normally be carried out where aerodynamic smoothness is not important.
 Loose material and fibres should be removed from the damaged area.
 The surface should be prepared as described earlier for an area extending 12mm (0.5 inches) x
number of plies from the edge of the damage.
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 Repair patches, equal in number to the original number of laminations, should be cut from cloth
of the same type as the original.
 The repair is now laid up, commencing with the largest patch, using the laminating technique
described in chapter ‘Technique of Laminating Glass Fibre’. Airbubbles and excess resin should
be removed by covering each patch with cellophane and using a rubber or plastic squeegee.
 The repair is finally covered with cellophane and pressure is applied whilst the resin cures.
7.6.21.2 Flush Repair Patch
When a smooth external finish is required, a flush repair may be carried out in the following manner.
Full or partial penetration can be repaired as shown in the diagram below.
 If penetration is complete, the size of hole in the inner lamination must be determined after
trimming away damaged fibres. If only partial penetration has occurred, the size of hole in the
last lamination to be damaged must be determined after carefully removing damaged outer plies.
Note. Care must be taken not to damage underlying plies during trimming.
 Knowing the size of hole and number of laminations affected, the size of the outermost limits of
the repair can be calculated.
 Mark out and remove the damaged material.
i. In 12mm steps when using a stepped repair.
ii. Using a 30:1 scarf ratio when using a scarfed repair.
 Surface preparation as previously described, is now carried out.
 When carrying out a repair where full penetration has occurred, the backing patch should be
positioned first by applying a thin coat of resin to the patch (ensuring full impregnation of the
fibres) and locating it centrally over the hole. The patch should be stippled into place with a
brush, covered with a sheet of cellophane and supported.
 A thin coat of resin should now be applied to the inside of the backing patch. (If the repair is to
damage which does not completely penetrate the laminate, this coat of resin would be applied to
the first, undamaged ply).
 The repair is now built up using the laminating technique previously described.
 After the final insert has been fitted it should be coated with resin, covered in cellophane and
smoothed using a roller type squeegee to remove air and excess resin. Leave cellophane in
position.
 Apply light, even pressure to the repair until cured.
 When cured remove pressure, remove cellophane, smooth off irregularities using sandpaper and
wipe area with MEK.
 Apply surface finish.
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7.6.22 REPAIR TO SANDWICH STRUCTURE
The term ‘sandwich’ construction describes any structure which consists of two skins attached to each
side of a core material. The skins may be of aluminium alloy, wood, GRP and CFRP and the core
material may be of balsa wood, polystyrene foam or a honeycomb matrix of aluminium foil, paper or
nylon.
It is not possible to cover all combinations in these notes and so they will deal with typical repairs to
sandwich structures of GRP skin with honeycomb and form cores.
Damage to GRP sandwich structure may affect the outer laminations only, the outer laminations and
core or the outer and inner laminations and core.
7.6.22.1 Repairs to Core and One Skin
When the core and one facing have been damaged by physical impact, delamination or water
contamination, the core and laminated skin must be cut back to sound material. The following
procedure is typical.
 Damage to the lamination and core is cut out in the smallest area which will include all the
damage. The shape of the cut out may be specified as circular, oval, square or rectangular.
The core may be cut out using a sharp Stanley knife taking great care not to cut the fibres of the
underlying skin.
A circular cut out may be made by using a circular cutter fitted into a hand drill. The lower skin
should be supported during this operation and the minimum pressure applied to the hand drill to
prevent separation of bonding in the surrounding structure.
 The laminations should be cut back to produce a stepped or scarfed depressions, described
earlier in chapter ‘Flush Repair Patch’.
Note. Repair to a balsa wood or foam core may call for the edges to be scarfed to produce a
longer bond. Typical core scarf ratio 4:1.
 The exposed edges of the core, the inner surface of the lower skin and the stepped or scarfed
upper skin should now be sanded to remove rough edges and loose material. The area for
about 25mm around the edge of the repair should be sanded to remove the gel coat and top
layer of resin to expose the fibres of the top lamination.
 The whole area should be cleared of dust and then cleaned with MEK.
 A section of replacement core material is cut to fit exactly into the cut out. Where scarfed edges
are used, accurate fittings of the scarf joint is essential. The direction of grain of the balsa wood
core should correspond with the original material.
Note. When repairing a honeycomb core, some manufacturers call for the repair piece to be
made slightly (about 3mm) oversize. It is then driven into the cut out, after resin has been
applied by using a wooden spreader and a hammer. The underside of the job must be supported
during this operation and the loose pieces of honeycomb removed with tweezers after fitting.
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 All mating surfaces of the cut out should now be coated with resin and the core pressed into
position.
 The laminations are now built up as described in chapter ’Technique of Laminating Glass Fibre’.
 After curing, the repair area is sanded down and the surface finish is applied.
7.6.22.2 Repair to Core and Both Laminations
When the damage affects the core and both lamination, the repair procedure is the same as for
damage to the core and one lamination but is carried out in two stages as follows.
 Remove damaged area in accordance with ‘Repair to Core and One Skin’ first section, scarfing
or stepping the outer laminates only.
 Sand surfaces of core cut out, laminations and the surface area within 25mm of the edge of the
cut-out.
 A mould should now be made to fit the shape of the inner facing.
 A distance piece should now be cut to fit exactly into the cut out of the inner facing having a
thickness equal to the thickness of the facing.
 The core is now inserted and the outer laminations repaired in accordance with chapter
‘Technique of Laminating Glass Fibre’.
 The mould and distance piece may be removed and the inner laminations stepped or scarfed in
the same manner as the outer laminations.
 The outer facings should now be supported by a mould and the inner laminations built up.
 After curing the repair is sanded smooth and the surface finish applied.

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MODULE 6 - Materials and Hardware
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8.
CORROSION
What Is Corrosion? Metal corrosion is the deterioration of the metal by chemical or electro-chemical
attack, which results from the tendency for a metal to revert to it's more stable natural state. e.g. Iron
Oxide (rust). Corrosion weakens primary structural members and is a major source of structural failure
if left unchecked. Water or water vapour containing salt or other impurities combines with oxygen to
produce the main source of corrosion to aircraft parts.
The appearance of corrosion varies with the type of metal. On aluminium alloys and magnesium it
appears as a pitted or etched surface, often with a grey or white powdery deposit. On Copper and
copper alloys, the corrosion forms a greenish film (Verdigris), on steel, a reddish rust. When the
deposits are removed, each of the surfaces may appear etched and pitted. If this pitting is not deep it
will not significantly affect the strength of the metal; however, these pits may become sites for crack
development. The corrosion may also travel below the surface coatings unseen until the part fails.
8.1
CHEMISTRY & MECHANISMS
Corrosion can be caused by either: Direct chemical attack, or Electrochemical attack. In both cases,
the metal is converted into a metal compound such as an oxide, hydroxide or sulphate.
The corrosion process involves two simultaneous changes. The metal that is attacked or oxidised,
suffers an Anodic change and the corrosive agent is undergoing a Cathodic change. The result is that
material is lost from the Anode and gained by Cathode forming an ionic bond.
8.2
CAUSES
8.2.1 DIRECT CHEMICAL ATTACK
This is caused by direct exposure of the metal surface to Caustic Liquids or gaseous agents such as:
 Spilled battery acids or battery fumes. Spilled acids are less of a problem now Nickel Cadmium
batteries are in common use.
 Flux deposits from inadequately cleaned joints. Flux residues are Hydroscopic (absorb
moisture).
 Entrapped caustic cleaning compounds. Caustic cleaning solutions should be kept capped when
not in use. Many corrosion removal solutions are themselves, corrosive agents and should be
carefully removed after use.
8.2.2 ELECTROCHEMICAL ATTACK
This method of corrosion may be likened to the reaction which takes place in a dry cell battery, in
electroplating or anodising. In its basic form, we need two dissimilar metals in the presence of an
electrolyte which is usually water containing impurities. This forms a simple electric cell in which the
less noble metal is the anode and is eaten away.
All metals and alloys are electrically active and have a specific electrical potential in a given chemical
environment. The constituents of an alloy such as Duralumin (mainly Aluminium and Copper) also
have different potentials and exposure of the alloy to a corrosive medium will cause the more active
metal to become anodic and the less active to become cathodic. The greater the difference in
electrical potential between the two metals, the more severe the chemical attack, if proper conditions
exist for corrosion. For example in the case of Dural, the Copper is more cathodic than the Aluminium
and so the aluminium will erode.
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8.3
TYPE AND SUSCEPTIBILITY
8.3.1 FORMS OF CORROSION
There are many forms of corrosion. The form may depend on metals involved, their function,
atmospheric conditions and corrosive agents present. The following are the more common found on
airframe structures.
Surface
Dissimilar Metal
Intergranular
Exfoliation
Stress
Fretting
Crevice
Filiform
Microbiological
8.3.2 SURFACE
General roughening, etching or pitting of the metal surface, frequently accompanied by a powdery
deposit of corrosion products may be caused by direct chemical or electrochemical attack. Corrosion
may spread under the surface coating unnoticed until the paint or plating is lifted off the surface by the
corrosion products or forms blisters.
8.3.3 DISSIMILAR METAL CORROSION
This may be taking place out of sight and may result in extensive pitting. It may or may not be
accompanied by surface corrosion. It is caused by a potential difference existing between two metals
where plating or jointing compound has been removed or omitted. Examples may be found where
steel bolts, nuts or studs contact magnesium alloys.
8.3.4 INTERGRANULAR CORROSION
This is a direct attack along the grain boundaries of an alloy and results from a lack of uniformity in the
alloy structure, caused by changes occurring during heating and cooling during manufacture or during
heat treatment. Intergranular corrosion often exists without visible surface evidence and so the
structure may be weakened considerably. Intergranular corrosion may often be detected by ultrasonic,
eddy current or radiographic inspection techniques.
8.3.5 EXFOLIATION CORROSION
Very severe Intergranular corrosion may cause the surface of the material to exfoliate. This is lifting of
layers of metal due to de-lamination at the grain boundaries. Exfoliation corrosion often attacks 7000
series alloys - with an appreciable amount of Zinc. Spars, stringers and other high strength parts which
are extruded or hot rolled are often susceptible to this kind of corrosion if they have been poorly heat
treated, because the grains tend to form in layers. Corrosion attacks the grain boundaries, the alloying
agents being more reactive than the base metal. The corrosive by-products tend to force the metal
apart and cause delamination of the material.
8.3.6 STRESS CORROSION
This is a particular form of Intergranular corrosion and results from the combined effect of stresses on
the structure. In highly stressed parts like landing gear components, cracks may originate from a
stress riser such and a scratch or surface corrosion. It is characteristic of aluminium, copper, stainless
steels and high strength alloy steels. It may occur along lines of cold working and signs of stress
corrosion are minute cracks radiating from areas of the greatest stress concentration. Likely areas for
this type of corrosion are U/C jacks, shock absorbers, bellcranks with pressed in bushes or other areas
where parts are a force fit, highly stressed or have residual stresses induced during the forming
process.
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8.3.7 FRETTING CORROSION
This occurs where two mating surfaces are subject to slight relative motion. The movement wears
away protective coatings and removes minute particles which oxidise to form hard, abrasive
compounds which expose further areas of metal to corrosive attack. In it's early stages the debris of
this corrosion forms a black powder. The most likely areas are gears, screw jacks, loose panels,
splined hydraulic pump drives and rivets (when they become loose). It may be serious enough to
cause cracking and fatigue failure.
8.3.8 CREVICE CORROSION
Severe localised corrosion at narrow openings or gaps between metal components, often due to
flexing. Corrosive agents are able to penetrate into the joint.
8.3.9 FILIFORM CORROSION
This is corrosion occurring beneath the paint or protective finish in the form of random threadlike
filaments. The pain or coating often bulges or blisters. Found mostly in clad aluminium (Alclad) where
the cladding has been pierced and corrosion has spread under the surface in thread like lines
unnoticed until it becomes quite severe.
8.3.10 MICROBIOLOGICAL CONTAMINATION
This usually occurs in fuel tanks, on turbine-engined aircraft. A fungal growth may occur in temperate
climates when water is present in the fuel tank. (unless the fuel has an additive to protect against it)
Where fungal growth has formed, there is a probability that corrosion of the tank will occur. (See
A.W.N. 21)
8.3.11 FACTORS AFFECTING CORROSION
Many factors will affect the type, speed of attack, cause and seriousness of metal corrosion. Some are
beyond the control of the aircraft designer or maintenance engineer. Some of them can be controlled!
8.3.12 CLIMATIC
The environmental conditions under which the aircraft is operated and maintained cannot normally be
controlled. The following factors will effect the rate at which corrosion will occur.
 Marine environments (exposure to salt water) will increase rate of corrosion.
 Moisture laden atmosphere as against a dry atmosphere. The USA store hundreds of aircraft in
a desert (dry)atmosphere for emergency war use.
 Temperature considerations i.e. Hot climate against cold climate. High temperatures will
increase the rate of corrosion (all chemical reactions occur faster at higher temperatures).
The worst conditions would be in a hot, wet, maritime environment.
8.3.13 SIZE AND TYPE OF METAL
Some metals corrode more easily than others. Magnesium corrodes readily, whilst Titanium is
extremely corrosion resistant because it oxidises readily. Thick structural sections are also more
susceptible than thin sections because variations in physical characteristics are greater. They are also
more likely to have been cold worked and therefore susceptible to stress corrosion.
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8.3.14 CORROSIVE AGENTS
Foreign materials may adhere to metal surfaces. These 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
8.4
CORROSION REMOVAL
General treatments for corrosion removal include:
 Cleaning and strip protective coat on the corroded area.
 Remove as much of the corrosion products as possible.
 Neutralise the remaining residue.
 Check if damage is within limits
 Restore protective surface films
 Apply temporary or permanent coatings or paint finishes.
If corrosive attack has not progressed beyond the point requiring structural repair, the following points
may be noted.
8.4.1 CLEANING AND PAINT REMOVAL.
It is essential that the complete suspect area be cleaned of all grease, dirt or preservatives. This will
aid in determining the extent of corrosive spread. The selection of cleaning materials will depend on
the type of matter to be removed. Dry cleaning solvent (trichloethane – genclean) may be used for oil,
grease or soft compounds. Heavy duty removal of thick or dried compounds may need solvent
emulsion type cleaners.
General purpose, water removal stripper is recommended for most paint stripping. Adequate
ventilation should be provided and synthetic rubber surfaces such as tyres, fabric and acrylics should
be protected. Remover will also soften sealants. Rubber gloves, acid repellent aprons and goggles
should be worn by personnel carrying out paint removal operations. The following is the general paint
stripping procedure:
 Brush area with stripper to a depth of 1/32 to 1/16 inch. Ensure brush is only used for paint
stripping.
 Allow stripper to remain on surface long enough for paint to wrinkle. This may take 10 minutes to
several hours.
 Re-apply the stripper to areas which have not stripped. Non-metallic scrapers may be used.
 Remove the loosened paint and residual stripper by washing and scrubbing surface with water
and a broom or brush. Water spray may assist, or steam cleaning equipment.
Note. Strippers can damage composite resins and plastics, so every effort should be made to 'mask'
these venerable areas.
8.4.2 CORROSION OF FERROUS METALS
Atmospheric oxidation of iron or steel surfaces causes ferrous oxide rust to be deposited. Some metal
oxides protect the underlying base metal, but rust promotes additional attack by attracting moisture and
must be removed.
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Rust shows on bolt heads, nuts or any un-protected hardware. It’s presence is not immediately
dangerous, but it will indicate a need for maintenance and possible corrosive attack on more critical
areas. The most practical means of controlling the corrosion of steel is the complete removal of
corrosion products by mechanical means. Abrasive papers, power buffers, wire brushes and steel
wool are all acceptable methods of removing rust on lightly stressed areas. Residual rust usually
remains in pits and crevices. Some phosphoric acid solutions may be used to neutralise oxidation and
convert active rust to phosphates, but they are not particularly effective on installed components.
8.4.3 HIGH STRESSED STEEL COMPONENTS
Corrosion on these components may be dangerous and should be removed carefully with mild
abrasive papers or fine buffing compounds. Care should be taken not to overheat parts during
removal. Protective finishes should be applied immediately.
8.4.4 ALUMINIUM AND ALUMINIUM ALLOYS
Corrosion attack on aluminium surfaces give obvious indication, since the products are white and
voluminous. Even in its early stages aluminium corrosion is evident as general etching, pitting or
roughness. Aluminium alloys form a smooth surface oxidation which provides a hard shell which may
form a barrier to corrosive elements. This must not be confused with the more serious forms of
corrosion.
General surface attack penetrates slowly, but is speeded up in the presence of dissolved salts.
Considerable attack can take place before serious loss of strength occurs. Three forms of attack are
particularly serious. These are:
 Penetrating pit type corrosion through walls of tubing.
 Stress corrosion cracking under sustained stress.
 Intergranular attack characteristic of certain improperly heat treated alloys.
Treatment involves mechanical or chemical removal of as much of the corrosion products as possible
and the inhibition of residual materials by chemical means. This should be followed by restoration of
permanent surface coatings.
8.4.5 ALCLAD
Pure aluminium has more corrosion resistance than the stronger aluminium alloys. To take advantage
of this, a thin sheet of pure aluminium is laminated to both sides of the aluminium alloy. The Alclad
surfaces offer good protection and can be maintained in a polished condition. Care should be taken
not to remove too much of the aluminium layer by mechanical methods as the core may be exposed.
8.4.6 TYPICAL PAINTED CORROSION TREATMENT SEQUENCE

Remove oil and surface dirt with the appropriate solvent.

Paint strip the area to be treated.

Remove the products of corrosion using scrapers (taking care not to remove metal) or abrasive
paper (wet or dry) or wire wool.

Neutralise any residual with the appropriate chemical cleaner and then wash off with water. Many
chemical cleaners exist. Deoxidine 202 is a phosphoric acid cleaner used on aluminium alloys. It
should not be used on magnesium alloys. Chromic acid is recommended for magnesium alloys.

Apply protective treatment. This may be Alochrom 1200 for aluminium alloys or Chromic acid
treatment for magnesium alloys.

Restore surface finish.
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8.4.7 PERMANENT ANTI-CORROSION TREATMENTS
These are intended to remain intact throughout the life of the component, as distinct from coatings
which may be renewed as a routine servicing operation. They give better adhesion for paint and most
resist corrosive attack better than the metal to which they are applied.
8.4.7.1 Electro-Plating
There are two categories of electro-plating and they are:
 Coatings less noble than the basic metal. The coating is anodic and so if base metal is exposed,
the coating will corrode in preference to the base metal. Commonly called sacrificial protection.
Cadmium plating or zinc on steel.
 Coatings more noble e.g. Nickel or chromium on steel. These nobler metals do not corrode
easily in air or water and are resistance to acid attack. If the basic metal is exposed, it will
corrode locally by electrolytic action. The attack may result in pitting corrosion of the base metal
or the corrosion may spread beneath the coating.
8.4.7.2 Sprayed Metal Coatings
Most metal coatings can be applied by spraying, but only aluminium and zinc are used on aircraft.
Aluminium sprayed on steel is frequently used for high temperature areas. The process (aluminising)
produces a film about 0.004” which prevents oxidation of the underlying metal.
8.4.7.3 Cladding
Hot rolling of pure aluminium onto duralumin produces Alclad which has good corrosion resistance and
the high strength of the alloy. If the cladding becomes damaged, exposing the core, the material will
corrode easily.
8.4.8 SURFACE CONVERSION COATINGS
These are produced by chemical action. The treatment changes the immediate surface layer into a film
of metal oxide which has better corrosion resistance than the metal. Among those widely used on
aircraft are:
 The Anodising of Aluminium Alloys by an electrolytic process which thickens the natural oxide
film on the aluminium. The film is hard and inert.
 The Chromating of Magnesium Alloys to produce a brown to black surface film of chromates
which form a protective layer.
 Passivation of zinc and cadmium by immersion in a chromate solution.
Other surface conversion coatings are produced for special purposes, notably the phosphating of steel.
There are numerous proprietary processed, each known by it’s trade name e.g. Parkerising,
Walterising.
8.4.9 ACID SPILLAGE
Acid spillage in aircraft can cause severe corrosion. Acids will corrode most metals used in aircraft and
will destroy wood and most fabrics. Aircraft batteries give off acidic fumes and battery bays should be
well ventilated, surfaces in the area should be treated with anti-acid paint. The correct procedure to be
taken in the event of a spillage is as follows:
 Mop up as much of the spilled acid using wet rags, try not to spread the acid.
 If possible, flood the area with large quantities of clean water.
 If flooding is not practical, neutralise the area with the following: 10% by weight Bicarbonate of
Soda with water.
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 Wash the area using this mixture and rinse with cold water.
 To check if acid has been cleaned up, test the area using universal indicating paper (or litmus
paper).
 Dry area completely and examine the area for signs of damaged paint or plated finish and signs
of corrosion especially where the paint may have been damaged.
 Restore damage as appropriate.
8.4.10 ALKALI SPILLAGE
This is most likely to occur from Nickel-iron batteries containing Potassium Hydroxide. These
compartments should be painted with anti-corrosive paint. Removal of the alkali spillage is as follows:
 Mop up as much as possible with a wet rag.
 Swab area with the following mixture which neutralises the alkali and passivates bare metal: 5%
by weight Chromic Acid in water.
 Flood area with clean water avoiding electrical gear.
 Check area for neutralisation with universal indicating paper or litmus paper.
 If okay, dry area and check for corrosion and damaged pain etc.
8.4.11 MERCURY SPILLAGE
Sources of mercury spillage are instruments, switches and test equipment. Mercury can rapidly attack
bare light alloys causing intergranular penetration and embrittlement which can start cracks and
accelerate powder propagation. Signs of mercury attack on aluminium alloys are greyish powder,
whiskery growth or fuzzy deposits. If mercury corrosion is found or suspected, assume intergranular
penetration has occurred and the structural strength is impaired. The metal in that area should be
removed and the area repaired in accordance with manufacturers instructions.
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8.4.11.1 Removal of Mercury Spillage
Ensure that toxic vapour precautions are observed at all times during the following operation:
 Do not move aircraft after finding spillage. This may prevent spread.
 Remove spillage carefully by one of the following methods:
i. Capillary brush method.
ii. Heavy duty vacuum with collector trap.
iii. Adhesive tape pressed onto globules will pick them up
iv. Foam collector pads.
 Try to remove evidence of corrosion.
 The area should be further checked using radiography to establish that all globules have been
removed and to check extent of corrosion damage.
 Examine area for corrosion using a magnifier, any parts found contaminated should be removed
and replaced.
8.4.12 IDENTIFICATION OF METALS
If the nature of a metal is unknown i.e. you don’t know what material it is, it may often be identified by
it’s reaction or lack of reaction to various chemicals.
8.4.12.1 Aluminium Alloys
These are light grey in colour, light in weight. Not affected by Nitric acid, Acetic acid or Ammonia.
Attacked by hydrochloric acid, Sulphuric acid and Alkalis.
8.4.12.2 Magnesium Alloys
These are light in colour, light in weight and attacked by saturated Sulphuric acid solution.
8.4.12.3 Bronzes (Aluminium & Phosphor)
These are usually coppery or reddish in colour. Attacked by Nitric acid to form a solution, which when
boiled produces a white precipitate.
8.4.12.4 Ferrous Metals
In most cases they have a ‘steely’ appearance, except Cast Iron which is black or grey. Most steels
are magnetic, except austenitic steels and some stainless steels.
Heating ferrous particles in near boiling Nitric acid until chemical action ceases produces:
 A yellow or light brown solution if the particles are carbon steel.
 A dark brown solution if the particles are cast iron.
Note. Stainless steel will not be attacked in this test.
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9.
AIRCRAFT FASTENERS
Most aircraft now use unified or metric threads, however, some older aircraft use obsolete British
Association (B.A.), British Standard Fine (B.S.F.) or Whitworth (B.S.W.) thread forms. None of these
are compatible with the unified thread forms. The B.A. (47½º) bolts are small in diameter, the largest
that was in common use was 2B.A .with a major diameter of approximately 3/16”. The B.S.W. and
B.S.F. are larger bolts with 55º thread angle.
It is often disputed as to the difference between a bolt and a screw, so for the benefit of these notes we
will a bolt as a threaded fastener used in conjunction with a nut which has a definite plain portion on the
shank. A screw is threaded all the way.
When defining the length of bolts, we are concerned with the length of plain portion with hexagonal
bolts and as shown in the diagram below for other shaped heads. Screw lengths are also shown in
diagram below.
9.1
SCREW THREAD NOMENCLATURE
The following terms and definitions are commonly used regarding screw threads. Reference shown be
made to the diagram below.
 Major Diameter. The largest diameter of the thread, measured at right angles to the axis.
 Minor Diameter. The smallest diameter of the thread, measured as right angles to the axis.
 Pitch. The distance from the centre of one crest to the centre of the next, measured parallel to
the axis.
 Depth of Thread. The distance between the root and crest , measured at right angles to the
axis.
 Lead. The distance a screw moved axially in one complete turn. In the case of multi-start
threads, the lead is equal to the pitch multiplied by the number of starts.
 Single Start Thread. This is when there is only one screw thread cut in the material.
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 Multi-Start Thread. This consists of two or more separate, parallel threads cut into the material
carrying the thread. This is used in order to achieve a quick acting motion between two threaded
items.
 Runout. The part of the thread where the minor diameter increases until it equals the major
diameter and merges with the plain portion of the shank. The runout cannot be used and nut
rotated on the runout would become thread-bound.
9.2
THREAD FORMS
9.3
BOLTS
9.3.1 BRITISH BOLTS
An extensive range of bolts and screws is provided for, in the specifications drawn up by the Society of
British Aerospace Companies.
Note. The following abbreviations are in common use:
A.G.S
Aircraft General Standard
A.S.
Aircraft Standards
AL. AL.
Aluminium Alloy
B.A.
British Association
B.S.F.
British Standard Fine
H.T.S.
High Tensile steel
H.T.S.S.
High Tensile Stainless Steel
L.T.S.
Low Tensile Steel
S.S.
Stainless Steel
U.N.C.
Unified National Coarse
U.N.F.
Unified National Fine.
9.3.1.1 Coding & Identification of Unified Threads
British Unified thread bolts are coded for length and diameter, i.e. by a number expressing the length of
the plain portion of the shank in 1/10" (excluding the runout in the case of Unified threads), and a letter
indicating the diameter. The following table gives examples of code numbers for unified threads.
Note. The code numbers for unified bolts are greater than 100.
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Unified Bolts and Screws
Standard No.
Description
Material
A102
Hex. Headed Bolt
H.T.S.
A104
Hex. Headed Bolt
S.S.
A111
Hex. Close Tol. Bolt
H.T.S.
A112
Shear Bolt
H.T.S.
A174
100º Csk. Hd. Bolt
S.S.
A175
100º Csk. Hd. Bolt
A1. A1.
A204
100º Csk. Hd. Screw
H.T.S.
A205
Pan Hd. Screw
H.T.S.
Diameter Code Letter
Code
Size
Code
Size
E
1/4"
P
9/16"
G
5/16"
Q
5/8"
J
3/8"
S
3/4"
L
7/15"
U
7/8"
N
1/2"
W
1"
Example. A102 9 E - H.T.S. Hex. Head bolt with unified thread, major  of 1/4" and plain portion of
9/10".
H.T.S.
(High Tensile Steel) fasteners will be cadmium plated.
C.R.S.
(Corrosion Resistant Steel) will be natural colour.
A1.A1.
(Aluminium alloy) fasteners will be dyed green.
The current methods of indicating that an item has a unified thread are as follows: (see diagram below)

Three touching circles marked in a convenient position (machine items).

