d) Dcam Part 66 - Module 6

DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE

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f T o MODULE 6: MATERIALS ANDg HARDWARE y 66 CATEGORY n r i (DCAMa PART B1.1) r t e e e i n r i p g o n r E P S A M For Training Purposes Only

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE

WARNING

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This document is intended for the purposes of training only. The information contained herein is as accurate as possible at the time of issue, and is subjected to ongoing amendments where necessary according to any regulatory journals and documents. Where the information contained in this document is in variation with other official journals and/or documents, the latter must be taken as the overriding document. The contents herein shall not be reproduced in any form without the expressed permission of ETD.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE TABLE OF CONTENTS

6 AIRCRAFT MATERIALS .................................................................................................................................................................. 1 6.1 FERROUS METALS.................................................................................................................................................................. 5 6.2 NON-FERROUS METALS....................................................................................................................................................... 21 6.3 COMPOSITE STRUCTURES.................................................................................................................................................. 21

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6.4 TYPES OF CORROSION ........................................................................................................................................................ 55

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6.5 FASTENERS ........................................................................................................................................................................... 90 6.6 PIPES AND UNIONS..............................................................................................................................................................184

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6.7 SPRINGS................................................................................................................................................................................201 6.8 BEARINGS .............................................................................................................................................................................208 6.9 TRANSMISSIONS ..................................................................................................................................................................212 6.10 CONTROL CABLES .............................................................................................................................................................233 6.11 ELECTRICAL CABLES ........................................................................................................................................................258

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE TABLE OF CONTENTS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE

6 AIRCRAFT MATERIALS Knowledge and understanding of the uses, strengths, limitation and other characteristics of structural metals is vital to properly construct and maintain any equipment especially airframes. In aircraft maintenance and repair, even slight deviation of from design specification of interior materials result in the loss of both lives and equipment. The selection of the correct material for a specific repair job demands familiarity with the most common physical properties of various metals.

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Strength, weight, and reliability are three factors which determine the requirements to be met by any material used in airframe construction and repair. The material must possess the strength required by the dimensions, weight and use. There are five basic stresses which metals may be required to withstand. These are: 1. 2. 3. 4. 5.

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Tension Compression Shear Bending Torsion

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Tension

Deformation

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE

Tensile Strength When a piece of sheet metal is pulled from each end, the resultant force is called tension. The ability to withstand tension is called tensile strength, and is measured in pounds per square inch. Yield Strength

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The ability of a metal to resist deformation is called yield strength. Example:

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M A simplified view of a material

The same material this time with an applied force

It breaks once the force exceeds ultimate strength of the material

When a tensile load is applied to a material, the material resists any deformation until its yield point is reached. However, once the yield point is reached, the metal stretches, and its molecular structure changes enough to increase the metal’s strength and therefore, resist further deformation. This continues until the ultimate load is reached, at which time, the material breaks.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE

Tension

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Tension

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Shear Strength

Shear strength describes a metal’s ability to resist opposing forces. A rivet that holds two or more sheets of metal together, resisting the force of the sheet trying to slide apart, is an example of shear load. When the rivets installed in a joint have more strength than the metal surrounding them, the joint is said to be loaded in shear. Example:

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Two simplified view of materials

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The two materials, joined with other materials, can withstand certain amount of force without deformation

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The joint breaks once the force exceed ultimate strength of the material for the joint

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE

Compression

Tension

Tension

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Bearing Strength

Bearing strength is the ability of a joint to withstand any form of crushing or excessive compressive distortion. Material under a compression load usually fails by buckling or bending. The force at which something buckles while being compressed varies with an objects length, cross sectional area and shape. Example:

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A simplified view of materials

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The same material this time with an applied force

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The material buckles once the force exceed the ultimate strength