A shallow recess in the head of a bolt, equal to the nominal diameter of the thread (cold forged
items).

A Dog point on the end of the thread (usually applies to screws).
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9.3.2 AMERICAN BOLTS
9.3.2.1 American Standards and Specifications
In aircraft production, it is essential that specific quality requirements are established for sizes, shapes,
materials and numerous other conditions for the manufacture of components and fasteners. For
components manufactured in America, there are numerous agencies for which aircraft parts are
manufactured, and each one has different standards. Among these standards are AF (Air Force), AN
(Air Force / Navy), MIL and MS (Military Standards), NAS (National Aerospace Standards). The most
widely used standards for aircraft use are AN, NAS and MS standards. Most items of aircraft hardware
are identified by a specification number or trade name and threaded fasteners by their AN, NAS or MS
numbers which usually give material and size information as well as the item type.
9.3.2.2 Classification of Threads
Aircraft bolts, screws and nuts are threaded in the NC (American National Coarse), the NF (National
Fine), the UNC (Unified National Coarse), the UNF (Unified National Fine) thread series. The thread
size is often coded to give the diameter and number of threads per inch. Example. 4-28 indicates a ¼”
diameter thread with 28 t.p.i.
Threads are also designated by class of fit to indicate the tolerance allowed during manufacturer.
Class 1
Loose Fit
Class 2
Free Fit
(used for aircraft screws)
Class 3
Medium Fit
(used for aircraft bolts)
Class 4
Close Fit
(close tolerance bolts)
A class 4 fit would require a spanner to turn the nut onto a bolt, a class 1 fit could be easily turned with
the fingers.
Aircraft bolts may be made from H.T.S., Corrosion Resistance Steel or Aluminium Alloy. Head types
may be hexagonal, clevis, eyebolt, internal wrenching and countersunk and head markings may be
used to indicate other features such as close tolerance, aluminium alloy, C.R.S. or type of steel.
9.3.2.3 AN Bolts
These come in three head styles, hex. Head, clevis and eyebolts. The table below gives an indication
of the various code number.
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AN Bolts - Types
AN No.
3 – 20
Type
Bolt, hex.
head
Material
Steel
C.R.S.
Al. Al.
Process
Thread
Size
Thread
Cadmium
Plated
Nil
Anodised
No. 10 to
1¼”
UNF
21 – 36
Bolt, Clevis
Steel
Cadmium
Plated
No. 6 to
1”
UNF
42 – 36
Bolt, Eye
Steel
Cadmium
Plated
No. 10 to
9/16”
UNF
73 – 81
Bolt, hex.
Drilled
head
Steel
Cadmium
Plated
No. 10 to
¾”
UNF or
UNC
173 – 186
Bolt, close
- tolerance
Steel
Cadmium
Plated
thread &
head
No. 10 to
1”
UNF
For identification purposes the AN number is used to indicate the type of bolt and it’s diameter and a
code is used to indicate the material, length and presence of split pin or locking wire hole as follows:
 Diameter. The last figure or last two figures of the AN number indicates thread diameter, 1 = No.
6, 2 = No.8, 3 = No.10, and 4 = ¼” with subsequent numbers indicating diameter in 1/16”
increments. Thus an AN4 is a hexagon headed bolt ¼” diameter and an AN14 is a hexagon
headed bolt 7/8” (14/16”) diameter.
 Lengths. The length of a bolt in the case of a hexagonal headed bolt is measured from under
the head of the first full thread as shown in the above diagram marked ‘Head Marking’ and is
quoted in 1/8” increments as a dash number. The last figure of the dash number represents
eighths and the first figure inches. So an AN4 – 12 is a ¼” diameter hexagon headed bolt 1¼”
long.
 Position of Drilled Hole. Bolts are normally supplied with a hole drilled in the threaded part of
the shank, but different arrangements may be obtained:
Drilled shank
= normal coding
e.g. AN24 – 15
Un-drilled shank
= A added after dash No.
e.g. AN24 – 15A
Drilled head only
= H added before dash No.
(replacing dash) A added after dash No.
Drilled head and shank = H added before dash No.
MODULE 6 - Materials and Hardware
e.g. AN25H15A
e.g. AN25H15
Page 9-5
 Material. The standard coding applies to a non-corrosion-resistance, 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
 Thread. Where the bolt is supplied as either UNF or UNC threads, a UNC thread is indicated by
placing an A in place of the dash, e.g. AN24A15
9.3.2.4 Special Bolts
See diagram below (Aircraft Bolts). The hexagon headed aircraft bolt (AN3 – AN20) is an all purpose
structural bolt used for applications involving tension or shear leads where a light drive fit is
permissible. Alloy steel bolts smaller than 3/16” diameter and al. Alloy bolts smaller than ¼” are not
used on primary structure. Other bolts may be used as follows.
 Close Tolerance Bolts. This type of bolt is machined more accurately than the standard bolt.
They may be hexagon headed (AN173 – AN186) or have a 100º countersunk head (NAS80 –
NAS86). They are used in applications where a tight drive fit is required (the bolt requires the
use of a 12 – 14 oz hammer to drive it into position.
 Internal Wrenching Bolts. (MS 20024 or NAS 495) these are fabricated from high strength
steel and are suitable for tensile or shear applications. The head is recessed to allow the
insertion of a hexagonal key used for installing or removing the bolt. In dural material, a heat
treated washer must be used to provide an adequate bearing surface for the head.
 Clevis Bolts. The head of a clevis volt is round and either slotted for a standard screwdriver or
recessed for a crossed-pointed screwdriver. This type of bolt is used only for shear loads and
never in tension. It is often inserted as a mechanical pin in a control system.
 Eyebolt. The eye is designed for the attachment of cable shackles or turnbuckles and the bolt is
used for tensile loads. The threaded end may be drilled for safety.
9.3.2.5 Screws
These differ from bolts in that they generally:

Are made from lower strength materials.

Have a loose fitting thread i.e. Class 2.

Are turned into a nut or anchor nut (bolts usually are fixed and the nut is turned).

Have no clearly defined plain portion.
Commonly used screws are classified in three groups:
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i. Structural screws which have the same strength as equal size bolts. They are made of
alloy steel, are properly heat treaded and have a defining grip. They are found in the NAS
204, AN 509 and AN 525 series and may have round, brazier or countersunk heads. AN
509 100º countersunk head screws may be driven with either Phillips or Reed and Prince
screwdrivers. The AN 525 washer-head screw is used where raised heads are not
objectionable and a large contact area is required.
ii. Machine screws which include the majority of types used for general repair. They may be
countersunk, round-head, washer head or filister head and made from low carbon steel,
brass, C.R.S. and Aluminium Alloy. The countersunk type in particular may have a wide
variety of drive slots as illustrated in the diagram below.
iii. Self tapping screws, which are used to attach lighter parts, such as nameplates to castings
and parts in which the screw cuts it’s own thread. A typical type is the Parker-Kalon which
is blunt at the end. Self tapping screws should never replace standard screws, nuts, bolts
or rivets.
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Page 9-7
9.3.2.6 Identification and Coding
Identification and coding used for screws is similar to that used for bolts. An AN or NAS code number
with letters following to indicate the type of screw, material, length and diameter. Examples of AN and
NAS code number follows:
AN 501B – 416 – 7
AN
= Air Force-Navy Standard
501 = Fillister head, Fine thread
B
= Brass
416 = 4/16” diameter
7
= 7/16” length
The letter ‘D’ in place of the 2B2 would indicate the material is 2017 – T aluminium alloy. The letter ‘C’
designates corrosion resistance steel. An ‘A’ placed before the material code letter would indicate the
head is drilled for safety.
NAS 144DH – 22
NAS = National Aircraft Standard
144 = Head style: diameter and thread ¼” – 28, internal
9.4
DH
= Drilled head
22
= Screw length in 1/16” - 1 3/8” long.
wrenching
NUTS
9.4.1.1 Aircraft Nuts
These are made in a variety of shapes and sizes. They can be made of Cadmium plated carbon steel,
stainless steel or anodised 2024 – T aluminium alloy and have right or left hand threads. As they do
not have any identifying marks or lettering, they are usually identified by colour and construction.
Familiar types include the Plain, Castle (Castellated), Slotted, Thin, Light hexagon and Wing nut. The
figure below shows a selection of typical nuts.

Castle Nut. (AN310) These are used with drilled shank hexagon headed bolts or studs, eye-bolts
and clevis bolts. It is fairly rugged and can withstand large tensile loads. Slots (castellations) are
designed to accommodate a split (cotter) pin.

Slotted Nut. Similar in construction to the castle nut and used in similar applications except that
they are normally used for engine use only.

Plain Hexagon Nut. AN315 and AN355 (fine and coarse thread) is of rugged construction and
suitable for large tensile loads. Since it requires an auxiliary locking device, it’s use on aircraft is
limited.
MODULE 6 - Materials and Hardware
Page 9-8

Light Hexagon Nut. AN340 and AN345 is a much lighter nut used for miscellaneous light tensile
requirements.

Plain Check Nut. AN316 is employed as a locking device for plain nuts, threaded rod ends and
other devices.

Wing Nuts. AN350 are used where the desired tightness can be obtained with the fingers and
where the assembly is frequently removed.
9.4.1.2 Hexagonal Stiffnuts
(see diagram below ‘Hexagonal Stiffnuts & Anchor Nuts)

Nyloc. This looks like a standard hexagonal nut, but with a nylon insert in the end. This insert is
initially not threaded and has an internal diameter slightly smaller than the nut thread. As the nut is
screwed on the bolt, the nylon insert is displaced and a high degree of friction is set up on the nut
threads. Another type of nyloc nut is named the ‘capnut’; this type is completely sealed and used in
pressurised compartments and fuel and oil tanks etc.
Note. As the insert is nylon, this type of stiffnut should not be used in high or low temperature
areas. A typical maximum temperature would be 120ºC. A similar type of stiffnut has a fibre insert
instead of nylon and is called a ‘fibrelock nut’. Neither nylon or fibrelock stiffnuts should be re-used.

Oddie. The top of this nut has a slotted end forming six tongues which form a circle slightly smaller
than the bolt or stud diameter. As the nut is fitted, a friction load is imparted onto the thread.

Philidas. This nut has a circular crown which is slotted horizontally in two places. The thread on
the slotted part is slightly ‘out of phase’ with the rest of the thread, so that increased friction is
achieved when the nut is fitted.

Aerotight. This is similar to the Philidas except that the slots are vertical. It’s locking method is
also the same.

Lightweight. The locking section of this stiffnut is slightly oval in shape and so causes increased
friction when the thread passes through it.
Note. Metal hexagonal type stiffnuts may be re-used, provided they are not being used in vital
areas such as flying controls and they still retain their friction effect. A recognised rule for
serviceability is that they are discarded when they can be screwed all the way down using the
fingers.
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Page 9-9
9.4.1.3 Anchor Stiffnuts and Stripnuts
(see diagram below ‘Hexagonal Stiffnuts & Anchor Nuts)
Anchor nuts may be supplied with single or double attachment points and may be fixed or floating in a
cage. The anchor nut may be one unit – stiffnut integral with the base plate or it may be an assembly
comprising stiffnut, cage and base plate. Single attachment types are used in corners or where space
is limited and have two adjacent fixing points. Double anchor nuts have a hole either side of the
stiffnut. They are fitted to the structure by riveting. Where a number of anchor nuts are required, to
secure panels etc. a number of stiffnuts may be fitted into metal strips for ease of securing. Strip
anchor nuts are usually of the floating variety.
Common application are for both types of stiffnut are:
 Attachment of anti-friction bearings and control pulleys.
 Attachment of accessories, anchor nuts around inspection holds and fuel tank openings.
 Rocker box covers and exhaust manifolds.
MODULE 6 - Materials and Hardware
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9.5
STUDS
Studs are metal rods, threaded at each end and are used in situations where it is not desirable or
possible to drill through a part for fitting a bolt and nut. One end of the stud would be screwed into a
blind threaded hole, the other part is held in position by a nut screwed onto the other end of the stud.
Different types of studs are available as follows:
Waisted. The diameter of the plain portion of the waisted stud is reduced to the core diameter of the
threaded ends, making the stud lighter in weight without impairing it’s strength.
Stepped. This is made with one threaded end larger than the other. The large end screws into the
components, which is usually of soft metal, so providing greater holding power. Stepped studs may
also be used as replacements for damaged studs where the stud hole has been drilled out and tapped
to a larger size.
Shouldered. The shoulder on the plain portion of the stud enables the stud to seat firmly onto the
surface of the component, providing more rigidity than that obtain with a normal stud.
9.5.1 FITTING STUDS
A stud must be a good fit and remain in positional when the nut is removed. The use of a locking agent
such as Loctite may be recommended in the M.M., but care should be taken to use the correct grade.
Studs may be inserted by using a stud box and spanner (see diagram below) by fitting locknuts, or by
the use of a stud tool (see diagram below), which can also be used for stud removal. When using the
stud box, an aluminium or copper disc is placed between the locking bolt and the stud. This prevents
damage to the stud.
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The illustration shows an exploded view of the stud tool. When assembles, the cam followers are
contained within the case and are free to move radially within the limits of the slotted holes. The end
plate is pressed into the end of the tool and located or by peening. The stud to be inserted or
extracted, is passed through the hole in the end of the plate until the plain portion of the stud is
positioned within the hole in the cage. The locating screw is adjusted to prevent further entry of the
stud into the tool. When the tool is rotated, the cage tends to remain stationary owing to the light
frictional grip of the cam followers on the stud shank. The rotating cam faces force the followers
inwards, thus providing a tight grip on the stud shank. The stud then turns with the tool in the direction
of rotation.
9.5.2 STUD REMOVAL
Loose or undamaged studs may be removed by the use of lock-nuts, using a spanner on the lower nut,
or by the use of the tool described in the previous paragraph. Those damaged or broken above the
surface may be removed by filing flats on the stud or cutting a slot in the top of the stud so that a
spanner, tap wrenth or screwdriver may be used to remove the broken stud. For stud broken flush with
or below the surface of the component, on the following methods should be used.

Centre pop the centre of the stud. Use a drill half the diameter of the stud, drill a hole centrally in
the stud. Lightly drive in a square taper drift till its edges cut into the stud, then unscrew by using a
spanner on the squared edge of the drift. Do not drive in the drift too hard as the stud will expand
and therefore be more difficult to remove.

Drill in as above method (previous paragraph) a tapping size hole. Tap with a thread of opposite
hand to that of the stud. Insert a bolt into the tapped hole and unscrew by applying a spanner to
the bolt head.

Select the appropriate size screw extractor (see diagram below) and using a drill, the size marked
on the extractor, drill a hole of suitable depth. Using a wrenth to operate the extractor, screw out
the broken thread.

As a last resort, drill through the remainder of the stud with a drill slightly smaller than the core
diameter of the stud, and very carefully re-tap the hole, picking up the original thread.

If none of the foregoing methods are practicable, it may be permissible to drill out the broken portion
of the stud and re-tap the stud one size larger to take an oversize or stepped stud.
MODULE 6 - Materials and Hardware
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9.6
FRICTIONAL LOCKING DEVICES
See diagram overleaf ‘Locking Devices and Methods’.
Vibration can cause a nut to slacken or even separate from it’s bolt or stud. A slack bolt or nut can
cause equipment failure and can become dangerous as a loose article. Various methods of locking
fasteners in position have been devised and are in common use on aircraft equipment.
Many locking devices used are ‘once only’ items and it is importance that the manufacturers
recommendations are followed regarding the number of times a particular device can be used. It is
also important that when replacements are made, the old items or parts of the items are not left in the
aircraft to become a potentially lethal loose article.
In addition to the ‘stiffnuts’ mentioned in chapter 9.4 the following are also utilised:

Spring Washers. These consist of
single or double coils of square
section spring with sharp corners. A
plain washer should separate the
spring washer from the face of the
component to avoid damage to the
component when the nut is tightened.
A spring washer may be re-used
provided it still retains it’s springiness
and sharp corners.

Shake Proof Washers. These are
spring steel washers which have
slanting serrations on their internal
or external edges. The angle of the
serrations is such that the nut will
ride over them when being
tightened, but any tendency to
unscrew will be resisted by the
sharp serrations biting into the
underside of the nut.

Lock-nuts. Thin hexagonal nuts, screwed down
tightly against the normal securing plain nut, or
against the part into which the male thread
component fits. They are also used on control rod
ends or where a rod is screwed into a fitting: they
are seldom used as a locking device for nuts in
aircraft construction.
MODULE 6 - Materials and Hardware
Page 9-13