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

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

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

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

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3. DUCTILITY The property of being able to be permanently extended by a tensile force. It is measured during a tensile, or stretching, test, when the amount of stretch (elongation), for a given applied load, provides an indication of a metal’s ductility. 4. ELASTICITY The ability of a metal to return to its original shape and size after the removal of any distorting force. The ‘Elastic Limit’ is the greatest force that can be applied without permanent distortion. 5. HARDNESS The ability of a metal to resist wear and penetration. It is measured by pressing a hardened steel ball or diamond point into the metal’s surface. The diameter or depth of the resulting indentation provides an indication of the metal’s hardness. 6. MALLEABILITY The ease with which the metal can be forged, rolled and extruded without fracture. Stresses, induced into the metal, by the forming processes, have to be subsequently relieved by heat-treatment. Hot metal is more malleable than cool metal. 7. PLASTICITY The ability to retain a deformation after the load producing it has been removed. Plasticity is, in fact, the opposite of elasticity. 8. TENACITY The property of a metal to resist deformation when subjected to a tensile load. It is proportional to the maximum stress required to cause the metal to fracture.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

9. TOUGHNESS The ability of a metal to resist suddenly applied loads. A metal’s toughness is tested by impact with a swinging pendulum of known mass. 10. STRENGTH There are several different measurements of the strength of a metal, as may be seen from the following sub-paragraphs

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10.1 TENSILE STRENGTH The ability to resist tension forces applied to the metal

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

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10.3 SHEAR STRENGTH The ability to resist side-cutting loads - such as those, imposed on the shank of a rivet, when the materials it is joining attempt to move apart in a direction normal to the longitudinal axis of the rivet. 10.4 BEARING STRENGTH The ability of a metal to withstand a crushing force.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

ALLOYING INGREDIENTS The main alloying agents of steel are: Carbon has a major effect on steel properties. Carbon is the primary hardening element in steel and allows heat treatment of steel to occur. Hardness and tensile strength increases as carbon content increases up to about 0.85% carbon. Low carbon steel contains 0.1 to 0.3 % carbon. Low carbon steels are used for the manufacture of safety wire and secondary structures. Medium carbon steel contains 0.3 and 0.5 % carbon. These steels are employed where a machining processes are required or where surface hardness is desireable. High carbon steels contain 0.5-- 1.05% carbon. These steels are used where extreme hardness is required, typical applications include springs, files and cutting tools.

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

Sulphur decreases ductility and weldability with increasing content. Sulphur levels are normally controlled to low levels. The only exception is free-machining steels, where sulfur is added to improve machinability.

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Manganese contributes to strength and hardness, but less than carbon. The increase in strength is dependent upon the carbon content. Increasing the manganese content decreases ductility and weldability, but less than carbon. Manganese has a significant effect on the hardenability of steel.

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Silicon is one of the principal deoxidizers used in steelmaking. Silicon is less effective than manganese in increasing as--rolled strength and hardness. In low--carbon steels, silicon is generally detrimental to surface quality.

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Phosphorous increases strength and hardness and corrosion resistance but decreases ductility

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Nickel increases the hardenability and impact strength of steels.

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Chromium is commonly added to steel to increase corrosion resistance and oxidation resistance, to increase hardenability, or to improve high-temperature strength. As a hardening element, Chromium is frequently used with a toughening element such as nickel to produce superior mechanical properties. At higher temperatures, chromium contributes increased strength.

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Molybdenum increases the hardenability of steel. Molybdenum may produce secondary hardening during the tempering of quenched steels. It enhances the creep strength of low--alloy steels at elevated temperatures.

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Vanadium increases the yield strength and the tensile strength of carbon steel. The addition of small amounts of Vanadium can significantly increase the strength of steels.

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Titanium is used to improve toughness

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Alloying elements

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

MATERIAL DESIGNATIONS Designations given to most low alloy steels are based upon an AISI (American Iron and Steel Institute) system that refers to the chemical composition of the alloy.