9.7
Loctite Sealant. This sealant is a penetrating liquid polymer which remains fluid when exposed to
air, bur hardens to a tough plastic when excluded from the atmosphere (anerobic). This hardening
effect is accelerated by contact with metal surfaces. Loctite is used mainly for bolt locking and for
the retention of inserts, such as roller bearings and bushes. Under the appropriate conditions,
Loctite will bond all common metals, glass, ceramics and phenolic plastics, but some plated metal
parts and plastics may require preliminary activation in a degreasing solvent containing a hardening
agent known as ‘Locquic’. Loctite when cured is virtually insoluble; it is resistance to aircraft fuels,
lubricating and hydraulic fluids. After curing, Loctite is comparable to phenolic resins as an
insulator and it’s shear strength is quite well maintained at temperatures between 18ºC and 149ºC
for short periods. Loctite is available in five grades, each varying in strength so that the correct
grade can be selected for a particular application. The shelf life of Loctite is approximately 12
months, but the entry of metal particles will cause the sealant to harden prematurely. Care should
be taken not to contaminate the Loctite in the bottles with Locquic activator.
POSITIVE LOCKING DEVICES
See diagram overleaf ‘Locking Devices and Methods’.
Positive locking can be defined as a locking method that uses a physical barrier to prevent nuts or bolts
loosening, in addition it can be physically checked after installation.
9.7.1 SPLIT PINS (COTTER PINS IN U.S.A.)
A split pin is used with a slotted or
castellated nut. The Nickel steel pin lies
in a slot in the nut and passes through a
pre-drilled hole in the bolt. The split pin
must be a good fit in both the hole and
the slot; the pin is then secured by
bending the legs are shown. Either
method being acceptable. Split pins are
classified by their diameter and length.
Split pins are used once only.
9.7.2 TERRY PINS
These are similar in design to a strong safety pin and are passed through a hole in a bolt, or nut and
bolt and then fastened. They are classified by their gauge and size.
9.7.3 TAB WASHERS
These are thin metal washers with two
or more tabs or projections; one tab
is bent over the work (as shown) or
fitted into a pre-drilled hole in the work
or against a projection, whilst the
other is bent against the face of the
nut. It is not permissible to straighten
the tabs of a tab washer and re-use,
but if a tab has not been previously
used, the tab washer may be re-used.
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Page 9-14
9.7.4 LOCKING PLATES
These are metal plates fitted around the nut
or bolt head after it has been fully tightened.
They are retained by a small set screw.
The set screw should be locked by a spring
washer or wire locking. Locking plates may
be re-used provided they are a good fit on
the nut or bolt head.
9.7.5 TAPER PINS AND PARALLEL PINS
Taper pins with a taper of 1 in 48 and parallel pins, are used on both tubular and solid sections, to
secure control levers to torque shafts and forked ends to control rods, etc. Some taper pins are
bifurcated and the legs spread for locking, whilst other taper pins and parallel pins are locking by
peening, or by forming solid rivet heads. To avoid slackness, the pins are usually assembles in
reamed holes, the heads being supported during the locking process. Careful inspection is required
after fitment of pins, through hollow tubes, to ensure that undue force during the peening has not bent
the pins, thus impairing security of the fittings.
9.7.6 CENTRE POPPING AND PEENING (BURRING)
Nuts can be locked in position by slightly damaging the screw thread of the nut or bolt. This is done
either by means of a centre punch or by peening over using a hammer. In this drastic method of
security, the appearance of the centre pop marks or peening is usually an indication that the particular
nut is nor normally removed. Slotted screws may also be locked in a similar way by punching a metal
burr from the surrounding metal into the slot. Note. This method should only be used when authorised
by servicing schedules, repair schemes or Mod. Leaflets.
9.7.7 WIRE LOCKING
The use of wire for locking has long been a feature of aircraft engineering. It should be understood that
there is more than one reason why wire may be used. Wire may be used to lock components, to
prevent inadvertent operation of a control or switch or to show whether a control or switch has been
operated. The different uses are known as:
 Standard wire locking.
 Restraint wire.
 Tell-tale wire.
 Tell-tale / restraint wire.
MODULE 6 - Materials and Hardware
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9.7.8 METHODS OF LOCKING
For aircraft, their
equipment, engines and
auxiliary equipment, locking
wire must be of an
approved type i.e. of the
correct gauge and be
corrosion resistance. The
following techniques are
equally effective and each
may be used in specific
situations:
 Double twist method.
 Single strand with twist (2.5 turns) at the origination end and closing loop at second end.
 Single strand with closing loop at both ends.
9.7.8.1 Double Twist Method
In this method, one piece of sire is threaded through the lock hole to approximately the mid-length of
the wire and then bent through 180º, the double strand so formed is then twisted together, keeping the
wire taught. The strands should be twisted until just short of the next locking wire hole. After inserting
the wire in the last hole, the wire should be locked off with about five twists and the remainder cut off.
The cut end should be bent to form a loop, thus preventing snags. Only three components should be
locked together by this method.
9.7.8.2 Single Wiring Method
This method may be used where lightly loaded adjacent parts may be locked together. A typical
application might be a circle of screws, or a series of electrical screws holding a cover plate on. This
method would be more convenient than the double twist method as more than three individual items
can be locked together. The maximum practical number of items is that which can be locked together
by a single 24” strand of wire.
MODULE 6 - Materials and Hardware
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9.7.9 WIRE LOCKING PRINCIPLES
A high standard of wire locking can only be achieved by practice, the following basic principles should
be adopted from the outset:
A. Locking wire should never be re-used and must therefore be renewed whenever disturbed.
B. Caution should be observed during twisting to keep the wire tight without over-stressing or
allowing it to become kinked, nicked or otherwise mutilated. Abrasions caused by pliers are
however acceptable.
C. Locking wire should not be installed in such a manner as to cause the wire to be subject to or
cause chaffing or fatigue through vibration, looseness or excessive tension, other than the
tension imposed to prevent loosening.
D. Wire locking of flexibly mounted components shall be so arranged that neither the flexibility of
the mounting or the efficiency of the locking is impaired.
E. Lengths of wire between points of contact should be kept to a minimum and wherever
possible, less than 3 inches.
The lay of the wire should be such as to resist any tendency for the locked parts to work
loose, taking care to differentiate between left hand and right hand threads.
G. Where locking wire is inserted through a locking hole and bent round the head of an item, the
direction of wrap and twist should be such that the loop round the part comes under the
strand protruding from the hole, so that the loop will not tend to slip up and form a slack loop.
F.
H.
The angle of approach of the
wire is not to be less than 45º
to the rotational axis of the
component being locked,
whilst the line of approach
should be as near as possible
tangential to the arc of
maximum radius.
MODULE 6 - Materials and Hardware
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Where locking tabs are used, they should be aligned with the locking wire in such a manner
as not to impair radial movement of the tab. Wherever possible the close end of the wire
should be in the tab and the open end at the component to be locked.
J. Where a pipe adapter is used, the pipe union is to be locked to the component, not to the
adapter. Adjacent union nuts may be locked together.
K. The use of Lead Seals attached to locking wire is not permitted.
I.
9.7.10 TURNBUCKLES AND ADJUSTABLE STRUT (CONTROL RODS)
See diagram ‘Turnbuckle Barrel or Strut End Safety Indication’ and ‘Control Cable Turnbuckles’ below.
A. Before locking with wire, ensure that the turnbuckles are correctly adjusted, locked, tensioned
and that the fork ends, eye ends or swaged ends have sufficient thread engagement to
ensure safety.
Note. The turnbuckle is considered to be in safety if any portion of the threaded end obscures the
‘safety’ hole. This check may be carried out by inserting into the hole a probe, approximately the same
size as the hole. Obstruction would indicate a safe condition. It is not recommended that locking wire
be used as a probe.
B. Where locking wire is used in an aircraft control system, the working clearances of the system
must account for the dimensions occupied by the locking wire. It is advisable where possible
to tuck the extremities back into the turnbuckle holes to avoid fouling.
C. The normal method of wire locking turnbuckles is known as the figure of eight or double figure
of eight as shown the diagram.
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9.7.11 LOCKING, RESTRAINING & TELL-TALE WIRE ON CONTROLS & SWITCHES
The term ‘Wire Locking’ has frequently been used in connection with controls and switches to describe
three distinct applications, i.e. locking, restraint and tell-tale. The distinction between these functions is
not always clear and is explained as follows:
A. Locking. This term should only be used when a control or switch is held in such a way that it
cannot be operated by the application of any reasonable degree of force. Switches and
controls should be locked if there is a need for them to be operated in flight. Locking wire
should be sufficiently strong to prevent it from being broken by any deliberate or accidental
attempt at operation. When it is necessary to release a control from the locked position, the
wire will normally be cut.
B. Restraint. Restraint wiring is the description applied when a control or switch may be
required to be operated in flight, but requires a protection against inadvertent operation. This
usually means being able to break the wire by the application of a reasonable degree of
manual force. Selection of the correct material and gauge of wire requires considerable care.
Many incidents have occurred because switches and controls have been excessively
restrained rendering normal operation impossible. For restraining purposes it is usual to use
a soft metal such as copper or aluminium which can easily be broken by a reasonable
application of force.
C. Tell-Tale. A tell-tale device is used to indicate that a control or switch has been operated,
even though it may be subsequently returned to it’s original position. Unless it is also
required to provide a restraint, a tell-tale should not restrict operation of the control or switch
in any way. Enamelled copper or bright anodised wire is preferred as these materials assist
in making the state of the tell-tale more apparent.
MODULE 6 - Materials and Hardware
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9.8
MISCELLANEOUS FASTENERS
9.8.1 HI LOCK AND HIGH / TIGUE FASTENERS
These fasteners are basically a threaded fastener that combines the best features of a rivet and bolt. It
consists of two parts, a threaded fastener pin and threaded collar. The Hi Tigue fastener is an up
dated Hi-Lok fastener. The three primary design advantages are:
 Accurate preload and torque within 10%.
 Minimum size and weight.
 Rapid, quiet, single handed operation.
Because the collar breaks off at design preload, use of torque wrenches is eliminated.
The threaded end of the Hi-Lock
pin contains a hexagon shaped
recess. The hex wrench tip of
the driving tool engages the
recess to prevent rotation of the
pin whilst the collar is being
installed, when the
predetermined preload is
reached the hex section of collar
shear off.
The basic part number indicates the assembly of the pin and collar number, example is shown below
(this is for reference only, do not try to remember).
HL 1870–8–12
max grip length in 1/16” (3/4”)
nominal diameter in 1/32” (1/4”)
collar part number
pin part number
Hi-Lok designation
9.8.2 SPECIAL PURPOSE FASTENERS
In addition to the rivets already described, other rivet type fasteners, each designed for a particular
application, are often used in the manufacture and repair of aircraft. Some of these are designed for a
specific use, others may be categorised as ‘High Strength Fasteners’. These fasteners include:
Tubular Rivets, Rivnuts, High Shear Rivets, Jo-bolts and Huck bolts.

Tubular Rivets. These are used primarily to save weight when riveting through tubular or hollow
members when a large part of the rivet is merely passing through space. Tubular rivets are often
used on control rods for connecting end fittings. The rivets are made to AGS drawing
specifications in several materials. The drawing number indicates the type of rivet and the
following letter denotes the material. The number after the letter denotes the dimensions of the
rivet, but has no particular significance as in the case of other types of rivet.

For example AGS 501/H/49 is a tubular rivet in Mild Steel, 1” long with a wall thickness of 26 SWG.
The table below shows the letters used to indicate different materials and the features by which the
materials may be recognised.

MODULE 6 - Materials and Hardware
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Letter
Identification
Material
A
Identification Feature
Protective
Treatment
Physical
Characteristic
Aluminium (L54)
Anodic film
Dyed black
D
Duralumin (L37)
None
Natural colour
H
Mild steel (T26)
Cadmium
plated
Magnetic
J
Nickel alloy
(DTD268) or Monel
metal (DTD204A)
Cadmium
plated
Only slightly
magnetic
K
Monel metal
(DTD204A)
None
Only slightly
magnetic


Rivnuts. These are internally threaded blind, expansion type
fasteners, which serve, primarily as captive nuts where access
is limited to one side of the structure. They receive the screws
for the attachment of de-icing overshoes to aerofoils, thread
material to walkways, carpets to flooring and sound-proofing to
the structure.
Rivnuts are manufactured in two head types, the flat head and the countersunk head, both with
open or closed (for pressurised areas) ends. They may also be manufactured with, or without small
projections (keys) to keep the rivnut from turning. Those without keys would be used where no
appreciable torque loads are imposed.
Installation of a keyed rivnut requires the use of two hand-operated tools; a notching tool for
cutting the locking keyway and a clinching tool for expanding the rivnut. The rivnut hole is first
drilled to the correct size and if a keyway is required, it is cut with the notching tool. The rivnut is
then screwed onto the clinching tool and inserted into the hole. Operation of the lever will then set
the tool.
Identification. Rivnuts are supplied in American thread sizes and in BA or BSF thread forms. To
avoid confusion we shall only consider the American sizes. The countersunk style rivnut is
available with two head angles 100º and 115º, both of these the flat head type are available in three
sizes, 6-32, 8/32 and 10/32. These numbers represent the actual machine screw size of the
internal thread of the rivnut. The outside diameters of the respective rivnuts are 3/16”, 7/32” and
1/4”. Rivnuts are available in six grip ranges, the minimum grip rivnut having a plain head and next
size has a radial dash mark on the head. Each succeeding grip range is indicated by an additional
radial mark on the head. The largest size having five radial dash marks.
MODULE 6 - Materials and Hardware
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The part number code is as follows:
10
KB 106
Grip length in thousandths. i.e. 0.106
Key (K) and Closed end (B). No. key and open end
would be indicated by a dash. i.e. 10-106
Screw thread size. i.e. 10-32
High Shear Rivets (Shear Pin, see diagram to right). As the name suggests, these are used in
situations where there is a high degree of shear loading. The pin is usually manufactured from
Alloy Steel with an Aluminium Alloy collar. Pins are available with flat or countersunk heads in a
range of diameters and lengths. The collar material is impregnated with a lubricant to facilitate it’s
closure and to provide corrosion protection.
Installation. Before placing a High Shear Rivet, ensure that the pin is the correct length. As
shown the diagram below ‘Shear Pin Grip Range’, the parallel part of the pin must be longer than
the total thickness of the work and the trimming edge of the pin must be inside the collar.
High Shear Pin’s are placed by supporting the head of the pin and using a special punch (set) to
close the collar into the groove of the pin. The collar is automatically trimmed to the correct length
by the shearing action of the punch against the trimming edge of the pin.
MODULE 6 - Materials and Hardware
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
Jo-Bolt. This is the trade name for a fastener which is used where a nut and bolt would normally
be fitted but where access is available only to one side of the work. A Jo-bolt (see diagram below ‘
Jo-Bolts and Fitting Sequence’) consists of three components: an alloy steel nut, an allow steel bolt
and a stainless steel sleeve.
Installation. The fastener is installed with a pneumatic or hand operated tool (see diagram ‘Jo-Bolt
Tool’) with which the bolt is rotated and the nut is held stationary. This action expands the sleeve
over the tapered end of the nut and draws the fastened items together. At a pre-determined torque,
the bolt breaks off at a notch-weakened point flush with the head of the nut. A different tool is
required for each of the two head forms and for each diameter bolt.
Identification. Jo-Bolts are made with either 100º csk. Or hexagonal heads to AGS specifications.
The specification drawing number is followed by a four figure size code, the first two figures giving
the diameter of the fastener, in 1/32” and the last two figures giving the mid-point of the grip range
in 1/16”.
Example:
AGS 3817 - 08 02
Normal grip length (1/8”)
Diameter (1/4”)
Alloy Steel, hexagonal headed

Huck Bolt (Lockbolt). These are not to be confused with a Huck rivet, the Huck bolt is a high
strength fastener manufactured from anodised aluminium alloy, or more commonly cadmium
plated steel. The huck bolt combines the advantages of a high strength bolt and a rivet. It is used
in wing joint fittings and landing gear fittings, longeron beams, fuel cell fittings and other major
structural attachments. It is more easily installed than conventional bolts and eliminates the need
for locking devices. Like the rivet, the Huck bolt usually requires a pneumatic hammer or pull gun
for installation and when installed, it is rigidly and permanently locked in place.
MODULE 6 - Materials and Hardware
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When removal of a Huck bolt becomes necessary,
the collar is first removed by spitting with a sharp
cold chistle. The use of a reaction bar on the
other side of the collar is recommended. The pin
may then be driven out with a drift.
Huck bolts are available in three main types, the
pull type, the stump type and the blind type (see
diagram below).
A. Pull Type. These fasteners are used
mainly in aircraft primary and secondary
structures. They are installed very
rapidly with a pneumatic pull gun which
simultaneously pulls the fastener into
position and swages the locking collar
into the locking grooves of the pin.
B. Stump Type. These do not have the extended stem with pull grooves and are used where
clearance will not permit installation of the pull type. A pneumatic riveting hammer with a
special set for swaging the collar into the locking grooves and a reaction (bucking) bar, are
the tools required for installation of stump lockbolts.
C. Blind Type. These lockbolts come as complete assemblies and are used where only one side
of the work is accessible. It is installed in the same manner as the pull type lockbolt.
Common features of the three types of Huck-bolt are the annular locking grooves on the pin and the
locking collar which is swaged into the pin locking grooves to lock the pin in tension. The pins of
the pull and stump type lockbolt are made of heat treated alloy steel or high strength aluminium
alloy.
9.8.3 TURNLOCK FASTENERS (¼ TURN FASTENERS)
This term describes a family of fasteners designed to secure inspection plates, doors and other
removable panels on aircraft. These fasteners may also be described as quick release, or panel
release fasteners. Their most desirable feature is that they permit quick and easy removal of access
panels for inspection and servicing purposes. Many manufacturers make this type of fasteners, typical
types being Dzus, Camloc and Airloc. Most quick release fasteners are designed for use on lightly
loaded panels, but some types are specifically designed for highly stressed panels.

Dzus Fasteners. The Dzus fastener consists of a stud, grommet and spring receptacle. The
grommet is made from Al. Al. and acts as a holding device for the stud. The stud is made from
cadmium plated steel and is available in three head styles; wing, flush and oval. The diameter is
measured in sixteenths of an inch and the length in hundredths of an inch from the head of the stud
to the bottom of the spring hole.
Body diameter, length and head type may be identified or determined by head markings.
MODULE 6 - Materials and Hardware
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Example.
F
6½
50
-
Flush Head
Body diameter in 16th’s of an inch.
Length (50/100th’s of an inch)
A quarter of a turn of the stud (clockwise) locks the fastener. The fastener may be unlocked only by
turning the stud counter-clockwise. A dzus key or a specially ground screwdriver may be used to
lock and unlock the fastener.
Note. It is good practice to mark the position of the correctly locked Dzus fastener by applying a
paint mark in line with the screw slot (on the panel).

Camloc Fasteners. These are made
in a variety of styles and are used to
secure aircraft cowlings, fairings and
panels. Included among this range of
fasteners are heavy duty fasteners
used for stressed panels.
The Camloc fastener consists of a stud
assembly, a grommet and a receptacle.
The stud and grommet are installed in
the removable part i.e. the cowling or
panel and the receptacle is riveted to
the structure of the aircraft. The stud
and grommet are installed in either a
plain, dimpled, countersunk or counterbored hole, depending on the thickness
of the material involved.
A quarter turn (clockwise) of the stud will lock the fastener. The fastener can only be unlocked by
turning the stud counter-clockwise.
MODULE 6 - Materials and Hardware
Page 9-25

Airloc Fasteners. The Airloc fastener
consists of three parts, a stud, a cross
pin and a stud receptacle. The studs
are manufactured from case-hardened
steel with a reamed hole for the cross
pin. The total thickness of material
which may be locked together is
stamped on the head in thousandths.
Studs are manufactured in three head
styles; flush, oval and wing. The
receptacles are manufactured in two
types, rigid and floating. Sizes are
classified by number; No.2, No.5 and
No.7. They are also classified by the
centre to centre distance between the
rivet holes of the receptacles i.e. No. 2
= ¾”. Receptacles are manufactured
from spring steel.

9.9
SOLID RIVETS
An aircraft, even though made of the best materials and strongest parts, would be of doubtful value
unless those parts were firmly held together. Several methods are used to hold parts together; welding
or soldering, threaded fasteners and rivetting being three of the main methods. The use of threaded
fasteners and soldering have been mentioned previously. Rivets are an alternative method of
fastening structure, a rivet being a metal pin on which a head is formed, during manufacture. The other
end or ‘shank’ is placed through two matching holes in the structure and a second head formed,
clamping the structure together. Rivets are normally strong in shear, but they should not be subjected
to excessive tensile loads. There are two main categories of rivet, solid rivets which are ‘set’ using a
rivetting gun on the manufactured head and a reaction (bucking0 bar on the other side, and blind rivets,
which may be installed where access is limited to the other side of the rivet.
Note. British and American rivets are not manufactured to identical specifications or from identical
materials, but British rivets are often used to repair American aircraft and vice versa. Care should be
taken to choose the correct specification rivet and both British and American rivets may be identified by
head and shank end markings or colour.
9.9.1 SOLID RIVETS (BRITISH)
Standards for British Solid rivets are issued by the Society of British Aerospace SBAC (As series) or
the British Standards Institute (SP series). The standards overlap to a certain extent with obsolete
rivets in the AS range being replaced by SP rivets. Rivets are identified by a standard number and a
part number. The standard number identifies the head shape, material and finish. This is followed by
a three or four figure code, the first one or two figures indicating the shank diameter in thirty-seconds of
an inch and the last two, the length in sixteenths of an inch.
Example. As 162-408 would be a 90 degree countersunk, aluminium alloy 5% magnesium rivet, 1/8
diameter and 1/2 inch long. The AS 162 indicating head type and material, the ‘4’ indicates 4/32”
diameter (1/8”) and ‘08’ indicating 8/16” lengths (1/2”).
Tables ‘Material Identification AS Rivets)’ and ‘Typical AS Rivet Specifications Numbers’ below, gives
details on materials and identification marks for the various types of AS rivets. Many of these rivets are
obsolescent and have been superseded by rivets conforming to SP standards.
Table ‘Material Identification (SP Rivets)’ below, gives details of material and identification information
for SP rivets with the standard numbers shown in table ‘Typical Specification Numbers’. It should also
be noted that SP rivets are also available in metric sizes.
MODULE 6 - Materials and Hardware
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Note. It should be noted that the colour coding of all British solid rivets is generally the same for the
same material. For example pure aluminium rivets are Black, Hidimium rivets are Violet, Monel rivets
are Natural and 5% Magnesium rivets are Green. This enables material types to be easily identified.
Material Identification (AS Rivets)
Mat. Spec.
Material
Ident. Marks
Finish
L37
Dural
‘D’ on shank end
Natural
L58
Al. Alloy
‘X’ on shank end
Dyed or Anodised
Green
(5% Mg.)
L86
Hidiminium
‘S’ on shank end
Dyed Violet
DTD204
Monel
‘M’ on shank end
Natural or Cadmium
Plated
Typical AS Rivet Specification Numbers
Material
Spec.
Snap
Mush
90º Csk
100º Csk
120º Csk
90º
Close
Tol.
L37
AS156
AS158
AS161
-
AS164
AS2918
L58
AS157
AS159
AS162
AS4716
AS165
-
L86
AS2227
AS2228
AS229
-
AS2230
AS3362
-
-
AS5462
-
AS465
-
DTD204
Material Identification (SP Rivets)
Mat. Spec.
Material
Ident. Marks
Finish
(On shank end)
L36
Aluminium
‘I’
Black Anodic
L37
Dural
‘7’
Natural
L58
Al. Alloy
‘8’
Green Anodic
(5% Mg.)
L86
Hidiminium
‘0’
Violet
BS1109
Steel
-
Cadmium
DTD204
Monel
‘M’
Natural or Cadmium
Typical SP Specification Numbers
Mat. Spec.
Snaphead
L36
SP77
L37
SP78
SP83
SP69
L58
SP79
SP84
SP70
MODULE 6 - Materials and Hardware
Mushroom
-
100º Csk Head
SP68
Page 9-27
L86
SP80
SP85
SP71
BS1109
SP76
-
SP86
DTD204
SP81
-
SP87
9.9.2 SOLID RIVETS (AMERICAN)
 These are generally used in normal construction and repair work. They are identified by the kind
of material they are made from, head type, shank size and temper condition. Typical head types
are Roundhead, Brazier head, 100º Countersunk head, Flat head and Universal head as shown
in the diagram below. The material used for the majority of aircraft solid rivets is aluminium alloy.
The strength of temper conditions of Al. Al. Rivets are identified by digits and letters in a similar
manner to that used in sheet Al. Al. The normal material grades are 1100, 2017-T, 2024-T,
2117-T and 5056. They may be identified as shown in the diagram below.
 The 1100 rivet is 99.45% pure aluminium and as such is very soft. It would be used for rivetting
lightweight soft aluminium structures where strength is not a factor.
 The 2117-T rivet is made from Aluminium Alloy and is known as the field rivet. It is the most
commonly used rivet mainly because it is ready to use as received and needs no further heat
treatment. It also has a high resistance to corrosion.
 The 2017-T and 2024-T rivets are made from high strength heat treatable Aluminium Alloys.
They are used where more strength is required than that obtained from the ‘field’ rivet. The rivets
need to be heat treated and if nor required immediately, they should be refrigerated until needed.
The 2017-T rivet should be driven within 1 hour of refrigeration (or heat treatment) and the 2024T within 10-20 minutes of refrigeration.
 The 5056 rivet is used for rivetting Magnesium alloy structures because of it’s corrosion resistant
qualities with magnesium.
 Mild Steel rivets are used for rivetting steel parts and Corrosion Resistant Steel rivets are used
for rivetting CRS components in fire-walls and exhaust areas etc.
MODULE 6 - Materials and Hardware
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 Monel rivets are used for rivetting nickel steel alloys. They may also be used as a substitute for
CRS rivets when specified.
 Copper rivets are also available, but their use is limited on aircraft. They may only be used on
Copper alloys or non-metallic materials such as leather.
Corrosion. Most metals including aircraft rivets are subject to corrosion. This may be the result of
local climatic conditions or the fabrication process used. It may be reduced to a minimum by using the
correct materials and by the use of protective coatings on the structure and the rivets. The use of
dissimilar metals should be avoided where possible and the rivet manufacturers usually apply a
protective coating on the rivets. This may be zinc chromate, metal spray or an Anodic finish.
Identification of American Solid Rivets. Some of the identification points have been explained
previously. Identification of a rivet can be aided by a combination of several features or by reference to
the rivet part number. Most of the important identification points are shown in table (12.4.1) and
explain as follows:

Head marking and Colour. These are used as an aid to indicate the material and protective
surface coatings used by the manufacturers. Zinc chromate is usually yellow, an anodise rivet is
usually pearl grey and a metal sprayed rivet is a silvery grey colour. The head markings are as
follows:
Head Markings
Materials
Plain Head
Pure aluminium, 110; Mild steel;
or Copper
Dimpled Head
2117-T Aluminium Alloy
Raised Teat
2017-T Aluminium Alloy
Raised Double Dash
2024-T Aluminium Alloy
Raised Cross
5056
Raised Triangle
Mild steel, countersunk head
Raised Dash
Corrosion resistant steel
Two Raised Teats
Monel
Aluminium Alloy

MODULE 6 - Materials and Hardware
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
Alloy Content. This is designated by letter(s) following the AN standard number describing the
rivet head type, as follows:
A
-
Aluminium Alloy, 1100 or 3003 composition
AD
-
Aluminium Alloy, 2117-T composition
D
-
Aluminium Alloy, 2017-T composition
B
-
Aluminium Alloy, 2024-T composition
C
-
Copper
M
-
Monel
Note. The absence of a letter following the AN standard number indicates a rivet manufactured
from mild steel.

Rivet Head Types:
 Roundhead – used in the interior of the aircraft and has a deep rounded top section. The head is
large enough to strengthen the sheet around the hole and to offer resistance to tension.
 Flathead – used on interior structures where there is insufficient clearance to use a roundhead
rivet.
 Brazier head – has a head of larger diameter, making them suitable for rivetting thin sheet. It
offers only a slight resistance to airflow and is often used on exterior skins, especially on aft
sections of fuselage and empennage. A modified brazier head rivet is also produced which has
a reduced head diameter.
 Universal head – rivet is a combination of roundhead, flathead and brazier head. 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.
 Countersunk head – this rivet is flat topped and bevelled towards the shank so that it fits into a
countersunk or dimpled hole and is flush with the material’s surface. The countersunk angle
may vary from 78º to 120º, the 100º rivet being the most common type. Countersunk rivets are
used to fasten sheets over which other sheets must fit. They are also used on exterior surfaces
of the aircraft because they offer only a slight resistance to airflow and therefore minimise
turbulence.