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The first two digits refer to the specific primary alloying elements, the last two digits (or the last three in a five-digit number) refer to the percentage of carbon contained in the alloy.  10XX -- refers to plain carbon steels (contain only carbon and manganese)  41XX -- refers to chromium and molybdenum alloy steels  43XX -- refers to nickel, chromium and molybdenum alloy steels  52100 -- refers to a chromium alloy with 1% carbon  93XX -- refers to a nickel, chromium and molybdenum alloy steel (with a different ratio between these elements than is contained in the 43XX alloys). For example, 4340 refers to a nickel-chromium-molybdenum alloy containing .40% carbon.

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9Ni - 4Co.30C is a specific trade name assigned to a nickel-cobalt alloy with .30% carbon. The 9 and 4 refer to the nominal percentages of nickel and cobalt in the alloy. The normally-used low alloy steels and their applicable strength ranges are shown. Use of these alloys is limited to the strength ranges shown. The European designations are slightly different. For further information refer to the ’Metallic Material List’ in the Structural Repair Manual (SRM) of the specific aircraft manufacturer.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Material designations

Metalworking Processes

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

After metal alloys are produced, they must be formed into useful shapes. Wrought objects are those formed by physically working the metal into shape, whereas cast items are formed by pouring molten metal into moulds. When it comes to mechanically working metal into a desired shape, there are three methods commonly used: 1. Hot-working 2. Cold-working 3. Extruding

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Hot – Working

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Hot-working is the process of forming metal at an elevated temperature when it is in its annealed or soft condition. Almost all steel is hot-worked from the ingot into a form which is either hot or cold worked to a finished shape. The ingot is then placed in a soaking pit to slow the cooling process until the molten interior gradually solidifies.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

After soaking, then it is worked into its desired shape through rolling and forging. Rolling consists of forming hot metal ingots with rollers to form sheets, bars and beams. Forging is a process where in a piece of metal is worked at temperature above its critical range. Forging is typically used to form shape through either pressing or hammering.

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Pressing

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Pressing is used to form large and heavy parts. Since a press is slow acting, its force is uniformly transmitted to the centre of the material being pressed. This affects the interior grain structure resulting in the best possible structure throughout.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Drop Forging

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Drop forging is a hammering process whereby a hot ingot is placed between a pair of formed dies in a machine called a drop hammer and a weight of several tons is dropped on the form upper die. This results in the hot metal being forced to take the form of the dies. Because the process is very rapid, the grain structure of the metal is altered, resulting in significant increases in the strength of the finished part.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Hammering

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Hammering is a type of forging that is usually used on small parts because it requires a metal worker to physically hammer a piece of metal into its finished shape. The advantage of hammering is that the operator has control over both the amount of pressure applied and the finishing temperature. Forging is usually referred to as smith forging and is used extensively where only a small number of parts are needed. In addition to the forming operation, hammering hardens the metal.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Cold – Working Cold-working is performed well below a metal’s critical temperature and ranges from the manual bending of sheet metal for skin repairs to drawing seamless tubing and wire. There are several cold-working processes; the two that are most common are cold-rolling and cold-drawing.

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Cold-Rolling

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Cold -rolling usually refers to the rolling of metal at room temperature.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Cold – Drawing

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Cold – drawing is used in making seamless tubing, wire and other forms of stock.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Extrusion

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Extrusion is the process of forcing metal through a die which imparts a required cross- section to the metal. Metals such as lead, tin and aluminium may be extruded cold, however most metals are heated. The advantage of the extrusion process is its flexibility. Example: Because of its workability, aluminium can be economically extruded to more shapes and larger sizes than is practicable with other metals. Many structural parts such as channels, angles, T-sections and Z-section are formed by the extrusion process.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Ferrous Metal (Iron) Any alloy containing iron as its chief constituent is called ferrous metal. The most common ferrous metal in aircraft structure is steel, an alloy of iron with a controlled amount of carbon added.

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Iron is a chemical element which is fairly soft malleable and ductile in its pure form. It is silvery white in color and is quite heavy. Iron combines readily with oxygen to form iron oxide, which is more commonly known as rust. Iron poured from a furnace into moulds is known as cast iron and normally contains more than two percent carbon and some silicon.

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Cast iron has few aircraft applications because of its low strength to weight ratio. However, it is used in engines for items such as piston rings where its porosity and wear characteristic allow it to hold a lubricant film.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Steel

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To make steel, pig iron is re-melted in a special furnace. Pure oxygen is then forced through the molten metal where it combines with carbon and burns. The molten steel is then poured into moulds where it solidifies into ingots. The ingots are placed in a soaking pit where they are heated to a uniform temperature of about 2,200º F/ 1204.4º C. They are then taken from the soaking pit and passed through steel rollers to form plate or sheet plate. Much of the steel used in aircraft construction is made in electric furnaces, which allow better control of alloying agents then gas-fired furnaces. An electric furnace is loaded with scrap steel, limestone and flux. The intense heat from the arcs melts the steel and the impurities mix with flux. Once the impurities are removed, controlled quantities of alloying agents are added, and the liquid metal in poured into moulds.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Stainless Steel

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Stainless steel is the classification of corrosion-resistant steels that contain large amounts of chromium and nickel. Their strength and resistance to corrosion make them well suited for high-temperature application such as firewalls and exhaust system components. The principal alloy stainless steel is chromium. The corrosion resistant steel most often used in aircraft construction is known as 18-8 steel because of its content of 18 percent chromium and 8 percent nickel. Stainless steel may be rolled, drawn, bent or formed to any shape.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Molybdenum One of the most widely used alloying elements for aircraft structural steel is molybdenum. It reduces the grain size of steel and increases both its impact strength and elastic limit. Molybdenum steels are extremely wear resistant and possess a great deal of fatigue strength and it’s used in high-strength structural members and engine cylinder barrels.

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Chrome-molybdenum (chrome-moly) steel is the most commonly used in aircraft. Its Society of Automotive Engineers (SAE) designation of 4130 denotes an alloy of approximately 1 per cent molybdenum and 0.30 percent carbon. It machines readily, is easily welded by either gas or electric arc, and responds well to heat treatment.

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Heat- treated SAE 4130 steel has an ultimate tensile strength about four times that of SAE 1025 steel, making for landing gear structure and engine mounts.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Heat Treatment of Steel Iron is an allotropic metal, meaning it can exist in more than one type of lattice structure, depending on temperature. Pure molten iron begins to solidify at 2,800 º F. Its structure at this point is known as the Delta form. If cooled to 2,554 º F, the atoms rearrange themselves into a Gamma form. Iron in this form is nonmagnetic. When nonmagnetic gamma iron in this form is cooled to 1,666 º F, another change occurs and the iron is transformed into a nonmagnetic form of Alpha structure.

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There are two basic forms of steel when it comes to heat treatment. They are ferrite and austenite.

Austenite

Ferrite is an alpha solid solution of iron containing some carbon and exists at temperature below the lower critical temperature. Above this lower critical temperature, the steel begins to turn into austenite, which consists of gamma iron containing carbon. As the temperature increases the transformation of ferrite into austenite until the upper critical temperature is reached.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Below the alloy’s lower critical temperature, the carbon which exists in the steel in the form of iron carbides is scattered throughout the iron matrix as a physical mixture. When the steel is heated to its upper critical temperature, this carbon dissolves into matrix as a physical mixture.