Part Number. Each type of rivet is identified by a part number so the user can select the correct
rivet for the job. The type of rivet head is identified by an AN or MS standard number. The most
common numbers and head types are:
AN426 or MS20426
-
Countersunk head rivets (100º)
AN430 or MS20430
-
Roundhead rivets
AN441
-
Flathead rivets
AN456
-
Brazier head rivets
AN470 or MS20470
-
Universal head rivets.
MODULE 6 - Materials and Hardware
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After the standard number, the material composition letter(s) are followed by a figure expressing
the diameter of the rivet shank in 32nds of an inch. (see diagram below) The last number(s),
separated by a dash from the diameter number, expressed the length of the rivet shank in 16ths
of an inch.
Example of complete part number:
AN470 AD 3-5
AN470
Universal head rivet
AD
2117-T Aluminium Alloy
3
3/32” diameter
5
5/16” shank length
Note. In the case of countersunk rivets, the length is the overall length.
9.9.3 HEAT TREATMENT
Heat Treatment of Rivets. Metal temper is important in the rivetting process, especially with
Aluminium Alloy rivets. These generally have the same heat treating characteristics as sheet alloys
and can be Annealed and Hardened in much the same manner. The rivet must be soft, or
comparatively soft before a good head can be formed. The 2017-T and 2024-T rivets must be Solution
Treated before being driven and they harden with age.
The process of heat treatment of rivets (Normalising) may be carried out in either an electric air furnace
or salt bath. The temperature range is 495ºC - 505ºC, depending on the alloy. For convenient
handling the rivets are heated on a tray or in a wire basket. After heating for the required period they
are quenched in cold water.
Refrigeration. The heat treated rivet will begin to age harden immediately after treatment and if the
rivets are not to be set immediately they may be refrigerated to delay the age hardening process. The
solution treated rivets are stored at low temperature (below freezing) and under these conditions will
remain soft enough for driving for up to 2 weeks. Any rivets not used in that period should be removed
and 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 the refrigerator. 2017-T
rivets must be driven within 1 hour of refrigeration and 2024-T rivets, within 10 minutes.
9.10
BLIND RIVETS
There are many places in an aircraft where access to both sides of the structure is impossible, or
where limited space will not permit the use of a reaction (bucking) bar. Also in the attachment of many
non-structural parts, such as aircraft interior furnishings, flooring material, de-icer boots etc., the full
strength of solid shank rivets may not be necessary. For use in such places, special rivets have been
designed which can be set from one side only. They are often lighter than solid rivets, yet amply strong
enough for their intended use. The rivets are produced by several manufacturers, both in this country
and in the U.S.A. and have unique characteristics requiring special installation tools and procedures.
The same general basic information about their fabrication, composition, uses, selection, installation,
inspection and removal procedures applies to most of them. The majority of ‘Blind’ rivets can be
described as Mechanically Expanded Rivets and are one of three main types. i.e.
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Self Plugging (friction lock) rivets
Self Plugging (mechanical lock) rivets
Pull Through rivets
9.10.1 FRICTION LOCK RIVET
These are generally fabricated in two parts: i.e. a rivet head with a hollow shank and a stem that
extends through the hollow shank. The diagram ‘Friction Lock Rivets’ shown below, shows typical
‘friction lock’ protruding head and countersunk head rivets. Several events occur in sequence when a
pulling force is applied to the stem of the rivet.
1. The stem is pulled into the rivet shank
2. The mandrel part of the stem forces the rivet shank to expand
3. When friction (pulling action) becomes great enough it caused the stem to fracture at the weakest
point. The bottom end of the stem is retained in the shank giving much greater shear strength than
could be obtained from a hollow rivet.
Note. With this type of rivet, the stem is often designed to break above the rivet head,
necessitating a further action, i.e. cutting off the extra portion of the stem with snips (or a
specialised pneumatic gun) and milling the exposed portion flush with the head. This type of rivet is
going out of style because of the extra processes involved with it’s fitting.
9.10.2 MECHANICAL LOCK RIVETS
See diagram ‘Mechanical Locking Rivets’ below. This type of rivet is similar in design to the friction
lock rivet previously described, except in the manner in which the mandrel is retained in the rivet. This
type of rivet has a positive mechanical locking collar to resist the vibrations that may cause the friction
lock rivet mandrels to loosen and fall out. Also the mechanical locking type rivet stem breaks off flush
with the head and usually does not require further stem trimming when properly selected and installed.
Self plugging mechanical lock rivets display all the strength of solid rivets and in most cases can be
substituted rivet for rivet. Three operations are performed when the rivet is installed (generally using a
pneumatic gun):
1. When pulling force is exerted on the stem, the stem is pulled in forming the blind head and
clamping the sheets of metal together.
2. At a pre-determined point, the inner anvil, incorporated in the gun forces the locking collar into
position.
3. The rivet stem snaps off approximately even with the head of the rivet.
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9.10.3 PULL THROUGH RIVETS
These rivets are sometimes called hollow rivets. When installed, the rivet mandrel is pulled through the
rivet leaving a hollow rivet of much lower strength than the self plugging types. Different types of these
rivets are supplied complete with individual mandrels or individual rivets used with a re-usable steel
mandrel which is drawn completely through the rivet. In some cases, the rivets may be plugged with
sealing pins which give them additional strength as well as sealing them.
MODULE 6 - Materials and Hardware
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9.10.4 GRIP RANGE
Unlike a solid rivet, the part of a blind rivet available to form a head cannot always be seen. It is
therefore necessary to know the range of total material thickness that a given rivet can fasten together.
This is known as the ‘Grip Range’ of the rivet. The diagram ‘Grip Measuring Gauge’ below, illustrates
the use of a gauge to measure the material thickness, used in conjunction with a rivet table.
9.10.5 EXAMPLES OF BLIND RIVETS

Avdel - Friction Lock (British). These are available in Snap head and 100º and 120º countersunk
head, supplied complete with mandrel. When the rivet is fitted, the stepped mandrel fractures,
leaving part of the mandrel in the rivet to form a plug. A pin tester is sometimes used to check the
plug security.
These rivets may be placed using a manually operated tool (Avdel pliers) or by the use of an Avdel
Riveter. The rivet mandrels may be ‘cropped’ using a Cropping tool. After cropping, the exposed
mandrel is trimmed using the river Miller. Avdel rivets are manufactured in L86 Aluminium Alloy
(Hidiminium) either natural colour or dyed violet with anodised Aluminium Alloy mandrels. The rivet
is identified by it’s A.G.S. specification reference which includes a size reference code. For example
AGS 2066 / 508 refers to a 100º csk. Head L86 rivet. The size code is given after the / sign.

5
-
The rivet diameter in 1/32” i.e. 5/32” diameter.
08
-
The rivet length also in 1/32” i.e. 1/4” long.
Chobert – Pull Through (British). These rivets are made with 100º or 120º countersunk heads and
snap heads. They have part parallel and part tapered bores and are expanded with a re-usable,
hardened steel mandrel which is drawn completely through the rivet. They are fitted using either a
hand manipulated tool or a pneumatic gun.
MODULE 6 - Materials and Hardware
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
The rivet differs from the normal blind rivet in that many rivets can be loaded on the mandrel and
the rivets are often supplied pre-wrapped in tubes so that they can easily be loaded. During use,
the mandrel wears and wear limits should be checked with a Go, No-Go gauge to ensure it is
serviceable before use.
Identification. Chobert rivets are
manufactured in Duralumin L37, Hidiminium
L86 and Steel. The Dural rivets are
Anodised (grey), the Hidiminium rivets are
Anodised and dyed violet. The steel rivets
will be Cadmium plated. To give additional
strength and to seal the rivet, sealing pins
are inserted after the rivet is broached. The
rivet is coded with an AGS code number,
the first part of which gives the head type
and material. The second part codes for
diameter and the length as for the Avdel
rivets. Example an AGS 2040 / 410 is a
Cadmium plated steel snap head rivet, 1/8”
diameter and 5/16” long. An AGS 2044 /
619 is a Chobert, duralumin rivet with 120º
csk. Head, 3/16” diameter and 19/32” long.

Tucker Pop (Pop) – British. The rivets are supplied mounted on steel mandrels, the head is pulled
into the rivet expanding it before the mandrel fractures at the waisted portion. This waisted portion
may either be close to the head of the rivet, or part way up the stem. In the first case the rivet will
be classified as ‘Break Head’ (BH) and in the second case, ‘Break Stem’ (BS). The rivets are
placed using a pair of ‘Pop Pliers’ or by the use of a hydro-pneumatic gun. Pop rivets are less
suitable for use on aircraft as they tend to loosen with vibration and then become increasingly
difficult to remove because of the looseness and the presence of the steel mandrel. (They tend to
spin when trying to drill them out).
MODULE 6 - Materials and Hardware
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Installation. This may be a problem when using pop rivets because the mandrel head is not
positively retained within the rivet or drawn completely through it. The mandrel head is often ejected
and may become a loose article. When placing these rivets, the mandrel heads must, if possible be
collected, or driven out and collected. Break head rivets must not be used if the structure is not
accessible to retrieve the mandrel heads. It is sometimes permitted for the mandrels of Break Stem
rivets to be dipped in an adhesive so that they will not vibrate loose after installation. If Tucker Pop
rivets are to be used externally on aircraft, the heads must be sealed to prevent the ingress of dirt
and moisture. Cellulose Metallic Filler is often recommended for this purpose.
Identification. The rivets are manufactured in either Aluminium Alloy or Cadmium plated Monel
with either Dome heads or 100º and 120º countersunk heads. The AGS reference number consists
of the AGS number identifying the material and head type, a three figure size code and letters
specifying Break head or Break stem. In the size code the first figure gives the diameter in 1/32” as
normal, the last two figures gives length in 0.01”. Example AGS 2051 / 537 / BS:
AGS 2051 -
Tucker Pop in Monel with 120º Csk. Head.
537
-
Rivet diameter 5/32” - Rivet length 0.37”
BS
-
Break Stem.
 Imex Rivets – British. This rivet is similar to the Tucker Pop rivet except that it’s shank end is
permanently sealed. The main purpose of this is to make the rivet pressure tight, but it has the
secondary effect of retaining the mandrel head. The rivets are fitted using the same tools as for
pop rivets and are supplied with either ‘long break’ or ‘short break’ mandrels. The long break
mandrels fracture proud of the rivet and need to be trimmed and milled down after broaching.
The short break fracture inside the rivets.
The rivets are manufactured in L58 Aluminium Alloy with either Domed or 120º countersunk
heads. The reference code number is of a different format to normal British rivet - AD / 46 R.
A
-
Aluminium Alloy
D
-
Dome head (K for countersunk)
4
-
Diameter in 1/32” (1/8”)
6
-
Max. thickness of riveted material in 1/32” (3/16”)
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 Cherry Rivets (USA). These rivets are manufactured in all the categories i.e. Friction Lock,
Mechanical Lock and Pull Through. The rivets are broached with individual mandrels which
fracture at the end of the broaching operation. The most commonly used Cherry rivet is
manufactured under the trade name ‘CherryLock’ which indicates that it is a mechanically locked
rivet. Cherry also manufacture Friction Lock and Pull Through rivets under the ‘Cherry MS’
name. The last main type of Cherry rivet is the ‘CherryMax’ which is the most modern type and
is a sophisticated mechanically locked rivet. All of the rivets may be set using a hand riveter or
one of a selection of Cherry pneumatic riveters (see diagram below).
Types of Cherry Rivet
 CherryLock. This is a mechanically locked (contains locking ring) rivet and may be a
Standard CherryLock or a Bulbed CherryLock. The Bulbed CherryLock was developed from
the Standard variety initially for high vibration area’s and thin sheets because it has a larger
than normal ‘bulbed’ blind head. See diagram below.
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 CherryMax Rivets. This is the latest type of Cherry rivet and gives the strongest and most
vibration resistant riveted joint. The main feature of this rivet is that it has an individual
driving anvil for each rivet, ensuring correct fitment of the locking collar every time.
Installation procedure
1. Insert CherryMax rivet into prepared hole. Place pulling head over rivet
stem and apply firm, steady pressure to seat the head. Actuate the tool.
2. Stem pulls into the rivet sleeve and forms a
large bulbed blind head; seats the rivet head
and clamps the sheets tightly together. Shank
expansion begins.
3. 'Safe-Lock' locking collar moves into rivet sleeve recess. Formation of
blind head is completed. Shearing has sheared from cone, thereby
accommodating a minimum of 1/16" in structure thickness variation.
4. Driving anvil forms 'Safe-Lock' collar into head recess, locking stem and sleeve securely together.
Continued pulling fractures stem, providing a flush, burr-free, inspectable installation.
MODULE 6 - Materials and Hardware
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Head Styles
 Cherry MS. These are made with three standard head styles 100º countersunk, universal
and modified truss head.
The rivets are also available as self plugging (friction lock) or hollow (pull through). This
type of rivet was superseded in 1960 by the Mechanically locking CherryLock and more
recently, the CherryMax rivet. The rivets are, however, still widely used.
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Materials and Identification. Cherry rivets are made from a variety of materials, mainly
5056 Aluminium Alloy and Monel metal. The stems are made from Cadmium plated Alloy
Steel or Inconel (Nickel Alloy). They may be identified by an NAS part number or a Cherry
Rivet part number. Examples of head styles and NAS / Cherry code numbers, with the rivet
materials are shown in the table ‘Bulbed CherryLock Rivets’ below. An example of a typical
part number is also shown below.
NAS1738 B 5 - 4
Maximum Grip Length in 16ths of
an inch (-4 = 4/16” = 1/4”)
Shank dia. In 32nds of an inch i.e. 5/32”
Material B = 5056 Al. Alloy
NAS1738 = Rivet type & head style
i.e. Bulbed, cherrylock, universal head
Cherry rivet grip length
MODULE 6 - Materials and Hardware
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MODULE 6 - Materials and Hardware
Page 9-41
Bulbed CherryLock Rivets
Head Style
NAS Number
Universal
Head
NAS 1738B
1738E
1738M
1738MW
1738C
1738CW
Countersunk
Head
NAS 1739B
1739E
1739M
1739MW
1739C
1739CW
Unisink
Head
Countersunk
Head (156º)
Cherry
Number
CR2249
2239
2539
2539P
2839
2839CW
CR2248
2238
2538
2538P
2838
2838CW
Rivet Material
5056 Aluminium
5056 Aluminium
Monel
Monel, Cad. Plt’d
Inconel 600
Inconel 600, Cad. Plt’d.
5056 Aluminium
5056 Aluminium
Monel
Monel, Cad. Plt’d
Inconel 600
Inconel 600, Cad. Plt’d.
Alloy Steel,
Cad. Plt’d
Inconel 600
Inconel 600
Inconel 600
A286 CRES
A286 CRES
Alloy Steel,
Cad. Plt’d
Inconel 600
Inconel 600
Inconel 600
A286 CRES
A286 CRES
-
CR2235
2245
-
2545
2845
Monel
Inconel 600
Inconel 600
Alloy Steel,
Cad. Plt’d.
Inconel 600
A286 CRES
-
CR250
2840
Monel
Inconel 600
Inconel 600
A286 CRES
MODULE 6 - Materials and Hardware
5056 Aluminium
5056 Aluminium
Stem Number
Page 9-42
 Huck Rivets (see diagram below). These are similar to Cherry Rivets in design. The most
popular type have mechanically locked rivet mandrels and are not to be confused with Huck Bolts
which are a type of High Strength fastener described.
MODULE 6 - Materials and Hardware
Page 9-43
10. PIPES & UNIONS
10.1
PIPELINES
Pipelines used for hydraulic circuits in aircraft fall into two basic categories; rigid and flexible. Flexible
pipelines, often referred to as hoses, are used where relative movements or vibration, occurs between
pipeline and component.

Rigid Pipelines (CAP 562 5-6). Rigid pipelines used in aircraft are normally manufactured from
stainless steel, tungum (high tensile brass) or aluminium alloy.

Flexible Pipelines (CAP 562 LFT 5-5). Modern high pressure hose assembles are designed for
the widest possible application in aircraft. The lining of the hose is manufactured from materials
which are designed to withstand the effects of high pressures, temperatures, oil, fuels, solvents and
other fluids.
The hose is strengthened by the application of high tensile steel wire braiding which ensures
maximum resistance to bursting.
When required for use in areas where high temperatures may be encountered (fire zones), the
hoses are covered with a fireproof sheath which forms an insulation against the flames.
Most flexible hose assemblies are marked along their length with a continuous thin coloured line
which allows the hose to be checked for freedom from twist when it is installed.
10.1.1 PIPELINE CONNECTORS
Standard aircraft parts consisting of union nuts, sleeves, collars, nipples and adapters are used to
connect high pressure rigid pipelines to each other or to components. The end of the rigid pipe is
flared so that it will:
1. Fit over an externally coned adaptor.
2. Accept a spherical ended nipple to fit into an internally coned adaptor.
3. Accept a nipple to join two flared pipes using union nuts and sleeves.
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A flexible hose assembly normally has an end fitting which incorporates a spherical nipple to fit into an
internally coned adaptor or a flared pipe and sleeve.
Sometimes it is necessary to have pipelines which are frequently disconnected for servicing purposes.
In such circumstances it is desirable to use pipeline couplings which will allow disconnection to be
made without the loss of fluid. Self sealing couplings are used which close the pipe ends
automatically.
When tightening, or disconnecting a pipe coupling, two spanners must always be used, one to hold the
sleeve or adaptor and one to turn the union nut. Overtightening must be avoided and when specified,
tightening torques must be strictly observed.
10.1.2 LOCKING
Pipeline couplings must be correctly locked, using the specified locking wire, to prevent them working
loose due to vibration with the resulting loss of fluid.
10.1.3 PRECAUTIONS

Cleanliness. Scrupulous cleanliness of pipelines is essential to protect the system against ingress
by dirt, dust etc., which will cause wear and possible failure of a system.
Prior to assembly the pipeline must be blown through with clean dry air and flushed out with clean
filtered hydraulic oil, of the type in the system.
If a pipe is to be left disconnected for some time the open end must be blanked using plugs and
caps as specified.
Note. The use of rag, tape or paper for blanking off purposes is strictly forbidden.

Support. Pipelines must be supported along their length, by prescribed clips or clamping devices,
to avoid excessive flexing which may cause kinking and distortion. Excessive movement of a pipe
may also cause chafing against the structure which may lead to a pipe failure.
Correct fitting of a flexible hose is essential to avoid sharp bend radii and to ensure that the pipe is
not under tension when fitted. This is achieved by ensuring that the hose is longer (by about 3%)
than the distance between the end fittings to which it is to be connected.
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10.2
HOSES AND HOSE ASSEMBLIES
Hose is the general term applied to flexible tubing used as pipelines in fluid systems for connecting
components which mover or vibrate relative to each other. The hose usually consists of a length of
flexible tube with union end fittings attached to enable it to be connected into a fluid system. When
complete with end fittings the hose becomes a hose assembly with an effective length that is the
distance between the nipple extremities of the end fittings.
10.2.1 CLASSIFICATION
The hose assembly is usually classified by the amount of internal pressure that it can withstand (the
maximum design operating pressure). For most hose the maximum operating pressure varies
inversely with the bore diameter and the larger the internal diameter of the hose, the lower is the
operating pressure. The table below gives the maximum operating pressures for the three classes into
which hose is usually divided.
HOSE PRESSURE CLASSIFICATION
Hose Classification
Maximum Operating Pressure
-lbf / in2
bar
Low Pressure
17.2
250
Medium Pressure
103.4
1500
Over 103.4
Over 1500
High Pressure
10.2.2 CONSTRUCTION
In general, all aircraft system hose is constructed in a manner similar to that illustrated in the diagram
below. Each hose consists of a liner, called an inner tube, which retains the fluid and is reinforced by
external layers of cotton and / or steel wire braid. The reinforcing layers enable the hose to withstand
the internal fluid pressure and often a further layer of cotton reinforced rubber is added to protect the
pressure layers from impact, scuffing and damp.
10.2.2.1 The Inner Tube
The material used for lining (inner tube) a flexible hose should have the following properties:

Flexibility at all temperatures.

Impervious to the effects of, but compatible with, the system fluid.

A smooth bore surface which offers little resistance to fluid flow.

Ability to retain these properties under all operational conditions.
Inner tube lining is classified either as rubber or non-rubber material. The rubber type tube may be
made from any of the following materials.

Buna N. Buna N is a synthetic rubber compound which has excellent resistance to oils and
solvents.

Neoprene. This is a synthetic rubber compound which has an acetylene base. Its ability to resist
the deleterious effects of oil and solvent is not quite as good as Buna N, but it is a tougher material
and because of this, it is used where resistance to abrasion is required.

Butyl. Butyl rubber is manufactured from products of petroleum and it is not suitable for use with
petroleum-based substances. Butyl is highly resistant to ester-based fluids and it is commonly used
in hydraulic system hoses when Skydrol hydraulic fluid is used.
There is a smaller selection of non-rubber tube material and the following material is the one you are
most likely to encounter.
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
Teflon. Teflon is the trade name given to a tetrafluorethylene resin non-rubber material.
This material is suitable for nearly all fluids used in aircraft systems and has a smooth wax-like
surface which offers very little resistance to fluid flow. To further reduce the flow resistance, the
Teflon used in aircraft hydraulic systems hose is impregnated with carbon which gives the material
a black surface. Carbon impregnated Teflon is unsuitable for use in oxygen systems and must
never be used. Plain Teflon hose, which is white in colour, is the only non-rubber hose permitted to
be used in oxygen systems.
Braid Reinforcement. The basic materials used for reinforcement of flexible hose are:
Cotton thread
High carbon steel wire
Stainless steel wire
The threads or wires are grouped in strips and wrapped around the inner tube of the hose to form a
braid as shown the diagram to the left. Various combinations of materials and numbers of
reinforcement layers are possible and the amount used is governed by the internal pressure that
the hose must withstand.
The braid angle is critical and it must be such that, when the hose is pressurised by the fluid, the
tension in the braid caused by the tendency of the hose to swell is balanced by the tendency of the
hose to lengthen so that the hose is stable under load. The neutral braid angle chosen prevents the
hose from continually flexing with changes in pressure, in consequence, its service life is greatly
extended.
The outer cover. The protective outer cover does not add to the pressure resistant capabilities of
a hose assembly, but it protects the reinforcing braid from abrasion and from the deletrious effects
of moisture and other liquids. The outer protective case is not normally used when stainless steel
braid is used.
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Lay lines. It is very important that a hose assembly is never twisted. Any twists, however small,
may damage the hose lining. To enable a twist to be readily detected, most hoses (other than
uncovered metal braided types), have straight lines on the outer cover. These are ‘lay lines’ and
they are moulded into, or painted on, the cover. The lay lines are parallel with the bore and they
may be continuous for the length of the hose or broken by the hose identity markings (see diagram
below). The lay lines will provide instant visual indication of twist (spiralling) which is most likely to
take place as the union nuts are tightened.
Straight hose connections. To allow for shrinking, vibration, movement of parts and ‘whip’, all
straight assemblies should be at least 3 percent greater in length than the maximum distance
between the end fittings to which they are to be connected. In no circumstances must a hose
assembly be under any form of tension.
Where the hose enters the end fittings, sharp bends must be avoided as this causes considerable
local strain and rapid failure of the hose (see diagram below).
10.2.3 PRE-INSTALLATION CHECKS
Before a hose assembly is fitted to an aircraft, it should be examined for evidence of damage and
corrosion and for cleanliness. The part number and date stamp should also be verified. Where
specified by the manufacturer, hose assemblies should be pressure tested before installation.
Where possible, every hose assembly must be examined internally to ensure that the bore is free from
obstruction or damage. One suitable method is to lay the hose straight, place a light at one end and
examine the bore from the other.
If the end couplings have been welded, brazed or soldered, they should be examined for evidence of
any corrosion which may have developed during manufacturer. An Introscope should be used in cases
where direct vision is impracticable.
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10.2.4 INSTALLATION
When installing hoses it should be ensured that they are not permitted to come into contact with other
parts of the aircraft structure or installation, which might cause chafing or electrolytic corrosion. It must
be borne in mind that flexible hoses tend to alter their run when pressure is applied and considerable
‘whip’ occurs under surge conditions, the force experienced during ‘whip’ of a hose is often sufficient to
cause damage to the hose and surrounding equipment.
The serviceability and strength of flexible pipes is considerably affected by the amount of bend. As the
variation in the connecting distances may be considerable between different installations, a check
should be made to ensure that bend radii are not less than the minimum given on the drawing.
There are two classes of minimum bend radii recommended by the manufacturers, the minimum bend
radii recommended for hoses where there is no movement is smaller than that recommended where
there is relative movement between the connections, e.g. a hose assembly connected to a flap or
undercarriage actuating jack will have a greater minimum radii than a hose connecting two rigid
couplings at different angles.
It is important to ensure that the minimum bend radii of hoses fitted to moving parts is never less than
the required minimum throughout the travel of the parts. Correct and incorrect methods of installation
are shown in the diagram below, where the different alignment of the hoses due to movement of
attached parts, will be noted.
10.2.5 HOSE ASSEMBLIES WITH RE-USABLE END FITTINGS
The flexible hose pipe usually becomes unserviceable because of damage to, or deterioration of, the
flexible hose, whilst the end fittings often remain in good condition. The end fittings amount to 80% of
the total cost and therefore, a salvage scheme for such fittings is necessary. Serviceable end fittings
can be removed from unserviceable hose and fitted to a length of new hose. Aeroquip hose
assemblies, now fitted to many types of aircraft, are designed with re-usable end fittings. These end
fittings may be obtained for flared or flareless joints.
An end fitting consists basically of two components, a socket fits tightly over the hose and a tapered
nipple (or insert), when screwed into the hose bore, expands the hose and clamps it firmly against the
socket. The is the most common method and is known as a ‘compression seal’ (see diagram ‘Typical
Re-usable End Fittings (a) & (b) below), but a somewhat different method of attachment, known as a
'lip seal'’(see diagram ‘Typical Re-usable End Fittings (c)below), is used by some manufacturers; the
nipple in this case has a cutting spur or separate collar which separates the inner hose from the braid
during the assembly operation. The re-use of end fittings is satisfactory if precautions are taken to
ensure that no damage is caused to the hose bore during the assembly and testing.
A brief description of the assembly technique follows specific instructions in the aircraft Overhaul
manual should be followed with great care.
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10.2.5.1 Hose Assembly
The new hose must first be carefully measured and cut to length with fine-tooth hacksaw, ensuring that
the cut-ends are square and smooth. It should then be thoroughly cleaned and blown out with dry
compressed air.
To minimise fraying when cutting off hose which has a cloth or metal sheath, it is advisable to wrap the
hose with masking tape and saw through the tape.
High pressure hose usually has a metal braid sheath and when this has protective rubber cover, the
cover must often be removed to enable the hose to enter the socket. Using a sharp knife, the cover
should be cut off to the depth of the socket and the exposed braid carefully cleaned up with a wire
brush.
10.2.5.2 Fitting Nipples
To complete the hose assembly, nipples must be screwed into the previously assembled hose and
sockets. This operation must be carried out with extreme care, as misalignment of the nipple could
easily result in its tapered end cutting into the hose wall. Slices of rubber dislodged in this way have
been known to cause malfunction of associated components (see diagram below).
Nipples are usually tapered over approximately half their length and are often provided with a plain pilot
extension to guide the nipples accurately into the hose.
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10.2.6 THE END FITTINGS
Aeroquip hose assemblies may have straight, elbow or adjustable elbow end fittings, as shown in the
diagram ‘Hose End Fittings (a)’ below. The part of the fitting which connects the hose assembly into
the system may take the form of a flared seal, a flareless seal (Globeseal), or a flange-seal (see
diagram ‘Hose End Fittings (b)’ below).
10.2.6.1 Standard End Fittings

Low Pressure Fittings. A typical Aeroquip standard low pressure end fitting is illustrated in the
diagram below. It consists of a nipple, a union nut and a socket, all manufactured in anodised
aluminium alloy.
MODULE 6 - Materials and Hardware
Page 10-8

Medium Pressure Fittings. The component parts of a standard medium pressure end fitting (see
diagram below) are similar to those of the low pressure type but, for a given bore, the nipple and
socket are considerably longer. The additional length allows for more contact area to secure the
hose against the higher pressure. All three parts of medium pressure fittings are in anodised
aluminium but, for the smaller size fittings the nipple and union nut are made in cadmium plated
steel.

High Pressure Fittings. A standard high pressure fitting (see diagram below) is longer than a
medium pressure fitting of comparable bore and is in two parts, an anodised aluminium alloy socket
and a cadmium-plated steel nipple assembly.
10.2.6.2 ‘Super Gem’ End Fittings
Super Gem end fittings are used only on hose with uncovered stainless steel braid reinforcement.
These fittings, which are made in both medium and high pressure types, consists of a socket, a sleeve
and a nipple assembly (see diagram below). Except for certain medium pressure applications, where
aluminium alloy nipple assemblies are used, all the end fittings are in stainless steel.
MODULE 6 - Materials and Hardware
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The diagram below shows a Super Gem end fitting attached to a hose. The sleeve fits between the
inner tube and the reinforcing braid of the hose and has been forced into the tapered bore of the socket
by screwing in the nipple assembly. This action has secured the reinforcing braid and at the same time
produced a leakproof joint by compressing the hose inner tube between the sleeve and the nipple. An
additional seal has also been produced by contact between the outer edge of the sleeve and the radius
on the plain portion of the nipple.
10.2.6.3 Little Gem End fittings
Little Gem end fittings are basically the same as the Super Gem type and like the latter, are used only
on hose with uncovered stainless steel braid reinforcements. The main difference between the two
types is that, with a little gem fitting (see diagram below), the sleeve is an integral part of the nipple
assembly. The sleeve is undercut to provide an annular space for accepting the hose inner tube and it
is also tapered to a sharp edge to form a circumferential cutting spur. During the initial stage of
attaching the end fitting, the cutting spur enters the end of the hose on a line between the reinforcing
braid and the inner tube. As attachment proceeds, the inner tube enters the sleeve undercut and is
forced home to form a positive seal. At the same time the reinforcing braid is lifted over the sleeve and
is secured between the sleeve and the socket.
MODULE 6 - Materials and Hardware
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10.2.7 THE NEW HOSE
The hose used in the manufacture of a replacement hose assembly must be of the same type and
length as the original. If, for some reason, the old hose cannot be used as a pattern for the length of
new hose it will be necessary to refer to the aircraft Overhaul Manual.
Having determined the length of hose required, cut the hose squarely with a hose-cutting machine or a
fine tooth hacksaw (see the diagram below). To avoid ragged ends and to minimise the wire braid flare
(braid fraying), the hose should be wrapped with adhesive tape at the cut-off position before starting
the cut. After the hose has been cut, the cutting residue should be cleared away until the pipe is clean.
MODULE 6 - Materials and Hardware
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10.2.8 PREPARING THE HOSE
It is necessary to strip off a length of the outer cover from each end of the hose, before attaching
Aeroquip standard high pressure end fittings, so that the sockets make direct contact with the
reinforcing braid of the hose. The work can be done manually by following the sequence shown in the
diagram below.
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10.2.9 MATING HOSES & END FITTINGS
The procedure for mating a hose and its end fitting differs with the type of fitting and with the type of
hose to be used. An example of one attachment procedure is shown in the diagram below. Illustrated
are the main stages in the attachment of a standard type high pressure end fitting to an Aeroquip hose
type 309 or 611. The hose is shown cut to length and with the appropriate amount of outer cover
removed.
Joining the pipe end fittings and the hose is completed in the following stages:
1. Cover the cut edge of the hose, the exposed wire braid and the hose fitting with a uniform coat of
the approved sealant (e.g. Aeroquip sealant type AE13696-001 is used with a 309 hose and a
special sealant, made by Dow-Corning, is used with a 611 Skydrol hose).
2. Protect the socket with soft clamps and fit in the jaws of a vice taking care not to distort the socket
by using too much vice pressure. Screw the hose anti-clockwise into the socket until it bottoms
firmly.
3. Liberally lubricate the bore of the hose and the thread and taper of the nipple, with lubricant that is
compatible with the system fluid.
4. Using a ring spanner on the nipple hexagon, screw the nipple into the socket until only a small gap
remains between the hexagon and the end of the socket. The gap must not exceed 1.6mm (1/16”).
The maximum gap for Standard, Medium and Low pressure fittings is also 1.6mm, but other end
fittings differ from this measurement and for details covering the full range, consult the relevant
drawings or Overhaul Manual.
10.2.10
EXAMINATION OF LOCALLY MADE UP ASSEMBLIES
After completing a hose assembly, the inside of the bore should be cleaned of excess lubricant by
using a suitable pencil brush and then blowing through with dry compressed air. Use compressed air
with care and wear goggles to protect the eyes from flying debris caused by the air blast. After
cleaning the hose assembly it must be examined and tested inside and out.
External Examination Details

Ensure that neither the hose nor the end fittings are damaged. Hose with Teflon inner tubes are
particularly easy to damage by kinking. This damage will show as distortion of the reinforcing
braid.
MODULE 6 - Materials and Hardware
Page 10-13

Examine each end fitting for the proper gap between the nipple hexagon or union nut and the
socket.

Check that the union move freely.
The hose must be checked internally for damage as follows:

If the end fittings permit direct viewing, hold the hose out straight towards a good light and visually
examine the inner tube for:
Cleanliness and freedom from obstruction.
Cuts or bulging.
If a visual examination is not possible, the absence of obstructions can be verified by ensuring that a
steel ball, of suitable diameter (80% of diameter), will roll through the hose assembly.
10.2.11
TESTING LOCALLY MADE UP ASSEMBLIES
Locally made up hose assemblies must be tested under pressure and proved serviceable before they
are released for service. This applies to all hose assemblies except those intended for oxygen and
pitot / static systems. Such hoses are usually tested after installation in the aircraft when the complete
system is under test. For other hose assemblies a proof pressure test is carried out on a standard
static hydraulic test rig. The pressure used for this test is normally one and half times the maximum
working pressure of the system into which the hose is to be fitted. For hydraulic hose assemblies the
test fluid should be the same type as the fluid used in the aircraft system.
10.2.12
PROOF TESTING
For proof testing a hose assembly on a hydraulic test rig, use the following procedure:
1. Loosely blank one end of the hose assembly and connect the other end to the rig pressure
manifold.
2. Induce a fluid flow to bleed the hose and when all air has been expelled, tighten the blank.
3. Lay the hose flat and straight and cover with a strong transparent material. Build up the pressure
in easy stages and maintain each stage to check for leaks. If the intermediate stages show no
signs of a leak, achieve the proof test procedure and retain for three minutes. The pressure should
hold without pumping.
When testing a hose assembly that is fitted with an additional protective sleeve, the sleeve may be
unclipped from the end fittings and pulled back to expose the end fitting junctions. This makes
examination for leaks less difficult. However, it is unlikely that there will be any leaks from the hose
assembly if there is no pressure drop over the test period.
After a hose assembly has been proof tested it should be drained of test fluid and the ends sealed
immediately with approved metal protective caps. Hose assemblies which have been proof tested with
a fluid that differs from the system fluid should be purged with clean kerosene and then dried with
filtered compressed air before the protective caps are fitted. If the hose assembly is not for immediate
use it should be stored in a sealed polythene bag.
10.2.13
INSTALLING HOSE ASSEMBLIES
When installing a hose it is important that it is routed so that it has sufficient freedom of movement to
allow flexing but it must be supported so that the end fittings do not carry the weight of the hose. If the
aircraft manual does not include installation details and no layout or general arrangement (GA) drawing
is available, the new hose assembly should be routed and supported in the same way as the original
hose. For general guidance some installation information is given in the following paragraphs.
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10.2.13.1 Pre-Installation Examination
Before installing a new hose assembly, or one which has been removed during servicing, it should be
examined externally to ensure that it is not damaged. Internal examination will depend upon the
circumstances and the bore of a new hose should be examined visually, if the end fittings allow. A
used hose with a permanent set must not be straightened for internal viewing, and if there is any doubt
about the internal condition, the hose assembly must be renewed.
10.2.13.2 Installation Procedure
When a hose assembly is routed correctly and the threads of the unions are in good condition, it should
be possible to engage both union nuts on their mating unions and screw them up with your fingers,
using spanner only for final tightening. Great care must be taken during installation, particularly during
final tightening because the inner tube may be damaged, However, twisting can be avoided simply by
observing the following points:

Position the hose assembly so that the lay line is readily visible and ensure that the line remains
straight as the hose is connected into the system. For uncovered metal braided hose, ensure that
the weave of the braid remains straight.

Hold the hose firmly against rotation as the union nuts are tightened. If the end fittings have
spanner hexagons use a second spanner to prevent the hose from rotating with the union nut. Do
not use a spanner on the socket of the hose (see diagram above).

Do not overtighten the union.
10.2.14
PROTECTIVE SLEEVES
There are some conditions of use when a hose assembly requires additional protection. Examples of
such conditions are found where contact between the hose and the aircraft structure is unavoidable
and where there is considerable moisture, heat and fire potential. For such situations, additional
protection is provided by encasing the hose in a sleeve with abrasion or fire resistant properties. The
protective sleeve is fitted so that it completely covers the flexible hose and overlaps onto the end fitting
where it is secured by special strip steel clamps.
Abrasion-resistant sleeves (see diagram below) are made in the following listed materials:

Synthetic Rubber. Synthetic rubber provides a tough anti-scuff covering sleeve that is resistant to
fuel, oil and ozone. This type of sleeving is used for aircraft hose assemblies and also for ground
equipment hose.

Teflon. Teflon sleeving is used for high temperature applications and is resistant to all fluids used
in aircraft systems or during servicing.
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
Nylon. Nylon protective sleeves are used mainly where a resistance to moisture is required. The
sleeving is made in the form of a spirally-wound strip and can thus be fitted after the hose assembly
has been manufactured. This is in contrast with the rubber and Teflon sleeves which must be fitted
over the hose before the second end fitting is attached.

Polyolefin Tube. Polyolefin is a transparent, heat-shrinkable material and is used where a thin,
closely-fitting abrasion-resistant sleeve is required. The sleeve is large enough, initially, to be fitted
over the hose assembly end fittings and is then shrunk on to the hose by applying heat, usually by
means of a portable electrically-operated heat gun.
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10.2.14.1 Fire Resistant Sleeves
A fire-resistant sleeve, or firesleeve as it is sometimes called, does not increase the long-term heat
resistance of a hose but protects it from the intense heat of a flame long enough for the fire to be
detected and extinguished. Firesleeves are made of knitted or braided asbestos yarn which is
impregnated and coated externally, with a fire-resistant synthetic rubber. The sleeving is produced in
both tubular and strip form as shown in the diagram below. Asbestos, particularly asbestos dust, is a
health risk.
10.3
RIGID PIPES
10.3.1 MANUFACTURE OF RIGID PIPES
These notes give guidance on the manufacturer, testing and inspection of rigid pipes and should be
read in conjunction with CAIPs leaflet BL/6-15 ‘Installation of Rigid Pipes’, also with relevant manuals
for the aircraft concerned.
The efficiency and safety of an aircraft depends to a large extent on the integrity of its pipe systems. It
is essential to ensure that the manufacture of pipes are carried out in accordance with the requirement
of the respective drawings. The method of working metal tubes are dependent on the type of material
and heat treatment.
10.3.2 MATERIALS
Metal tubing for aircraft pipelines is available in various materials and sizes.

Copper and Copper Alloy. Can be flared, brazed or silver soldered. Slight hardening can be
induced during bending operations they can be annealed. Rarely used on modern aircraft.

Aluminium Nickel-Silicon Brass Tubes. Also known as ‘Tungum’, supplied in the annealed
condition can be flared brazed or silver soldered.

High Nickel Copper Alloy. This is used for high to low pressure systems.

Steel Tubes. Available in various grades. Supplied in half hard condition and may be softened for
working purposes and re-heat treated after working. Stainless steel must be used with skydrol.

Aluminium Alloy. The working and heat treatment of these tubes vary. All manufacturing
operations must comply with drawing instructions.
10.3.3 INTRODUCTION
Currently, policy permits pipelines to be made up and tested locally. Authority is normally given when
completed spares are not available and then only proprietary parts of the same specification as the
original equipment are to be used. The material specification, outside diameter and wall thickness of
the tubing to be used will be quoted in the aircraft Overhaul Manual. Making pipes from rigid tubes
usually involve tube bending and swaging or tube flaring.
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10.3.4 BENDING TUBES
When shaping pipelines for aircraft systems, all bends in the tube should be formed on a tube bending
machine. To prevent cracking, some materials need to be softened by heat treatment before making
the bend and after the bend, a second heat treatment may be needed to restore the material to normal
During bending, the tube will require an internal support to prevent it from flattening in the bend area.
When it is available, the original pipeline should be used as a pattern for shape and bend radii or, if it
badly distorted, a manufacturing schedule and drawing will be necessary. However, in tube bending,
no radius should be less than four times the outside diameter of the tube.
For small quantity work it is generally possible, depending on the nature of the pipe material, to make
large radius bends by hand in pipes up to ½ inch outside diameter. For pipes larger than ½ inch
outside diameter, or where considered necessary due to the type of material or radius of bend, it is
usual to fill the pipe with fusible alloy of low melting temperature and to bend the pipe either by hand or
in a bending machine. Sand is generally used as a filler when bending oxygen pipes, so that
contamination with oil is avoided.
Fusible alloy of the type recommended for use on aircraft pipes has a melting point below 100ºC.
Boiling water can, therefore, be used to melt the alloy for loading and unloading. Since the alloy must
not be subjected to a temperature above 100ºC, the alloy containers should be completely surrounded
by water maintained at 85ºC. to 95ºC. Under no circumstances must a flame be used in conjunction
with fusible filler alloys. If a tube requires heat treatment, this must be carried out before filling
operations are started.
10.3.5 UNLOADING
After a tube has been bent to the desired shape, it should be unloaded by completely immersing it in
boiling water to melt the filler alloy. The hot water enters the tube at this stage and care must be taken
to preserve the protection oil film.
As visual examination of the bores of bent pipes is impracticable, if contamination is suspected,
radiographic inspection techniques should be used.
10.3.6 PIPE BENDING MACHINE
These are normally either compression or mandrel machines and may be hand operated, power
assisted or fully automatic.
Compression bending machines are provided with circular formers in various diameters, grooved
around their circumferences to fit a particular diameter pipe. The pipe is bent by rolling a similar
grooved guide round the former, the semi-circular grooves exactly fitting the outside of the tube and
preventing distortion from taking place. When the mean radius of the bend is larger than four times the
outside diameter of the pipe, bending is possible on a compression bending without using a filler but
the insertion of a close-fitting spring may be recommended. A compression bender can also be used
when the mean radius is less than four times the outside diameter provided that a fusible filler alloy is
used to maintain full bore diameter.
10.3.7 COMPRESSION BENDING MACHINES
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The compression bending machine consists of a mounting base, a pulley-shaped bend former, a
grooved tube guide, a tube stop and an operating lever.
10.3.8 PREPARING TUBE ENDS
Rigid tubes are usually joined together by metal faces that require perfect contact if they are to provide
a satisfactory leak-proof joint. Therefore, it is important that the end is trimmed at right-angles to the
axis of the tube. There are special cutting tools used for cutting soft tube accurately. When the end of
a tube has been cut to length and filed true, chamfered and deburred, the tube should be cleaned
internally by using cleaning fluid and a pull through.
10.3.9 FLARING OPERATION
Before a pipe is flared, it must be ascertained that it is of the specified material and in the correct heat
treatment condition for this operation. It is advisable that the pipe should be bent to shape before
flaring.
The pipe end should be square, smoothly finished and clean, a rough or burred edge may cause the
pipe to spit when flared.
The sleeve or union nut and collar, should be assembled on the pipe, then the appropriate half bushed
fitted to the pipe end and clamped in the flaring tool with the pipe end level with the faces of the
bushes. It is most important that the half bushes used are dimensionally accurate and carefully
maintained. If a gap exists between the bushes when they are fitted to the pipe, diametrically opposed
flash lines may be formed on the pipe flare, representing a potential source of failure.
For all materials except stainless steel, the expanding cone of the flaring tool should then be screwed
in until it starts to expand the end of the pipe. At this stage the expanding cone should be rotated by
the handle provided and gently fed inwards until the pipe end is expanded to the limit imposed by the
countersunk half bushes.
When the flare is formed the pipe should be freed from the half bushes and inspected for cracks, splits,
thinning, eccentricity, or other visible faults.
To check the flaring it is recommended that each coupling is connected to a coned adaptor test fitting
and then dismantled. If the test fitting is made from steel, this will allow it to withstand repeated use,
but over-tightening is to be avoided. When assembled, the flare should pass through the union nut
thread with not more than 1/64 inch clearance.
With the collar as far as it will go towards the flared end of the pipe, the projection of the pipe beyond
the collar face must be measured. The tolerances are given in the table below, are permissible.
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10.3.10
Tube Outside Diameter
(inches)
Projection Tolerance
(inches)
3
1
to
16 4
0 to 0.010
5
3
16 to 8
1
1
to
64 33
7
1
16 to 12
1
1
to
32 16
STAINLESS STEEL
Under no circumstances should the expanding cone be used with stainless steel, since ‘pick-up’ and
subsequent damage to the flare may occur. For this material, a single operating tool and a special
buffer lubricant (e.g. Trilac Lacquer) are recommended.
A good flare should be concentric, free from cracks and provide a good seat for the countersink of the
collar. The flare should provide a maximum diameter that is nearly equal to the threaded bore of the
union nut. If the tube has been correctly positioned in the split die the flare should stand proud from
the collar to provide a good grip.
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10.4
IDENTIFICATION
Each pipeline in a hydraulic system is identified by marker tape carrying the word hydraulic and the
international identification system as shown below.
10.5
TESTING & LIFE
10.5.1 HOSES
Test Medium. Pressure tests are usually made with a fluid similar to that which the pipe will carry in
service. However, there are some exceptions, for example, paraffin is usually recommended for testing
petrol pipes as it is safer and more searching. Pneumatic and oxygen pipes are usually tested with
water, this being followed by a further test with air, in which pressure is limited to maximum system
pressure.
Locally made up hose assemblies must be tested under pressure and proved serviceable before they
are released for service. This applies to all hose assemblies except those intended for oxygen and
pitot / static systems. Such hoses are usually tested after installation in the aircraft when the complete
system is under test. For other hose assemblies a proof pressure test is carried out o a standard static
hydraulic test rig. The pressure used for this test is normally one and half times the maximum working
pressure of the system into which the hose is to be fitted. For hydraulic hose assemblies the test fluid
should be the same type as the fluid uses in the aircraft system.
10.5.2 RIGID PIPES
Pipe assemblies should be marked to indicate that they have passed the prescribed pressure test. The
marking should confirm to the method specified for the identification of pipes by the aircraft
manufacturer and will usually be by rubber stamp on the pipe itself, or by metal stamping on a label
attached to the pipe.
10.5.2.1 Pre-Installation Checks
Before pipes are fitted into aircraft, they should be inspected for evidence of damage to the pipe
assembly or the protective treatment, and for external and internal corrosion. If damage or deformation
to the pipe is suspected, the pipes should be pressure tested, or the roundness of the bore checked
(as applicable) as outlined in leaflet BL/6-15.
Such checks are extremely important, since dented or otherwise damaged pipes may cause a
restriction of flow which could have serious consequences.
Checks should be made to established that the pipes are of the specified type and that there is
evidence of their prior inspection and design approval. Approval assembly drawings, installation
instructions and Approved Certificates should be held for reference and record purposes. The
inspector’s stamp should normally appear adjacent to the part number, the method of applying the
stamp should be stated on the drawing.
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Dirt, swarf, dust, etc., introduced by dirty pipes, not only may put out of action the various services of
which the pipe system forms part of, but will increase the wear of the components of those services
and may cause early and complete failure. It is of the utmost importance, therefore, that adequate
precautions are taken at all times to ensure the scrupulous cleanliness of individual pipes and the
complete pipe system.
Prior to assembly, all pipes must be blown through with clean dry air and where applicable, flushed out
with clean filtered fluid of the type to be used in the particular system in which the pipe is to be
installed.
If the pipe is not to be installed immediately, its ends must be blanked, since dirt, swarf or dust may
render a system unserviceable or increase wear in certain components to such an extent as to cause
premature failure.
10.6

UNIONS
Pipe To Pipe. One form of coupling is now standard for the direct coupling of lengths of metal
tubing used in aircraft pipelines.
The standard types of pipeline couplings have the pipe ends ‘flared’ by a special tool, after a collar
an union nut have been fitted over the pipe. Three basic types are illustrated below ‘Standard Metal
Couplings’.
When tightening the pipe couplings, the components of which are made of light alloy (BS L85),
avoid overstressing the screw threads of the sleeve and union nut by excessive tightening. If the
pipes have been flared and the coupling assembled correctly, a pressure-tight joint will be made
without difficulty.
Note. The couplings in the range 1/8 in. to 2 in. outside diameter are easily over-tightened and
careful avoidance of over-stressing must be made. Special care in this respect must be exercised
when tightening the 1/8 in. and 3/16 in. outside diameter sizes.