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Heat Treatment

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Heat treatment is a series of operations involving the heating and cooling of metal in the solid state. Its purpose is to make the metal more useful, serviceable, and safe for a definite purpose. By heat treating a metal can be made harder, stronger and more resistant to impact. Heat treating can also make a metal softer and more ductile. All heat-treating processes are similar in that they involve the heating and cooling of metals. They differ however in the temperatures to which the metal is heated and the rate at which it is cooled. A pure metal cannot be hardened by heat treatment because there is little change in its structure when heated.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Heat –Treating Equipment

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Successful heat treating requires close control over all factors affecting the heating and cooling of metals. The furnace must be of the proper size and type and must be so controlled that temperature are kept within prescribed for each operation. Even the atmosphere within the furnace affects the condition of the part being heat–treated. The quenching equipment and the quenching medium must be selected to fit the metal and the heat–treating operation. There are many different types and sizes of furnaces used in heat treatment. Furnaces are designed to operate in certain specific temperature ranges and attempted use in other rangers frequently results in work of inferior quality. Furnaces heated by electricity the heating elements are generally in the form of wire or ribbon. Such furnaces commonly operate up to a maximum temperature of about 2000 º F.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Heating The object in heating is to transform parasite (a mechanical mixture of iron carbide that exists in a finely mixed condition) to austenite as the steel is hated through the critical range. Steel begins to appear dull red at about 1000 º F and as the temperature increases the colour changes gradually through various shades of red to orange, to yellow and finally to white.

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Soaking

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The temperature of the furnace must be held constant during the soaking period, since it is during this period that rearrangement of the internal structure of the steel takes place. The length of the soaking period depends upon the type of steel and the size of the part. As a general rule, a soaking period of 30 minutes to 1 hour is sufficient for the average heat-treating operation.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

Cooling Various rates of cooling are used to produce the desired results, still air is a slow cooling medium, but is much faster than furnace cooling. Liquids are the fastest cooling media and therefore used in hardening steels. There are three commonly used quenching liquids brine, water and oil. Brine is the most severe medium, water is next and oil is the least severe. Generally an oil quench is used for alloy steels and brine or water for carbon steels.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Portable Quench Tank

Quenching Media

Quenching solutions act only through their ability to cool the steel. Most requirements for quenching media are met satisfactorily by water. The rate of cooling is relatively rapid during quenching in brine, somewhat less rapid in water and slow in oil. Brine usually is made of a 5 to 10 percent solution of salt (sodium chloride) in water. In addition to its greater cooling speed, brine has the ability to “throw” the scale from steel during quenching.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS - FERROUS (DCAM 6.1 L1 & L2)

STEEL APPLICATIONS General The base material iron is a chemical element which, in its pure form, is a very soft, malleable and ductile metal which is easy to form and shape. It readily combines with oxygen to form iron oxide (rust), and so is alloyed, primarily with carbon, but also with other elements. When molten iron is alloyed with more than 2% Carbon and poured into a mould, cast iron is formed. Cast iron has limited uses in the aviation industry due to low strength to weight ratio and brittleness.

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Iron is extracted from iron ore by mixing it with coke and limestone and heating it in a furnace. The process extracts the oxygen from the ore, and allows the iron to sink to the bottom of the furnace. The limestone reacts with any impurities in the molten iron and floats to the surface to form a slag.

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To make steel, the pure iron is remelted in a special furnace where carbon is introduced along with other alloying elements to achieve the desired characteristics. Description

Steel is an excellent engineering material with many applications. For aircraft use, however, it does have some significant problems. The main restrictions are its high density (approximately 3 times the density of aluminium) and its susceptability to corrosion. The corrosion of steel can be reduced by the addition of certain alloying elements, but this can have significant effects on properties and costs. Between 9 and 16% (Airbus A320: 9% , Boeing B777: 11%) of an aircraft’s structure is alloy steel and stainless steel. The high strength and high modulus of elasticity are the primary advantages of the high-strength steels. This is useful for designs with space limitations such as with some landing gear components. Alloy selection considerations include service temperature, strength, stiffness fatigue properties and fabricability.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Steel Application

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

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

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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.

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

Stress 

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

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

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

Stress 

Load Area

400 400   20  37 N / mm 2 2 2 r   25

Exercise 1 Calculate the tensile stress in a steel rod, with a cross-section of 10 mm x 4 mm, when it is subjected to a load of 100 N.

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Example 2 A structural member, with a cross-sectional area of 05m2, is subjected to a load of 10 MN. Calculate the stress in the member in; (a) MN/m2 and (b) N/mm2 (a)

(b)

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Stress 

Load Area

10  20MN / m 2 05

1N/mm  1 MN/m 2

2

So Stress  20 N/mm

2

As the load in the tensile test is increased from zero to a maximum value, the material extends in length. The amount of extension, produced by a given load, allows the amount of induced STRAIN to be calculated. Strain is calculated by measuring the extension and dividing by the original length of the material. Note: Both measurements must be in the same units, though, since Strain is a ratio of two lengths, it has no units.

Strain 

Extension Original Length

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

Note: When calculating stress in large structural members, it may be more convenient to measure load in Mega-Newtons (MN, or N6) and the area in square metres (m2). When using such units, the numerical value is identical to that if the calculation had been made using Newtons and mm2. i.e. A Stress of 1 N/mm2 = l MN/m2

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Example 3 An aluminium test piece is marked with a 20 mm gauge length. It is subjected to tensile load until its length becomes 2115 mm. Calculate the induced strain.

Extension  21 15 - 20  1  15 mm Strain 

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Extension 1  15   0  0575 (no units) Original Length 20

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

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig 1

The graph can be considered as comprising two major regions. Between points 0 and A, the material is in the Elastic region (or phase), i.e. when the load is removed the material will return to its original size and shape. In this region, the extension is directly proportional to the applied load. This relationship is known as ‘Hooke's Law’, which states:

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

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

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Aircraft structural designers’ interest in materials does not extend greatly beyond the elastic phase of materials. Production engineers, however, are greatly interested in material properties beyond this phase, since the forming capabilities of materials are dependent on their properties in the plastic phase.

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

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UTS

Point B Yield Point

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Load/Extension Diagram for Mild Steel Fig. 2

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

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

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Point of Fracture

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Plastic Region

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Small Elongation

Zero Elongation

Large Elongation

(a)

(b)

(c)

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

(a) represents a brittle material (glass)

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

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

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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

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 Reduction in area (at the neck)

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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.

Final Extension  100 Original Gauge Length

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Percentage elongation

Example 4 In a tensile test, on a specimen with 150.5 mm gauge length, the length over the gauge marks at fracture were 176.1 mm. What was the percentage elongation?

Elongation 

Final Extension 176.1 - 150.5  100   100  17.009%  17% Gauge Length 150  5

Proof Stress Many materials do not exhibit a yield point, so a substitute value must be employed. The value chosen is the ‘Proof Stress’, which is defined as: The tensile stress, which is just sufficient to produce a non-proportional elongation, equal to a specified percentage of the original gauge length.

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

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If the 0.2% Proof Stress is required, then 0.2% of the gauge length is marked on the extension axis. A line, parallel to the straight-line portion of the graph, is drawn until it intersects the non-linear portion of the curve. The corresponding load is then read from the graph. Proof Stress is calculated by dividing this load by the original cross-sectional area. 0.1% Proof Stress will produce permanent set equivalent to one thousandth of the specimen's original length.

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

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Acquiring the Proof Stress from a Load/Extension Graph Fig. 4

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

Stress is a measure of stiffness Strain

ie .

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This value, which is of great importance to designers, is known as ‘the Modulus of Elasticity, or Young’s Modulus’, and is signified by use of the symbol E. Thus E = Stress divided by Strain and, since Strain has no units, the unit for `E' is the same as Stress. i.e. lbf/in2, tonf/in2 (also psi and tsi), or the European and SI units such as kN/m2, MN/m2 and kPa or MPa.

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The actual numerical value is usually large, as it is a measure of the stress required to theoretically double the length of a specimen (if it did not break first).