Pipe to Internally-Coned Adapter. (see diagram ‘Standard Metal Couplings (b)’ above). This type
of coupling provides a means of connecting a pipe to an externally-threaded adapter which is
countersunk at the mouth. The coupling comprises a union nut, a collar and a spherical-ended
adapter nipple. To assemble the coupling, the nut and collar are pushed over the pipe end which is
then flared.
The skirt of the nipple is placed in the flared pipe and the union nut screwed onto the externallythreaded adapter until the nipple is tightly gripped between the collar and the internally-coned
surface of the adapter. To ensure that the flaring of the pipe is satisfactory dismantle the coupling
and examine it.
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
Pipe to Externally-Coned Adapter. (See diagram ‘Standard Metal Couplings (c)’ above), This
type of coupling provides a means of connecting a pipe to an adapter having an externally-coned
nozzle, and comprises a union nut and collar only. To assembly the coupling, the nut and the collar
are pushed over the pipe end which is then flared. The flared end is pushed over the conical end
of the adapter and the union nut is tightened, thus compressing the flared pipe between the collar
and coned adapter. The coupling must be dismantled and examined.

Standard Adapter. (See diagram ‘Bulkhead Pipe Fittings and Banjo-Type Couplings’ (a) on next
page). This is a double-ended fitting with a central hexagonal collar to enable it to be held securely
in the jaws of a spanner whilst pipe connections are made at either end. Standard adapters are
threaded externally and are coned internally or externally at each end. The range of adapter
comprise fittings with both similar (diagram (a)) and dissimilar (diagram (b)) ends to enable various
types of pipe fittings to be coupled when required. Adapters are normally made of light alloy to BS
L85.

Standard Hexagonal Fittings for Bulkheads. (See diagram ‘Bulkhead Pipe Fittings and BanjoType Couplings’ (c) on next page). These are similar to standard adapters except that, at one end,
a longer threaded portion accommodates a nut which is used to clamp the fitting to a bulkhead
through which the pipeline must pass. Two washers are provided to prevent the hexagon collar
and the nut from being tightened against the surfaces of the bulkhead. Bulkhead fittings are made
of light alloy to BL S85 normally are obtainable with similar ends. This type of fitting must be held
with a spanner to prevent rotation when connecting or disconnecting pipes.

Standard Flanged Fittings for Bulkhead. (See diagram ‘Bulkhead Pipe Fittings and Banjo-Type
Couplings’ (d) on next page). These are similar to the hexagonal type except that each comprises
one fixed and one loose flange, each with bolt holes, to enable the fitting to be mounted rigidly on
the bulkhead. In addition to the externally and internally coned fittings, a flanged fitting is available
which has a long internally threaded body in which a banjo union can be screwed (diagram (e)).
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
Standard Banjo Union. (See diagram ‘Bulkhead Pipe Fittings and Banjo-Type Couplings’ on
previous page). This comprises an externally-threaded hollow screw fitting into a single or
doubled-ended pipe connection to which it is sealed top and bottom by bonded seals or soft metal
joint rings. Banjo unions are made in a range to suit sizes from 1/8 to 1.1/4 in. outside diameter
and are normally made of light alloy to BL L85. In the diagram (f), (g) and (h) show the various
types of banjo unions, each of which may either internally or externally-coned connections.

Standard Pipe Union. (See diagram ‘Standard Pipe Unions’ below). A series of elbows, T-pieces
and four-way pipe unions having connections which are either similar or dissimilar in type and size,
are illustrated.

Standard End Plugs and Caps. (See diagram ‘Plugs and Caps for Pipe Ends’ below). When it is
desired to blank off a section of a system, it is normal practice to do so at a union body or at a
component, pipe couplings, nipple plugs and cone caps to AGS1140 and AGS1159 are available
for the purpose. If, however, it is desired to blank off a pipe on which is fitted an inner sleeve, an
adapter nipple must be used in conjunction with a cop cone and an outer sleeve (diagram (d)).
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10.7
FLARES & FLARELESS
10.7.1 FLARELESS COUPLINGS
(See diagram ‘Flareless Pipe Coupling’ below). Flareless couplings are superceding flared couplings in
many applications. The assembly consists of a sleeve, union and pipe nut. Basically, the sleeve is a
close fit over its pipe and the nut forces one end of the sleeve against a (coned) union so the sleeve is
caused to bow and its pilot bits into the pipe.
10.7.2 FITTING PROCEDURES
(New items have to be pre-set)
1. Assemble items onto square and de-burred pipe.
2. Push pipe into its union, or special (steel) pre-setting tool, as far as it will go and screw its nut
finger tight pilot to bite.
3. Tighten a further full turn, this causes sleeve to bow and pilot to bite.
4. Dis-assemble to inspect the sleeve should be correctly bowed and pilot close to or touching pipe, it
should have bitten into pipe.
5. End (axial) play should be less than 0.010 inches. It is permissible if it is able to rotate on pipe.
6. Re-assemble onto union and tighten until distinct increase in torque is felt and tighten a further one
to two hexagon flats (1/16th to 1/3rd turn).
7. If leaking, dis-assemble and inspect. Excessive tightening will damage pipe and sleeve, maybe to
breaking point.
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11. SPRINGS
A spring can be defined as a device capable of deflecting so as to; store energy; absorb shocks;
source of power; measure force and maintain pressure between contacting surfaces.
11.1
TYPES IN USE
The main types of springs in use are leaf, helicoil, spiral and torsional.
11.2
MATERIALS
Most springs are made from tempered steels, but other materials with high Youngs Modulus such as
filament wound carbon fibre are also utilised.
11.3
APPLICATIONS

Leaf springs. These are used on the undercarriages of some small aircraft such as the Cessna
150 and 172 models.

Coil springs. These are the most common form, amongst many other applications are engine
poppet valves and hydraulic relief valves.

Spiral springs. These springs are employed in instruments and plate type valves.

Torsional springs. These springs are used on many undercarriage doors such as those on the
airbus aircraft.
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12. BEARINGS
Bearings are broadly classified into ball bearings and roller bearings. Ball bearings employ steel balls
rotating in grooved raceways, whilst roller bearings utilise cylindrical, tapered or spherical rollers
running in suitably shaped raceways. Example of both types are shown in the diagram below.
12.1
PURPOSE
The purpose of a bearing is to support (generally) moving, components at minimum friction. Some
simple bearings are merely brushes, other multi compound bearings are described below.
12.2
CONSTRUCTION
12.2.1 BALL BEARINGS
There are four main types of ball bearings:

Radial. This is the most common type and is found in all forms of transmission assemblies such as
shafts, gears and control rod end fittings. The bearings are manufactured with the balls in single or
doubled rows and may be rigid or self aligning. They may also be provided with metal shields or
synthetic rubber seals to minimise ingress of dirt or other foreign matter and to retain the lubricant.
They may also be provided with circlips, grooves or flanges to help retain bearing elements. The
balls may be retained in a cage, but filling slots in the outer rings may permit individual insertion of
balls. It is important to note that cages restrict the number of balls and load capacity.

Angular Contact. These are capable of accepting radial and axial loads in one direction. The
axial loading capacity depending largely on contact angle, but running clearances will be greater for
this type than for radial bearings. In applications where axial loads are only in one direction, a
single angular contact bearing may be used, but where axial loads vary in direction, an opposed
pair of bearings is often used. A particular type of angular contact bearing, known as a ‘Duplex’
bearing, is fitted with a split inner or outer ring, and is designed to take axial loads in either
direction. Duplex bearings should never run unloaded. They are not adjustable and radial loads
should always be lighter than axial loads. This is a most efficient form of thrust bearing and is not
speed limited as in the washer type thrust bearing.
Note. When we say ‘axial’ loading , we are concerned with loads parallel to the axis of rotation of
the bearing. Radial loads will be at right angles to the axis of rotation. An example of both types of
load may be illustrated using an aircraft wheel bearing. The normal rotation of the wheel will give a
radial load, but the side loads during turning of the aircraft will give axial loads.

Thrust Bearings. Thrust bearings are designed for axial loading only. The balls are retained in a
cage and run between washers having flat or grooved raceways. Centrifugal loading has an
adverse effect on the bearings and they are, therefore, most suitable for carrying heavy loads at
low speeds.

Instrument Precision Bearings. These bearings are manufactured to a high degree of accuracy
and finish. They are generally of the radial type.
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12.2.2 ROLLER BEARINGS
These may be divided into three groups:

Cylindrical Roller Bearings. These bearings are capable of carrying greater radial loads than ball
bearings of similar dimensions, due to the greater contact area of the rolling elements. The type of
cylindrical roller bearing most commonly used is that in which the diameter and length are
approximately the same proportions. The ‘Needle’ roller bearing has a roller length several times
greater than it’s diameter. These are designed for pure radial loads and are often used in locations
where the movement is oscillatory rather than rotary, such as universal couplings. Needle
couplings are particularly useful where space is limited and may even use the shaft of component
as the inner ring. These bearings are particularly susceptible to the effects of mis-alignment and
lack of lubricant and may also be subject to ‘Brinelling’ due to lack of rotational movement.

Tapered Roller Bearings. These bearings are designed so that the axis of the rollers forms an
angle with the shaft axis . They are capable of accepting simultaneous radial loads and axial loads
in one direction, the proportions of the loads determining the taper angle. Axial load on the rollers
results in high rubbing contact, so good lubrication is essential.

Spherical Roller Bearings. These may one or two rows of rollers which run in a spherical
raceway in the outer ring, thus enabling the bearing to accept a minor degree of mis-alignment.
The bearing is capable of withstanding heavy radial loads and moderate axial loads from either
direction.

Internal Clearances. Radial ball bearings and cylindrical roller bearings are manufactured with
various amounts of internal clearances, so that different tolerances and conditions, are allowed for.
Standard bearings are available in four grades of fit, namely Group 2, Normal Group, Group 3 and
Group 4. Instrument Precision bearings are only available in the first three groups. Bearings are
usually marked in some way to indicate the class of fit, a system of dots, circles or letters often
being used. Replacement bearings should be the same standard.
 Group 2 (1dot) bearings have the smallest internal clearance and are normally used in precision
work where minimum axial and radial movement is required. These should not be used where
operating conditions, such as high temperatures, could reduce internal clearances and are not
suitable for use as thrust bearings or high speed.
 Normal Group (2 dot) bearings are used for most general applications where only one ring is an
interference fit and where no appreciable transfer of heat to the bearing is likely to occur.
 Group 3 (3 dot) bearings have a greater radial internal clearance than Normal Group and are
used where both rings are an interference fir, or where one ring is an interference fit and some
transfer of heat must be accepted. They are also used for high speeds and where axial leading
predominates.
 Group 4 (4 dot) bearings have the largest internal clearances; they are used where both rings are
an interference fit and the transfer of heat reduces internal clearances.
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12.2.3 MAINTENANCE OF BEARINGS

Lubrication. Adequate lubrication is essential for all types of bearings. The purpose of lubrication
is to lubricate the areas of rubbing contact, to protect the bearing from corrosion and to dissipate
heat. For low rotational speeds, or for oscillating functions such as found in a number of airframe
applications (controls surface bearings, control rods), grease may be a suitable lubricant. At higher
rotational speeds where higher temperatures are generated, grease will tend to beak down and
therefore, oil is more suitable. It is important that only those lubricants recommended in the
approved maintenance manuals are used.
 External bearings are often of the pre-packed, shielded or sealed types and are usually packed
with anti-freeze grease because of low temperatures encountered. These bearings cannot
normally be re-packed and when un-serviceable must be rejected. Grease nipples are provided
on some open bearings so that the grease may be replenished at specified intervals or when
grease is lost through the use of solvent, detergents or de-icing fluid. Nipples should be wiped
clean before applying the grease gun, to prevent the entry of dirt. Grease forced into the bearing
will displace the old grease and any surplus grease exuding from the bearing should be wiped
away with a lint-free cloth.
 Wheel bearings are normally tapered roller bearings and should be packed with the correct
grease when re-fitting the wheel.
 Bearings fitted in engines and gearboxes are generally lubricated by oil spray, splash, mist, drip
feed, or controlled level oil bath and loss of lubricant is prevented by the use of oil retaining
devices such as labyrinth seals, felt or rubber washers, and oil throwers.

Installation. The majority of bearing failures are caused by faulty installation, unsatisfactory
lubrication, or inadequate protection against entry of liquids, dirt or grit. To obtain the maximum life
from a bearing, great care must be exercised during installation and maintenance.
 Where bearings carry axial loads only, the ring need only be a push fit on the shaft or in the
housing, but bearings which carry radial loads must be installed with an interference fit between
the revolving ring and its housing or shaft, otherwise creep or spin may result. Before installation,
a bearing should be checked to ensure that it is free from damage and corrosion and that it
rotates freely. In some cases bearings are packed with storage grease which must be removed
by washing in a suitable solvent. All open bearings should be lubricated with the specified oil or
grease before installation.
 Bearings should be assembled correctly; i.e. as specified in the appropriate drawing or manual
and should be seated squarely against the shoulders on shafts or housings. Damage to the
shoulders, bearing rings or the presence of dirt could prevent the correct seating and thereby
impose stress on the bearings and promote rapid wear. Some bearings are supplied as matched
pairs and should be mounted correctly.
 Bearings may often be installed using finger pressure only, but where one ring is an interference
fit, an assembly tool or press should be used. It may also be necessary to freeze the shaft or
heat the bearing in hot oil, depending on the degree of interference specified. If a press is not
available, the use of a soft steel or brass tube drift may be permitted in some instances. Diagram
below ‘Fitting Interference Ring’. Any force must be applied only to the ring concerned, since
force applied to the companion ring may result in damage to the rolling elements or raceways. On
no account should a copper drift be used as work-hardening could result in chips or copper
entering the bearing.
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12.3
INSPECTION
Wear and corrosion of bearings, once started, progress rapidly and so bearings showing evidence of
these faults should be discarded. Frequent removal of bearings from shafts or housings may result in
damage to the bearings and for this reason a routine inspection is normally carried out in situ. Wheel
bearings are normally inspected when the wheel is removed. It may not be possible to examine the
rolling elements and raceways while the bearing is in position, but it is usually possible to examine the
rings externally for overheating, damage and corrosion, after removing the surplus grease with a lintfree cloth. The internal condition of a bearing may be revealed by an examination of the lubricant
exuding from the bearing. Metal particles reflect light and give a rough feeling when the lubricant is
rubbed into the palm of the hand. Bearing wear may be checked as follows:
1. Actuate the moving parts slowly for smoothness of operation. Roughness may result from grit in
the bearing elements or surface damage to the rolling elements or raceways, caused by corrosion,
excessive wear or lightning strikes.
2. Check for wear by moving the inner race or shaft in both axial and radial directions. The amount of
clearance will depend on the initial grade of fit.
3. Check shielded bearings to ensure that there is no rubbing contact between the stationary and
rotating components.
Cleaning
1. Wipe off grease, air pressure should be used to dislodge entrapped grease, but do not allow the
bearing to spin.
2. Soak or swill in White Spirit, oscillate and turn slowly under a jet of White Spirit, but do not allow
bearing to spin.
3. Dry with warm dry compressed air but, again, do not allow the bearing to spin. Lubricate with oil
and allow to slowly rotate.
Note. All of the previous points emphasise that the bearing should not be allowed to rotate unless it is
adequately lubricated.
12.3.1 GENERAL INSPECTION PROCEDURE & FAULTS

Rolling elements and raceways: Visually check for corrosion, pitting, fracture, chipping and
damaged cages where fitted. Also for ‘Brinelling’ i.e. indentation of the raceway by the rolling
elements. (Caused by shock loading).

Discoloration. Any discoloration of the bearing indicates that high (tempering) temperatures have
been reached. Generally no discoloration is permissible, but some wheel bearing manufacturers
allow Straw Yellow discoloration only. The maintenance manual should be consulted for
instructions.

Check the inner and outer rings, outer surface for signs of ‘creep’ – which will show as scuffed or
polished surfaces. If creep signs exists, measure affected dimensions, including the bearing’s hole
or shaft.

Measure axial and radial internal clearance with suitable equipment.

Running Smoothness. Checked by oiling and mounting on a shaft to rotate at 500 – 1000 r.p.m.,
holding the outer ring and feeling / listening for smooth running and resistance.

Sealed Bearings. Reject if any grease is leaking from the seals. It is now known how much grease
has been lost.
12.4
STORAGE
1. Soak the bearing in preservative oil (mineral and lanolin mix) as specified. Turn and swill around to
ensure all over coating.
2. Wrap in grease-proof paper or oil paper; box and label. Store horizontally to prevent brinelling.
3. Clean, inspect and re-protect annually.
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4. Miniature bearings are immersed in oil in a closed phial.
5. Storage conditions should be clean and dry with an even temperature to minimise condensation.
12.5
TRANSMISSION
12.5.1 KEYS AND KEYWAYS
Where considerable mechanical power has to be transmitted from a shaft to a hub or vice-versa; the
two components may be locked together and secured by means of one or more keys and keyways.
The key itself is a solid piece of metal of square or rectangular cross-section, of uniform width, and
uniform or tapering thickness. This key fits into a matching recess formed between the shaft and the
hub. This combination is generally used in circumstances that do not call for frequent removal of the
shaft from the hub.
The following are the most common types in use:

Taper Key. These are made with a standard taper of 1:100 on the thickness, the tapering face of
the key matching the taper of the recess or keyway. This type of key is designed to resist axial
movement between the hub and the shaft and once fitted, the key should remain undisturbed
except in emergency. The following taper keys are in used:
 Hollow Saddle. This type is curved to
suit the shaft radius; when driven into
position its taper provides a friction
grip between hub and shaft that is
capable of taking moderate loads. The
absence of any form of keyway on the
shaft if a feature of this type of key.
 Flat Saddle. This key is rectangular or square in cross section and bears on a flat formed on the
shaft. It provides a more positive grip between shaft and hub than is achieved by the hollow
saddle key.
 Plain Taper and Gib-headed. These fit into keyways which
are formed partly in the shaft and partly in the hub. The Gibhead illustrated has an angle chamfered on the end. They
are capable of transmitting much greater power than saddle
types.
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 Feather. This type of key is used where axial movement is
required between shaft and hub, for example, a feather key may
be used if it is necessary for a pulley or gear-wheel to move
along a shaft while it is still being driven. The hub keyway is cut
to allow side and top clearance around the key, so permitting a
sliding fit of key in the keyway.

Woodruff. This key is made in the form of a segment of a parallel-sided disc; it fits into a cavity in
the shaft which conforms closely to the rounded portion of the key and, into a uniform keyway in the
hub in such a manner as to provide a push fit on the sides and a clearance fit at the top of the key.
These keys may be fitted to parallel or tapered shafts.
12.5.2 SPLINED & SERRATED DRIVES
These are used as a method of keying together rotating components. Such drives will transmit high
power and allow easy assembly and dismantling. Some drives are ‘wasted’ to allow them to shear
when the component they are driving becomes damaged and would put too much load on the driving
device. Serrated drives may be used in preference to splined drives in positions where adjustment
would be necessary, such as windscreen wiper blades.
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
Spline Drive. A spline is an integral feather key projection
which is machined on a shaft or in a hub, and is usually of
uniform rectangular cross-section, the splines on the shaft
mate with the recessed in the hub, and the dimensions are
such as to allow a sliding fit. Splined shafts usually have at
least four splines on each member. Typical uses are in
hydraulic pump drives, generator drives, fuel pump drives.

Serrated Drive. Similar in principle to the splined drive, the
serrated drive makes use of triangular projections on the shaft
and hub which mate together as in the splined device to give a
sliding fit.
12.5.3 MASTER SPLINE
These are used when a hub must be assembled to a shaft in a specific position for the purposes of
timing or balance. The master spline will generally take the form of a wide spline mating with a wide
recess.
12.5.4 EXAMINATION
Keys, keyways, splined and serrated drives must be subjected to examination for damage and wear.

Damage. The most common cause of damage to keys, keyways, splined and serrated drives is by
mishandling and not using correct assembly / dismantling tools and methods. Most types of
damage (dents, burrs and cracks) will be clearly visible to the naked eye or with hand magnifiers.
However, in some instances more sophisticated Non-destructive testing methods (N.D.T.) may be
required.

Wear. All moving parts tend to wear and the extent of this can be found by careful visual
examination. In some circumstances precision instruments may be employed to check if the wear is
within the limits stipulated for the component. In these cases reference should be made to the
relevant maintenance manual.

Fitting. Never attempt to force components together; if they do not fit correctly, the reason must be
investigated and resolved.
12.5.5 CHAINS
Chain provides a flexible, strong and positive connection and is used to change the direction of pull in
control runs where considerable force is exerted – for example, in aileron, elevator, trimming tab and
engine controls. The change of direction of pull is achieved by a short length of chain which is routed
around the teeth of a sprocket and is connected, at each end, to the steel rod or cable of the control
system.
The chain used in aircraft remote control systems is of the simple roller type which is similar to the welltried bicycle chain. The chain consists of a series of inner and outer links which are assembled as
shown the diagram below ‘Standard Chain Fitting’. The size of a chain is measured by the pitch of its
links, this being the distance between the centres of the bearing pins. There are four sizes of chain
used on aircraft, 8mm, 9.5mm, 12.7mm and 15.9mm, of which the most commonly used is the 8mm.
The illustration also shows the various end fittings which are used to attached the short lengths of
chain to the rest of the control run.
12.5.5.1 Installing Chain Control
The following points should be observed when installing a length of chain.
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1. Ensure that the chain is the correct size and engages smoothly with the sprocket without any
tendency to ride on top of the teeth.
2. Look along the chain to ensure that it is not kinked or twisted.
3. Check the end fittings and connectors are security.
4. Examine the sprocket mountings for security and correct position.
5. Ensure that the chain is clean and free from corrosion. Lubricate as specified in the servicing
schedule.
The incorrect assembly of chains should be rendered impossible by the use of non-reversible chains in
conjunction with the appropriate types of wheels, guards and connectors.
12.5.5.2 Servicing
1. On each servicing occasion all lengths of chain used in remote control systems are examined for
serviceability and smoothness of travel. If found to be dirty, corroded, worn or damaged in anyway,
the chain is to be removed for cleaning and closer examination. After the chain has been removed:
2. Clean by soaking in a bath of Kerosene and brushing with a stiff fibre brush.
3. Examine for cut, bent or corroded links.
4. Test the links for freedom or movement. If a tight joint is discovered, the end of the bearing pin
should be tapped lightly with a hammer. If this does not free the joint, the chain is to be discarded
and a new one fitted.
5. Suspend the chain from one end and visually check for any distortion.
6. If the chain appears loose on the sprocket, or if there is any reason to suspect that it has been
stretched, the chain is to be tested for excessive elongation.
Chains used for aircraft purposes are generally of the simple roller type and comply with the
requirements of British Standard B.S.288 1934, entitled ‘Steel Roller Chains and Chain Wheels’. A
complete schedule of dimensions and breaking loads for chains is given in this Standard.
Chain assemblies are produced to standards prepared by the S.B.A.C., these standards providing a
range of chains built up in various combinations with standard fittings, e.g. end connectors with internal
or external threads, bi-planar blocks for changing the plane of articulation of a chain through 90º (see
diagram ‘Typical Chain Assembly Arrangements’ later in this section) and cable spools for connecting
chains to cables having eye-splices. Such fittings are illustrated in the previous diagram ‘Standard
Chain Fitting’.
12.5.5.3 Chain Assemblies
A simple roller chain consists of outer and inner plates, rollers, bearing pins and bushes; component
parts are shown in the diagram below ‘Chain Details’. The chain has three principal dimensions
(known as gearing dimensions since they are related to the size of the wheels on which the chains
run), these being pitch, width between inner plates, and roller diameter. The position at which these
dimensions are measured are also shown the diagram below.
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A typical assembly for 3/8 in. and 1/2 in. chains, using a standard end connector with an internal
thread, is shown in the diagram below.
The pitch of the chain is the distance between the centres of the rollers and for aircraft purposes, four
sizes are standardised by the S.B.A.C., as shown in the table below. B.S.288: 1934 prescribes that
the proof-load for a chain should be one-third of the minimum breaking load; the relevant figures for
simple chains are also given in the table below.
Chain
Pitch
B.S.No.
Minimum Breaking
Load
8 mm
1
800 lb.
0.375 in.
2
0.50 in.
4
0.50 in.
6
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1,900 lb.
1,800 lb.
3,500 lb.
Proof Load
267 lb.
634 lb.
600 lb.
1,166 lb.
Page 12-9
12.5.5.4 Testing a Chain for Elongation
If excessive elongation is suspected, tension the chain and whilst it is under load, measure the
distance between bearing pin centres over the maximum convenient length of chain. The percentage
elongation can then be obtained from the following equation:
A-B
B x 100 = percentage elongation
Where
A = the measured length
B = the nominal length i.e. the chain pitch times the number of pitches measured on the suspect
chain.
If the elongation is more than 2% the chain must be renewed.
After examination, the chain should be thoroughly soaked in oil before refitting to the aircraft or putting
into storage, in accordance with the servicing schedule. Chains which are put into storage should be
coiled.
Chains assemblies for aircraft systems should be obtained as complete, proof-loaded units from
approved chain assembly manufacturers and not attempt should be made to break and reassemble
riveted links or riveted attachments. If it is necessary to disconnect the chain, this should be
undertaken only at the bolted or screwed attachments. Split pins must not be used and this applies
also to nuts and bolts which have been peinde.
Note. The procedure specified by S.B.A.C. standards for securing nut and bolt joints for Class 1
application is to peen the bolt end for 8mm. pitch chain and to split pin the bolts of the remaining
standard chains. In all cases the nut is actually a lock nut, since the hole in the loose outer plate is
also tapped.
The use of cranked links for the attachment of the chain to end fittings, etc., is not permitted, thus,
when a chain is required to terminate in a similar manner at each end, the length should be an odd
number of pitches. For the same reason, an endless chain should have an even number of pitches.
The use of spring clip connecting links is prohibited and the attachment of chains to other parts of the
system should be effected by positive methods such as pre-riveted or bolted joints.
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The diagram below ‘Typical Chain Assembly Arrangements’ illustrates typical arrangements of chain
assemblies. (a) shows the simple transfer of straight-line to rotary motion, (b) illustrates how a change
of direction of straight-line motion is obtained, whilst (c) shows a change of direction of motion in two
planes by the use of a bi-planer block.
A range of non-interchangeable end fittings is available as a safe-guard against the crossing of
controls. However, these connectors do not always prevent the possibility of reversing the chain end to
end on its wheel, neither do they prevent the possibility of the chain being assembled to gear on the
wrong face where two wheels are operated by the same chain. Such contingencies can be overcome
by use of non-reversible chains.
12.5.5.5 Non-Reversible Chains
Non-reversible chains are similar to standard chains except that every second outer plate is extended
in one direction in order to break up the symmetry of the chain. The complete system of nonreversibility involves the use of five features, i.e. the non-reversible chain, the shroud on the wheel,
correct positioning of the wheel on its shaft, the chain guard and non-interchangeable connectors. The
shape of the special outlet plates and the principle of non-reversible chains is shown in the following
diagram.
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It will be seen from the above diagram that by providing a shroud on one side of the wheel and by
making use of the chain guard, the reversing of the chain end to end on its wheel is not possible. It
should be borne in mind that in practice a special feature, such as an attachment collar, a key or a flat
on the shaft in conjunction with a specially shaped hole, is incorporated in the wheel mounting to
ensure that it can be assembled on its shaft in one definite position only.
The diagram below ‘Non-Reversible Chain with Jockey Pulley’ illustrates an instance where the use of
jockeys is necessary or where contra-rotation of the wheel is required; it will be seen that the feature
on non-reversibility does not affect the ability of the chain to gear on both sides.
12.5.5.6 Inspection After Assembly
After installation in the aircraft, the chain should be examined for freedom from twist, particularly in
instances where the attachment is made to rods by means of screwed end connectors, or where a twist
may inadvertently be applied to the chain during the locking of the assembly. Care should be also
taken to ensure that the chain is not pulled out of line by the chain wheel; the chain should engage
smoothly and evenly with the wheel teeth and there should be no tendency for the chain to ride up the
teeth.
The pre-tensioning of chains should not be excessive, as this will cause friction, but should be just
sufficient to prevent any back-lash in the system.
The guarding should be checked to ensure that jamming could not occur and that the chain would not
come off the wheel, should it become slack.
The security end connections should be checked, care being taken to ensure that the split pins in the
chain connecting bolts are correctly locked.
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The initial lubricant on new chains should not be removed and the chains should be further lubricated
after assembly by brushing all over, particularly on link edges, with lubricant complying with
specifications DTD 417A, unless otherwise specified.
The wheel or pulley mountings should be examined to ensure that the wheels or pulleys are firmly
secured to the shafts or spindles, that they are correctly located and are running freely.
12.5.5.7 Maintenance Inspection
Chain assemblies should be inspected for serviceability at the periods specified in the relevant
Maintenance Schedule; guidance on the recommended methods of checking chains is given in the
following paragraphs.
The continued smoothness of operation between the chain and the chain wheel or pulley should be
checked. If the chain does not pass freely round the wheel or pulley, it should be removed and
checked as detailed.
The chain should be checked for wear; if it is worn so that the links are loose and can be lifted away
from the wheel teeth, it should be removed and checked for excessive elongation as detailed.
The chain should be checked for damage, cleanliness, adequacy of lubrication and freedom from
corrosion. If the inspection shows the chain to be corroded or otherwise defective, it should be
removed.
In instances where it becomes necessary to adjust the tension of the chain in system incorporating
turnbuckles or screwed end connectors, care should be taken to ensure that the chain itself is not
twisted during the adjustment. The connectors should be held firmly while the locknuts are being
slackened or tightened.
12.5.5.8 Inspection of Chain Assemblies Removed at Overhaul Periods
When it is necessary to disconnect the chains the assemblies must be removed at design breakdown
points.
12.5.5.9 Check Elongation
If elongation through wear is suspected, the following procedure should be adopted.
1. The chains should be cleaned by immersion in clean paraffin and brushing with a stiff brush; after
cleaning, the chains should be dried immediately by hot air to ensure that no paraffin remains,
otherwise the chains will corrode. The chains should be measured when clean but before any oil is
applied.
2. The chains should be placed on a flat surface and stretched by the application of a tensile load.
The table below indicates the load applicable to the various sizes of chains. The length should
then be measured between the centres of the bearing pins, elongation being calculated by the
formula below.
Chain
Pitch
8 mm
B.S.No.
1
Tensile Load
(lb.)
12
0.375 in.
2
16
0.50 in.
4
28
0.50 in.
6
28
The percentage extension over the nominal length should be calculated by the following formula:
Percentage extension =
M - (X x P)
x 100
XxP
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Where
M = Measured length under load in inches.
X = Number of pitches measured.
P = Pitch of chain in inches.
It the extension is in excess of 2 per cent on any section of the chain the whole chain should be
replaced. Should localised wear by likely to occur in a chain run, additional checks should be made on
such sections and the percentage extension ascertained from the formula given. If the extension in
such sections is in excess of 2 per cent, the chain should be rejected.
The chain should be checked for tight joints by articulating each link through approximately 180º, the
most suitable method being to draw the chain over a finger. Tight joints may be caused by foreign
matter on the bearing pins or between the inner and outer plates; this may be remedied by cleaning as
described. If cleaning is not successful, the end of the bearing pin may be very gently tapped with a
light hammer, but if this does not clear the joint, the chain should be rejected. Tightness may also be
caused through lack of clearance between the inner and outer plates due to damage; if this is so, the
chain should be rejected.
The chain should be examined for damage, cracks and wear to plates and rollers and for evidence of
corrosion and pitting.
Note. It is nor permissible to break down or attempt to tighten a riveted link in a run of chain.
12.5.5.10 Proof Loading
It is not necessary to proof load a chain after removal for routine examination. However, if it is desired
to replace a portion only of the assembly, proof loading of the complete assembly is necessary. The
proof load should be evenly applied and unless this can be assured, it is considered preferable to fit a
complete new assembly.
12.5.5.11 Protection & Storage
After the chain has been cleaned, inspected and found acceptable, it should be thoroughly soaked in
an appropriate oil, time being allowed for the lubricant to penetrate to the bearing surfaces. If not
required for immediate use, the chain should be laid on a flat surface, carefully coiled and wrapped in
grease-proof paper, care being taken to ensure the exclusion of dirt and the prevention of distortion,
during storage.
12.5.5.12 Chain Wheels & Pulleys
During installation, chain wheels and pulleys should be checked to ensure that they are attached in the
manner and by the method specified by the relevant drawings. The correct positioning of chain wheels
is of particular importance when non-reversible chains are used. During maintenance, chain wheels
should be checked for security and wear on teeth. Pulleys should be checked for damage and
excessive wear on the walls and on the chain guide section. The continued efficiency of ball races
should also be ascertained (Leaflet BL/5-2).
12.5.6 GEAR & GEAR TRAINS – TYPES & USES
The term gears or gearing is applied to a system of moving parts in the form of toothed wheels which
are used to transmit motion. Gears may be referred to as:

Driving Gears
-
attached to the ‘power in’ shaft.

Driven Gears
-
attached to the ‘power out’ shaft.

Idler Gears
interposed between the driving and driven gear in order to maintain the
direction of rotation of the output shaft the same as the input shaft.
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The basic principle involved is essentially one of leverage in that it is easier to lift a heavy weight if a
lever used as the lever multiplies the effort applied to it.
In the example shown above, the diagram of wheel ‘B’ is twice that of wheel ‘A’. The thick lines running
from the centres of the wheels are imaginary levers and are equal in length to the radius of the wheel.
If wheel ‘A’ is turned by it’s shaft, the short lever of wheel ‘A’ will bear against the lever of wheel ‘B’.
Because this lever is twice as long as the lever of wheel ‘A’, the torque resulting in the shaft of wheel
‘B’ will be twice that applied to the shaft of wheel ‘A’.
Since gear ‘B’ is twice the size of gear ‘A’, it has twice the number of teeth, each of which acts as a
lever. As gear ‘A’ turns gear ‘B’ rotates in the opposite direction, but because of the difference in the
number of the teeth gear ‘B’ only rotates haft a turn for every complete turn of gear ‘A’, or half the
speed. The relationship between the number of teeth which mesh with one and other, is known as the
‘Gear Ratio’ and may be calculated for any gear train using the simple formula:
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1.
Number of teeth on Driving Gear
= Gear Ratio
Number of teeth on Driven Gear
2. Speed of output shaft = Speed of input shaft x Gear Ratio.
Gears therefore can be used to perform two main functions:
1. To multiply the torque of the driving shaft and to decrease the speed of rotation of the driven shaft
or visa versa.
2. To reverse the direction of the drive or to alter the direction of the driven shaft.
12.5.6.1 Types of Gears
See diagram on the following page.

Sour Gears. Common straight toothed gear wheels with teeth formed externally or internally.
External spur gears are used when a change of speed is required and the shafts lie parallel to each
other, internally toothed gears are used when a change of speed is required whilst maintaining an
overall minimum diameter.

Helical Gears. Teeth are cut on a helix and a sliding engagement is made, with more than one
tooth in mesh at any one time.
This tooth shape is smoother and quieter running than the spur type, but produces a heavy axial
loading on the shafts. This axial loading is proportional to the resistance to motion offered by the
driven gear and can be eliminated with gears that are in permanent engagement by using ‘Double
Helical’ gearing, where the teeth are cut with opposite helix.

Bevel Gears. These are used when the drive is required to be transmitted at an angle.
The teeth are formed on conical wheels and may be cut straight across in a ‘Straight Bevel’, or in a
helix in the ‘Spiral Bevel’.

Hypoid Gears. These are used when the axes of the two shafts do not intersect and are similar in
appearance to spiral bevel gears.

Worm Gears. These are often used when a large reduction in shaft speed is required and a high
resistance to turning is encountered. The worm pinion teeth are similar to a multi-start thread while
the worm wheel teeth are cut at an angle.

Skew Gears. These are used to connect shaft’s whose centre lines do not intersect or run parallel.
The teeth are cut on an acute helix.
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Associated Terms

Backlash. This term is used to describe the clearance which must exist between gear teeth at
point of mesh, essential with all forms of gearing to allow for expansion and lubrication.

Idler Gear. A gear which is interposed between the driving and driven gear, its function is to
connect the drive between two shafts. A spur idler gear is used between two parallel shafts to
maintain the direction of rotation and does not affect the ratio of the gears. A bevel idler may be
used where two shafts intersect and / or are co-axial.

Intermediate Gear. A gear which is positioned between the driving gear and one or more driven
gears in a gear train. It may function as an idler gear or transmit drive through its own shaft.

Compound Gear. This is a gear wheel which has more than one driving face. These faces may
be formed integrally on one casting or forging, or it may comprise two or more gears bolted or
splined together to transmit drive to a number or shafts.

Pinion. This term is usually applied to the smaller of two mating gears.

Layshaft. A shaft which supports an idler gear or intermediate gear, it may be integral with the
gear and be supported by bearings, or may be fixed and provide a bearing surface for the rotating
gear.

Rack and Pinion. A device in which a toothed rod (rack) meshs with a mating pinion to translate
the rotary movement of the pinion into linear movement.

Step-Up Drive. A drive through a gear train in which the speed of rotation of the output (driven)
shaft is increased.
Example: Used in aero-engines in a generator drive. It ensure that the generator has sufficient
rev/min to remain ‘on charge’ at engine idling rev/min.

Step-Down Gear. A reduction gear in which the rev/min of the output shaft is reduced while the
torque is increased.
Example: Used between the engine and propeller in order to allow the engine to develop its power
by running RPM while maintaining high propeller efficiency by avoiding the tips speeds reaching
Mach I.
12.5.6.2 Propeller Reduction Gears
The principle considerations which determine the choice of reduction gear for a specific application are:
1. The reduction ratio possible within a certain overall size.
2. The relationship between the input and output shaft axis.
Propellers driven by gas turbine engines require vary large reduction gear ratio’s to cater for the needs
of both the engine and the propeller. For example, the Dart engine develops its maximum power at
15000 RPM, and to avoid compressibility problems at the propeller tips the propeller must be limited to
approximately 1370 RPM. Thus a reduction gear ratio of about 11:1 is required.
A simple spur gear capable of such a reduction would be excessively bulky and also the propeller and
turbine shafts would lie on different axes which would cause problems in engine air intake design.
Both of these problems are overcome in current turbo-prop designs by the use of various forms of
‘Epicyclic’ Reduction Gears.
12.5.6.3 Simple Spur ‘Epicyclic’
A train consisting of a sun (driving) gear meshing with and driving three or more equip-spaced gears
known as ‘Planet Pinion’. These pinions are mounted on a carrier and rotate independently on there
own axles. Surrounding the gear train is an internally toothed ‘Annulus Gear’ in mesh with the Planet
Pinions.
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If the annulus is fixed as in the diagram below, rotation of the sun wheel causes the planet pinion to
rotate about their axes within the annulus gear, this causes the planet carrier to rotate in the same
direction as sun wheel but at a lower speed. With the propeller shaft secured to the planet pinion
carrier, a speed reduction is obtained with the turbine shaft (input shaft) and propeller shaft (output
shaft) in the same axis and rotating in the same direction.
If the annulus is free as in the diagram below, rotation of the sun wheel causes the planet pinions to
rotate about their axles within the annulus gear. With the planet pinion carrier fixed and the propeller
shaft attached to the annulus gear, rotation of the planet pinions causes the annulus gear and propeller
to rotate in the opposite direction to the sun wheel and at a reduced speed.
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12.5.6.4 Compound Spur Epicyclic
Compound epicyclic reduction gears enables a greater reduction in speed to be obtained without
resorting to larger component. They may be of either the fixed or free annulus type.
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13. CONTROL CABLES
Cables provide a strong, light and flexible method of control and are used extensively in aircraft control
systems. Cables operate in tension and can, therefore, only be used to pull the control. However, two
cables can be arranged in the form of a continuous loop to provide a pull in both direction (see diagram
below).
13.1
TYPES
Flying control cables are normally pre-formed; that is, the stands in the cable are formed into the shape
they will assume in the complete cable. The cables, which are made of galvanised or corrosionresistant steel, are impregnated with a friction-preventative lubricant during manufacture.
A cable is made up of steel wires which, in
turn are formed into strands, as illustrated
in the two examples below. Each strand
consists of several wires (7 or 19) which
are wound helically in one or more layers,
the centre wire being known as the core
wire or king wire. Each cable is made up
of several strands (usually 7), wound
helically around the centre or core strand.
The cable is described by the number of
strands it contains and by the number of
individual wires in each strand. Figure ‘A’
shows that a 7 x 7 cable consists of 7
strands, each having 7 wires; Figure ‘B’
shows a 7 x 19 cable – 7 strands, each
having 19 wires.
The number of wires in each strand, the number of strands and the overall diameter of the cable
determine the breaking load of the cable. For example, a 7 x 19 cable of 6.4 mm (¼ inch) overall
diameter has a minimum breaking load of 7000 lbf. Cables are classified either by the minimum
breaking load, which may be quoted in cwtf, lbf or kN, or by the nominal diameter in inches.
It is often necessary to coil a cable when handling it for assembly into an aircraft. The coil should be of
large diameter; never less than 50 diameters of the cable involved and with a minimum diameter of 150
mm. To avoid kinking the cable and thus making it unserviceable, uncoiling should be done by rotating
the coil so that the cable is paid in a straight line.
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13.2
FITTINGS
Each flying control cable has metallic end fittings (see diagram
below) swaged in position. Swaging is an operation in which a
metallic end fitting is secured to the end of a cable by plastic
deformation of the hollow shank of the fitting. The cable end is
inserted into the hollow shank and the shank is then squeezed in a
swaging machine so that it grips the cable. Only specialist units are
permitted to do this work. Normally a faulty cable should be
replaced by a serviceable one on an exchange basis.
13.3
PULLEYS & BELL CRANKS
13.3.1 PULLEYS
Pulleys are used to change the direction of operation of flying control cables and to give support on
long straight runs. A cable guide (or retainer) is fitted to the pulley to ensure that the cable remains on
the pulley. A typical pulley, with its retainer is illustrated in the diagram below. When adjusting a
control, it is important to ensure that the cable fittings do not foul the pulley; otherwise the cable
movement will be restricted. Also look for possible misalignment between the cable and pulley: This
must not exceed 2º (see ‘B’ in diagram below).
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13.3.2 FAIRLEADS
A fairlead restricts the sideways fluctuations of the cable caused by aircraft vibrations. This damping
effect of the fairlead is particularly important in helicopter control systems where the vibration frequency
is so low that it can set up ‘resonance’ conditions in the cable (i.e. the conditions are such that the
fluctuations progressively build up to impose very heavy loads on the cable). A typical fairlead is
illustrated in the diagram below. It is manufactured from a low friction material, such as fibre, tufnol or
nylon.
13.3.3 SCREWJACK
A cable-operated trimming tab control system operates a screwjack at the receiving end of the system.
The screwjack (see diagram below) is attached by means of an adjustable rod to the trimming tab. The
cable movement rotates the sprocket of the screw jack to reposition the trimming tab. This unit acts as
a lock, retaining the trimming tab in the desired position until the cockpit control is next moved.
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13.3.4 CABLE TENSIONING
Need for tension. For a wire cable control system to operate effectively and efficiently, the cable
tension must be correct. It should be just sufficient to operate the control – neither too taut nor too
slack; tension imposes an unnecessary load on the control system, tensioned to a pre-determined
value in accordance with the servicing instructions for the particular system. The value chosen is such
that sufficient tension is maintained over a range of operating temperatures. The range of
temperatures over which the tension remain satisfactory depends upon whether or not a cable tension
regulator is fitted in the system. Temperature change, cable stretch and general wear of supporting
parts affect the tension which must, therefore be checked and adjusted as necessary at specified
intervals. Some cable systems have compensating devices fitted which ensure effective operation over
a much wider range of temperatures than would otherwise be possible.
Turnbuckles. It is normal to use turnbuckles to adjust the tension of cables in flying control systems.
There are two types of turnbuckles in common use, (see diagram below). The type fitted will match the
end fitting on the cables.
When connecting the cables together using a turnbuckle the threads must be evenly engaged at either
end. It is also important to ensure that sufficient threads are engaged; otherwise the load on the cable
could strip the threads.
With the barrel type of turnbuckle (‘A’ in the above diagram), no threads should be visible.
Cable end fittings that engage with the tension rod type of turnbuckle (‘B’ in the above diagram),
have small ‘witness’ holes drilled in their shank. The turnbuckle thread must reach these holes
for the connections to be in safety.
All turnbuckles are locked in the approved manner using wires or clips as shown above.
Adjusting the tension in cable system. In an aircraft, there are many different types of metal, each
one of which expands with increasing temperatures at a different rate. The effect of this in a cable
system is that the tension tends to decrease with an increase in altitude. Thus, to retain sufficient
tension at altitude, the pre-determined load must be high. This required a strong structure, with a
resulting increase in weight. Furthermore, compared with a tension regulated system, stress and static
friction are also higher. While tensioning is being carried out to the correct value of pre-determined
load by evenly adjusting all the turnbuckles in the system, the correct relative positions of the pilot’s
control and the relevant control surface must be maintained. The cable tension is checked frequently
using a tensioner as the adjustments are made. Another type of cable operated flying control system
has a cable tension regulator fitted in the system.
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13.3.5 CABLE TENSION REGULATOR
A cable tension regulator is a mechanical device which, when fitted in a cable system, allows the
cables under all conditions of temperature change and structural deflection to take up and let out
equally on each side of the circuit, thus maintaining uniform tension. The compensating unit of a
tension regulator may be manufactured with one or two springs; a double spring unit is described and
illustrated below.
This type of regulator consists of a pair spring-loaded quadrants, with a pointer and scale to record
tension compensation. The control cables are fitted to the grooved cable quadrants shown in the
above diagram. The purpose of the regulator is to maintain the cables at their optimum tension, by
compensating for small changes in cable length and the variation in the size of the airframe structure
that occur with changes in temperature. A cable tension setting graph is used (also see diagram
above) when it is necessary to set the cable tension.
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The regulator operation is illustrated in the diagram below.
The diagram to the left shows the compensatory reaction of the
regulator to a fall in cable tension. The cable tension is
maintained in the control system by the pressure exerted by the
compression springs on the cable quadrants, to which the
cables are fitted. When a fall in cable tension occurs, the
quadrants are displaced radically by the compression springs.
Movement of the quadrants ceases when the pre-set tension
has been restored. Links connecting the quadrants to the
crosshead cause the crosshead to move freely outwards on the
locking shaft during quadrant displacement. This allows the
crosshead to take up a new position on the locking shaft when
the cables have reached their pre-set tension.
Figure ‘B’ shows the action of the cable regulator following an
increase in able tension. In this case, the regulator operates in
the reverse direction to that shown above. The radial movement
of the quadrants moves the crosshead via the links, along the
locking shaft until the pre-set cable tension has been reestablished.
Figure ‘C’ shows that when the pilot operates a regulated
control, the crosshead tilts on its locking shaft, causing it to lock
on to the shaft. Both quadrants are now locked together and
operate as a lever to give positive control of the system.
Adjustment of the regulator is carried out by using a cable tension setting graph, as shown in the
relevant aircraft publication. To adjust the cable tension, the following procedure should be adopted:
1. Determine the ambient temperature as described in the aircraft maintenance manual.
2. Refer to the setting graph and identify the scale position that corresponds to the ambient
temperature.
3. Adjust the cable tension equally on either side of the regulator by means of turnbuckles, until the
regulator pointer aligns with the previously identified scale position.
After the cables have been set to the correct tension, regulator compensation may be checked by
grasping both cables near their point of entry to the regulator and forcing both cables in towards each
other. The resulting movement of the quadrants should be smooth and even. If the regulator fails to
move or the movement is jumpy, it may indicate that the cables have been wrongly rigged so that the
tension is uneven, causing the crosshead to tilt and ‘lock’ the system.
The introduction of cable regulators into a control system ensures a relatively constant cable tension.
Hence, a lower cable tension can be used, which lowers static friction, improves control response and
permits a reduction in structural weight.
13.4
RIGID CONTROL CABLES
Many of the control knobs and leavers situated in an aircraft cockpit are designed to operate systems
and controls in different parts of the aircraft. These controls, or system, are known as ‘Remote Control
Systems’. Many different services are operated in this way, including those fitted to modern aircraft.
They include canopy jettison mechanisms, air conditioning and heating controls, aircraft parking brake
mechanisms and engine throttle and fuel controls, to name but a few. In each type of control system,
the end from which it is operated is called the transmitting end and the other end of the control that is
remote from the operating end, is called the receiving end.
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Many remote control systems are now operated electrically from the aircraft cockpit. However, there
are also many systems fitted to modern aircraft that are entirely mechanical in their operation and
employs cables, chains and rods in their make-up.
13.4.1 BOWDEN CONTROLS
13.4.1.1 Principle of Operation
The Bowden Control System is based upon the operation of a multi-stranded wire cable housed in a
flexible conduit (see diagram below). The system is usually designed to operate lightly loaded
components in a one way direction, by the application of a pulling action on the wire from a control or
operating lever. Components or services operated in this way are returned to their original positions by
the force exerted on the wire by a return spring. Bowden controls can also be designed to operate in
two directions; in this case two wires and a pulley are used to transmit the two-way operating force.
13.4.1.2 Cable
A Bowden cable is manufactured from a number of stainless steel wires that are helically wound
around each other to form a cable. There are two types of Bowden cable assembly; small lightly
loaded controls employ cable manufactured from seven wires, whilst more heavily loaded controls
employ cable manufactured from nineteen wires.
13.4.1.3 Conduit
The Bowden conduit consists of a close-coiled wire covered cotton braiding and finished externally with
a black waterproof coating. A cap is fitted to each end of the conduit to prevent the braiding from
becoming unravelled, and to reinforce the ends of the conduit. The length of a Bowden control is
usually restricted to 1.8 – 2.5 metres (6 –8 ft). However, should a longer control tun be required, the
conduit for the long straight parts to the run must be manufactured from rigid metal tubing, whilst
flexible conduit is used where bends in the control occur.
13.4.1.4 Cable Nipples
A nipple is fitted to each end of the cable. The cable and nipple assembly transmits the control
operating force from the control lever to the service or component to be operated. There are three
types of nipple in use (see diagram below), the choice of nipple being dependent upon the design of
the fitting at the transmitting or receiving end of the cable, to which the nipples are attached. The
nipples may either be swaged or soldered to the cables; cable assemblies made up by manufacturers
usually have swaged nipples fitted to them.
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13.4.1.5 Bowden Control System Components
There are relatively few component parts to a Bowden Control System. Each system is, however,
designed to meet a particular requirement and hence may differ in detail from the basic control, whilst
retaining the functional characteristics of the basic system. In the following paragraphs the basic
components of the system will be described and illustrated.

Hand lever. The hand lever (see diagram below) is mainly found in the aircraft’s cockpit and is
used to initiate system operation. The nipple at one end of the operating cable is located in a
recess in the lever, and is retained in position by a face plate through which a centre screw is
passed to clamp the plate into position.

Plain adjustment stop. An adjustment stop may be fitted to the control at the receiving end of the
cable and in some cases, a stop may also be fitted to the transmitting end of the control. The
purpose of the stop is to provide a means of adjusting the length of the control conduit, and to alter
the slackness in the operating cable. The stop consists of a special screw and nut assembly. It is
fitted to either the component to be operated, or to the aircraft structure adjacent to the component.
The screw is bored axially through its centre to permit the cable to pass through it and its hexagon
head is counter-bored to provide a seating for the metal cap fitted to the end of the conduit. An
adjustment stop similar to the type described is shown in the diagram below.
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
Double-ended stop. In some cases, it is not possible to fit a plain adjustment stop to a Bowden
control, due to the inaccessibility of the component to be operated. In such cases, a double-ended
stop is used (see diagram below). The double-ended stop is inserted into the control conduit at a
suitable position in the aircraft; this allows adjustments to be made to the control conduit. The
adjuster is made up of two main parts, one of which is screwed inside the other, and they are
locked together by a locknut. A hole is bored axially through the centre of the adjuster to allow the
cable to be passed through it and is counter-bored at either end to form a location for the conduit.
A safety hole is bored radially in the outer part of the adjuster and is used to check the minimum
thread engagement of the adjuster. The male, or inner part of the thread must always be visible
through the hole when the minimum thread engagement has been achieved.

Connectors. A connector is used when it is necessary to uncouple a Bowden control at some
point in the cable, or when a Bowden cable is used in conjunction with a different type of cable.
There are two types of connector in use, as illustrated below. In part ‘A’ of the illustration, a
connector is shown that may be used to join two cables together when they are not running through
a conduit; alternatively, it may be used to join a Bowden cable to a cable of a different type. Part
‘B’ of the illustration shows a connector that is used to join two Bowden controls and their
associated conduits together.
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
Junction boxes. A junction box is used to connect a single cable to two or more cables, enabling
a number of components to be operated by a single control. This principle may also be used in the
reverse order, in which two or more transmitting ends are used to operate a single control or
component. An example of the application of this principle is shown in the diagram below. In this
case, the Bowden control is connected to the Miniature Detonating Cord (MDC) system fitted to an
aircraft canopy; the system is used in an emergency to shatter the canopy transparency to allow
the aircraft pilot to escape from the aircraft. At the transmitting end of the control, the cables are
connected to three operating handles; one is situated inside the cockpit for operation by the aircraft
pilot and the other two are located on the fuselage sides for operation by rescue personnel. The
operation of any one of the handles will pull the appropriate control cable and through the junction
box, operate the MDC firing unit.
13.4.1.6 The Maintenance of Bowden Controls
The maintenance of a Bowden Control System may be divided into three separate phases, which are:
Servicing the system.
Removal of a cable.
Preparation and fitment of a replacement cable.

Servicing of the system. The extent of the servicing required on a Bowden control system is
largely dependent upon the frequency of its use and its location in the aircraft. Servicing
requirements are defined in the aircraft servicing schedule and may include the following
operations:
 Examine the ends of the cable for sign of wear, damage and fraying.
 Examine the conduit for kinks and signs of wear, particularly at the ends of the cable.
 If there is any slackness in the cable, adjust the adjustment stop to remove the slackness.
 Ensure that all attachments are secure and that the adjuster locknuts are tight.
 Operate the control and ensure that it operates correctly and is returned by the return spring
tension.
 Lubricate the control with the specified lubricant.

Removal of a cable. When a Bowden cable has been found to be unserviceable and is to be
removed from an aircraft, a sequence of operations is usually performed. Many of the tacks are
common to all types of aircraft and may include the following.
 Slacken the adjuster locknut and screw the adjuster in fully to slacken the cable to assist cable
removal.
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 Remove any cable clips or ties securing the cable conduit to the aircraft.
 Lift the cable nipple from the aircraft component.
 Dismantle the operating lever as required and lift the cable nipple from the lever recess.
 Carefully remove the cable from the aircraft and ensure that the cable does not become
entangled in the other controls and fittings.

Preparation and fitment of a cable. Bowden cable is usually supplied in the form of a complete
cable assembly and is comprised of a conduit, cable adjuster and an appropriate length of cable
with cable nipples fitted to either end. In some cases however, the cable and the conduit are
supplied separately in coils, and hence the cable must be correctly assembled with its component
part before it is fitted to an aircraft. When Bowden cable is supplied in this way, each end of the
cable is lightly tinned with solder for about 2.5 cm (1”) of its length; this prevents the wire from
becoming unravelled. It is then wound into large neat coils. Great care is needed when handling
Bowden cable to avoid distortion of the cable occurring. A distorted or kinked cable is very difficult
to straighten and can easily increase the operating load of a control, or prevent it from operating at
all if the damaged portion were to pass through a cable conduit. When work has to be performed
on a control, the work area must be kept clean and free from dirt and swarf; a contaminated control
can easily become stiff in operation and may ultimately seize altogether. Particular attention must
be paid to the correct axial alignment of a cable between the adjustment stop and the component to
be operated.
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A maligned cable will rub against the bore of the cable stop and cause excessive wear of the cable
wires to occur (see diagram above). When it is necessary to connect a cable to an operating lever,
the centreline of the conduit must be in a straight line with the mid potion of the rise and fall of the
arc of travel of the level.
When a replacement cable has been obtained and prepared for fitment to an aircraft, it should then
be fitted as follows:
 Clean and lubricate the cable.
 Install the cable into the aircraft and clip the conduit into position.
 Connect the cable nipples to the operating lever and to the component.
 Adjust the adjustment stops to remove all slackness from the cable and tighten the adjuster
locknuts.
 Operate the control to ensure that full and free movement is being achieved. Ensure that the
return spring re-sets the control to the OFF position, when the operating lever is set to OFF.
Periodically, it may become necessary to replace a Bowden cable assembly with a locally
manufactured item. The following notes describe a typical sequence of operations that would be
required under such circumstances.
 Remove the defective cable from the aircraft.
 Measure the exact length of the old cable conduit and cut a replacement from a new coil.
 Clean up the cut ends of the conduit and fit end caps.
 Measure the length of cable required and add about 5 cm (2”) to this dimension to aid assembly
of the cable. Measure out a length of new cable and lightly tin the cable along about 2.5 cm of its
length either side of the cutting point. Cut the cable to the required length.
 Ensure that one end of the cable is free from grease.
 Tin the cable with grace C solder over a cable length of at least 1½ times the length of the nipple.
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 Apply flux over the tinned portion of the cable.
 Heat the nipple with a hot soldering iron and then slide the nipple over the tinned cable, leaving
about 1.6 mm (1/16”) of cable protruding through the nipple. Sufficient heat must be applied to
the joint to ensure that the solder runs freely in it.
 Hold the soldered nipple in a vice and carefully spread the protruding cable wires around the preformed nipple cup (see diagram below) with a small ball pein hammer.
 Apply sufficient solder to the end of the cable to completely cover the wire strands and to fill the
nipple cup.
 Thoroughly wash the cable with hot water to remove any flux residue, dry and then lubricate the
cable with the lubricant specified in the servicing schedule.
 Thread an adjustable stop over the cable (if a stop if to be used at both ends of the cable) and
then slide the conduit over the cable, making sure that the protective caps remain in position at
each end of the conduit.
 Thread on the second adjustable stop if this is to be used. Screw in both stops to their minimum
length.
 Fix the cable assembly temporarily in position in the aircraft following the route that it is intended
to follow.
 Temporarily connect the previously soldered nipple to the operating lever and ensure that the
lever is selected to the OFF position.
 Ensure that the component, or service to be operated is in the OFF position.
 Thread the second nipple onto the cable, pull the cable taut and mark the correct position of the
nipple on the cable with a chinagraph pencil.
 Remove the control from the aircraft, degrease, tin and cut cable to length. Thread the second
nipple into position on the cable and solder using the previously described method.
 Install the cable in the aircraft.
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13.4.2 TELEFLEX CONTROLS
The Teleflex system of controls is used on many modern aircraft. Each system is controlled from the
aircraft cockpit and is used to operate such service as throttle, propeller and fuel cock controls. The
Teleflex control system has also been used in various other applications, such as canopy winding
mechanisms and aircraft trim controls. Unlike the Bowden control system, the Teleflex control provides
doe two-way operation of any component or system without the need for additional components to be
fitted to the system. It also provides a more accurate system of control than the Bowden system, a
feature which may be illustrated by its use for the operation of engine throttle controls. Apart from
system accuracy, the other main advantages of the Teleflex control system over the Bowden control
system are that it may be set in any position desired by the aircraft pilot, and may be locked in that
position if a locking device if fitted to the control.
The Teleflex control system consists of a special cable housed in a rigid metal conduit, or in certain
installations a flexible conduit may be used. Each control is made up of a number of special Teleflex
component parts, some of which may be modified to suit the particular function of the control. In some
systems, particularly those with very long control runs, ordinary control cables may be supplemented
with tie rods or chains for part of the control run. Such systems usually feature a Teleflex control at the
beginning, or transmitting end of the control run and also at the receiving end of the run.
13.4.2.1 The Teleflex Control Run
There are many variations to the basic Teleflex control run. Each control run is designed to impart
either a push-pull or a rotary motion to a system or component. In the diagram below, several
examples of Teleflex control runs are illustrated in their basic form; these show some of the
installations which may be fitted to a modern aircraft.
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13.4.2.2 The Teleflex Cable
The Teleflex control cable is unique in its construction, basically it consists of a steel core-wire
surrounding by a number of helically wound outer wires. Three types of cable are available for aircraft
use. The cables are significantly different in their construction and hence are not interchangeable with
each other. The three types of cable are:
DS 23/2 (also known as No 2).
DS 380.
DS 169330.

DS 23/2. The DS 23/2 type of cable is built up from a high tensile steel wire inner core that has
been wound with a close-pitched compression wire (see diagram below ‘Teleflex Cables’). A lefthand helix wire, interspaced by a spacer wire, is then wound around the compression wire/core
wire assembly. The complete cable assembly has a minimum breaking load of 204 Kg
(450 lb.).

DS 380. The DS 380 type of cable, although outwardly similar to the DS 23/2 cable in its
appearance, differs in its construction. The cable is not fitted with a compression wire, but it is
fitted with a high tensile steel inner wire; this has a greater diameter than the core wire fitted to the
DS 23/2 cable. A right-hand helix wire, interspaced by a spacer wire, is wound directly onto the
inner wire to form the complete cable assembly. This type of construction improves the efficiency
of the system with a reduction in backlash, particularly on the compression stroke, it also provides
an increased minimum breaking load strength of 454 Kg (1000 lb.).

DS 169330. The DS 169330 cable is similar in its construction and strength to the DS 23/2 cable,
with the exception that the spacer wire has been omitted from its design. The cable, which is
manufactured from Stainless Steel, is designed for use in the ‘hot’ areas of an aircraft, where
temperatures of up to 350ºC may be recorded.
13.4.2.3 Conduit
The Teleflex conduit is a metal tube whose function if to guide the cable through the aircraft, and to
enable compressive loads to be applied to the cable without it becoming kinked or distorted. Two basic
types of conduit are supplied, these are:
Rigid conduit.
Flexible conduit.
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
Rigid conduit. The rigid conduit is used for most of the Teleflex control systems. It may be
manufactured from Tungum, Steel or Light Alloy tubes. Most modern conduit is manufactured from
Aluminium and is lined with Polytetrafluoroethylene (PTFE) to reduce the friction generated by the
conduit. The addition of a lining, however, produced a greater outside diameter than that of the
standard conduit. Furthermore, PTFE lined conduit may not be used in the ‘hot’ zones of an
engine, (i.e. where temperatures exceed 100ºC). Teleflex conduit located in these areas must be
manufactured from either Tungum or steel tube.

Flexible conduit. Flexible conduit is used in Teleflex control installations to allow for the relative
movement of components whilst they are being operated. The conduit (see diagram below) is
usually kept as short as possible and is interposed between rigid conduit and the component to be
operated. The conduit consists of a continuous winding of metal strip, covered by a layer of cotton
interposed by fine wires running lengthways along the conduit. Finally, the assembly is covered by
a damp and oil resistant covering.
13.4.2.4 Conduit Connectors
Conduit connectors are used to join sections of conduit together to form a control run in an aircraft.
The connectors are similar in design to all-metal pipe couplings, with the exception that adapter nipples
are not fitted to them. There are several different types of connector, each of which is designed for a
particular application in an aircraft (e.g. to allow the conduit to pass through a pressure bulkhead into
an aircraft’s pressure cabin without the loss of cabin pressure). The standard type of conduit connector
consists of two externally threaded connector nipples that are screwed into each end of an internally
threaded connector body (see diagram below). To couple two lengths of conduit together, the
connector nipples are pushed onto the end of each piece of conduit and the ends of conduit suitably
flared with a flaring tool. The nipples are then screwed into the connector body and tightened to firmly
retain the conduit by gripping the flare, holding it firmly onto a shoulder at the base of each of the
connector body holes.
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A pressure type of bulkhead connector is illustrated in the diagram below. In this type of connector, the
conduit flare is retained on a coned seating by a collar located in a recess in the outer sleeve. When
the outer sleeve is tightened onto the connector body, the conduit flare is gripped tightly between the
cone and the collar to form an airtight seal. The connector is bolted to the cabin pressure bulkhead
and the joint between the connector and the bulkhead is sealed by a jointing washer.
13.4.3 TELEFLEX CONTROL UNITS & FITTINGS
To operate the Teleflex control system, the cable and the conduit are connected to control units at
each end of the control run, and sometimes at intermediate points along the run. At the intermediate
points, other control units and fittings are used to direct the run through the aircraft. The control unit at
the transmitting end of the system is located in the aircraft’s cockpit and is usually a lever operated unit,
or alternatively, a simple push-pull operated control. The control system movement at the receiving
end of the system is controlled by either a wheel unit, which is basically similar in design to the
transmitting unit, or by one of several types of sliding end fitting.
13.4.3.1 Wheel Units
The Teleflex wheel unit consists of a light alloy casing in which is housed one or more gear wheels.
Each gear wheel has teeth cut around its periphery to suit either the left-hand helical windings of the
type DS 23/2 cable, or the right-hand helical windings of the type DS 380 cable. Hence the units are
not interchangeable with each other. The casing of each unit is machined to accept the gear wheel
and cable meshed together, thus ensuring that the cable is kept in mesh with the gear wheel at all
times. There are various types of wheel unit, some of which are as follows:
Single entry unit.
Double entry unit.
Straight-lead unit.
Distributor box.
Junction box unit.

Single entry unit. The single entry
type of wheel is the type commonly
used at the transmitting end of a
Teleflex control run. The cable enters
the unit via a conduit connector and is
located in a slot in the gear wheel (see
diagram to the right). Rotary travel of
the unit is limited to 270 degrees of
gear wheel travel, and a minimum of
40 degrees cable engagement on the
wheel must be maintained at all times.
Conversely, at the extreme end of the
travel, the cable must not foul that part
of the cable already wrapped around
the gear wheel.
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

Double entry unit. In the double entry wheel unit, the cable
enters the unit by a conduit connector and after wrapping
around the gear wheel, the cable emerges via another
conduit connector at a point 90º, 120º or 180º from the point
of entry. These units usually known as 90º, 120º or 180º
wrap units, whichever is applicable (see diagram to the left).
At the point the cable emerges from the wheel unit, it enters
a short piece of conduit, known as the ‘spent travel tube’.
The tube is fitted to the unit to prevent the cable from fouling
the aircraft’s structure, and also to prevent the ingress of
moisture and foreign bodies into the wheel unit.
Straight-lead unit. In the straight-lead
wheel unit, the cable passes straight
through the unit and hence only
emerges with a few teeth on the gear
wheel (see diagram to the right).
Consequently, this type of unit is not
suitable for a heavily loaded control
system. The unit can, however, be
interposed in a control run without the
need to break the cable, and it can also
be fitted to the transmitting or receiving
end of a control system.

Junction box unit. A junction box unit is installed when it is
necessary to reverse the direction of travel of the Teleflex
control run, or to add a branch into the run to enable an
additional control to be operated (e.g. simultaneous operation
of a control on the port and starboard side of the aircraft). In
one type of junction (see diagram to the left), two cables pass
through the junction box body diametrically opposite to each
other, and engage on either side of a single gear wheel. The
function of this type of box is either to:
 Operate one cable and then transfer the movement to the
other cable via the gear wheel, but in the reverse direction.
 To rotate the gear wheel and thus move both cables simultaneously.
An alternative arrangement contains a double gear wheel in which the cables pass through the box
side by side, and thus transmit movement in the same direction.
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
Distributor box. The design of the distributor box is basically similar to the double entry wheel unit
described earlier, but with an extra gear attached to the face of the gear wheel. The additional gear
wheel drives a pinion on a cross shaft, which in turn engages with a torsion drive (see diagram
below)
Wheel unit damping device. In some transmitting wheel units, a damping device is fitted to the unit
enabling the friction in the control handle to be adjusted. The damping device, which is fitted to
controls such as throttle and propeller pitch levers, ensure that control settings are not altered by
aircraft vibration. Usually, the device consists of a spring-loaded friction plate that is pressed against
the gear wheel. Adjustment of the amount of friction generated in the unit may be effected by rotation
of a knurled hand nut fitted to the control lever pivot, to either increase or decrease the friction.
13.4.3.2 Pull-Push Control Units
Pull-push control units are often installed in an aircraft instead of wheel units. These are used where
the control is to operate against a light load, and where fine adjustment of the control is unnecessary.
There are various types of pull-push unit in use, one of which is illustrated below. The operating cable
is secured to the operating handle by a lock spring and plug. Some pull-push units have locking
devices fitted to them to retain the operating handle in a set position, whilst other are spring-loaded to
ensure that the control returns to its normal position.
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13.4.3.3 Sliding End Fittings
A sliding end fitting is used instead of a wheel unit when a push-pull action is required at the receiving
end of the control. Various fittings are available for use, the choice of fitting depending upon the
method used to attached the fitting to the component. In each case the fitting is comprised of a guide
tube, which terminates in: a forkend, an eye, a ball joint, an internal or an external threaded fitting. The
cable is attached to the fitting either by means of a special collet attachment, or by means of a lock
spring and plug. The diagram below illustrates some of the types of end fitting that may be found in a
modern aircraft installation.
13.4.3.4 Swivel Joint
A swivel joint is a form of universal joint that may be used instead of a wheel unit, provided that the
rotary movement imparted to the control mechanism lever does not exceed 90º. This type of joint (see
diagram below), consists of a ball and socket connector that is located inside a housing and attached
to the end of a rigid conduit. The housing itself is firmly attached to the aircraft’s structure. The ball is
welded to a length of tubing of the same dimensions as the rigid conduit, and a suitable sliding end
fitting is attached to the end of the control cable to ensure that the guide tube slides freely over the
swivel joint tube. The angular travel of the swivel joint is limited to 8º from the central axis of the
conduit and hence, it is essential during installation of the control to ensure that the angular travel of
the control, to which the swivel is attached, falls easily within this limit.
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13.4.3.5 Quick-Break Units
Quick-break units are used to enable Teleflex control systems to be dismantled easily at such points as
engine bulkheads, and fuselage and wing break points, without disturbing the setting of the controls.
Several types of break unit are in use, but all employ a similar type of construction to that shown in the
diagram below. The cable joint fittings consists of rods that are fixed to the cable and machined to form
interlocking slotted ends, that fit snugly into each other to form a joint.
13.4.3.6 Teleflex Cable Connectors
The Teleflex cable is coupled to the control system end fittings by one or two types of connector, which
are:
Screwed end split collet connector.
Lock spring connector.

Screwed end split collet connector. The screwed end slip collet connector (see diagram below)
consists of a body, which is bored and threaded to accept the helical wire of the Teleflex cable.
One end of the body is reduced in diameter and threaded to receive standard A.G.S. fittings, whilst
the other end is externally threaded to receive the outer sleeve and locknut. An inspection hole is
drilled through the body to enable the technician to check that the cable is inserted correctly into
the body of the fitting. The end of the body to which the cable is fitted is tapered to an angle of 40º,
drilled and then slotted to form a collet. The outer sleeve, or plug end, used for the No 380 size
control, is copper welded to the slider tube and screwed onto the collet radially onto the cable by a
tapered seating formed on the inside of the sleeve, locking the cable t the body. Finally, the locknut
is tightened onto the sleeve to prevent the body from becoming unscrewed. Alternatively, when the
No 2 size control is used, the tapered bore of the outer sleeve is used to form a housing for the
flared end of the slider tube. Tightening of the connection presses the flared end of the slider tube
against the collet on the body of the connection, thus locking the cable in the body in a similar way
to the No 3 connection.
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
Lock spring connector. The lock spring type of connector is used in many aircraft to attach
Teleflex cable to both sliding end fittings and pull-push control units. In this type of connector, the
sliding end attachment fitting is internally threaded for approximately one fifth of its length, and then
bored for a further two fifths of its length to house the cable lock spring. It is then bored to the end
of the fitting to provide a housing for the cable end. An inspection hole is drilled across the bore of
the fitting at a point just beyond the lock spring housing. A plug is permanently attached to the
sliding tube and threaded externally. It is then assembled into the end of the attachment fitting and
locked by a locknut and tabwasher. The cable is passed through the sliding tube and plug, and is
retained in position by a lock spring. The diagram below illustrated a typical lock spring connector.
13.4.3.7 Sealing of Teleflex Control Runs
A Teleflex control run must be sealed when it is installed in a pressurised aircraft cabin, to prevent air
from the cabin escaping to atmosphere through the control conduit. The control seal is formed by two
50mm (2”) lengths of 1.6mm (1/16”) diameter graphite coated asbestos string, 50mm apart, wound
around the cable between the raised helix wire. The seal packing is retained in position by the helix
wire and forms a permanent seal between the cable and the bore of the conduit. The point of entry of
the conduit into the pressure cabin is made pressure tight by means of a bulkhead connector. A
pressure type of greaser connection is fitted on the pressure side of the bulkhead connection. The
illustration below shows an example of a Teleflex control run that passes into a pressure cabin.
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13.4.4 AIRCRAFT FLEXIBLE CONTROL SYSTEMS
The cable used for control runs is extra flexible and made up into various lengths to suit the control
system, each length of cable has end fittings swaged in position. If a control cable becomes
unserviceable, a new cable complete with end fittings may be obtained, or a new cable may be made
by swaging new end fittings to a new length of cable. The length and tension of the cables is adjusted
by turnbuckles situated at convenient positions in the control run. If push-pull rods are used, they are
usually made of light alloy tube and have screwed end fittings which enable the length of the rods to be
adjusted.
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