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A typical value of E for steel would be 30 x 106 psi. or 210,000 MN/m2

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

2000 x

 Aluminium

10,000 x

 Steel

30,000 x

 Diamond

171,000 x

Tensile Testing of Plastics This is conducted in the same way as for metals, but the test piece is usually made from sheet material. Although the basic load/extension curve for some plastics is somewhat similar to metal curves, changes in test temperature or the rate of loading can have a major effect on the actual results. Even though the material under test may be in the elastic range, the specimen may take some time to return to its original size after the load is removed.

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Compression Test Machines for compression testing are often the same as those used for tensile testing, but the test specimen is in the form of a short cylinder. The Load/Deflection graph in the elastic phase for ductile materials is similar to that in the tensile test. The value of `E' is the same in compression as it is in tension. Compression testing is seldom used as an acceptance test for metallic or plastic materials (except for cast iron).

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

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

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1. Brinell Test

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

Diameter (Area) of resulting Indentation

The Brinell Hardness Test Fig. 5

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

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M The Vickers Hardness Test Fig. 6

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

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

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M The Rockwell Hardness Test Fig. 7

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No precise relationship exists between the various hardness numbers, but approximate relationships have been compiled. Some comparative values between Brinell Vickers and Rockwell are shown in Table 1. Table 1 COMPARATIVE HARDNESS VALUES MATERIAL Aluminium alloy Mild steel Cutting tools Note:

BHN 100 130 650

HV 100 130 697

ROCKWELL B 57 B 73 C 60

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

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

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

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

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

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

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Test Piece

Impact Test Fig. 8

.

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

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a. Creep Creep can be defined as the continuing deformation, with the passage of time, of materials subjected to prolonged stress. This deformation is plastic and occurs even though the acting stress may be well below the yield stress of the material.

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At temperatures below 0.4T (where T is the melting point of the material in Kelvin), the creep rate is very low, but, at higher temperatures, it becomes more rapid. For this reason, creep is commonly regarded as being a high-temperature phenomenon, associated with super-heated steam plant and gas turbine technology. However, some of the soft, low-melting point materials will creep significantly at, or a little above, ambient temperatures and some aircraft materials may creep when subjected to over-heat conditions.

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b. Creep in Metals When a metallic material is suitably stressed, it undergoes immediate elastic deformation. This is then followed by plastic strain, which occurs in three stages (refer to Fig. 9):  Primary Creep - begins at a relatively rapid rate, but then decreases with time as strain-hardening sets in.  Secondary Creep - the rate of strain is fairly uniform and at its lowest value.

 Tertiary Creep - the rate of strain increases rapidly, finally leading to rupture. This final stage coincides with gross necking of the component, prior to failure. The rate of creep is at a maximum in this phase.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Three Stages of Creep Fig. 9

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c.

Effect of Stress and Temperature on Creep

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

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

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

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

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e. 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. f. Fatigue An in-depth survey, in recent years, revealed that over 80 percent of failures of engineering components were caused by fatigue. Consequences of modern engineering have led to increases in operating stresses, temperatures and speeds. This is particularly so in aerospace and, in many instances, has made the fatigue characteristics of materials more significant than their ordinary, static strength properties. Engineers became aware that alternating stresses, of quite small amplitude, could cause failure in components, which were capable of safely carrying much greater, steady loads. This phenomenon of small, alternating loads causing failure was likened to a progressive weakening of the material, over a period of time and hence the attribution of the term ‘fatigue’. Very few constructional members are immune from it, and especially those operating in a dynamic environment. Experience in the aircraft industry has shown that the stress cycles, to which aircraft are subjected, may be very complex, with occasional high peaks, due to gust loading of aircraft wings. For satisfactory correlation with in-service behaviour, full-size or large-scale mock-ups must be tested in conditions as close as possible to those existing in service.

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

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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.

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

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

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

Test Piece

Load

Simple Fatigue Testing Fig 10

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

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

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

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Stress (S)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fatigue Limit

40 – 60 % UTS

Number of Cycles (N)

A S-N Curve for a Steel Material Fig. 11

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Many non-ferrous metals, however, show a different characteristic from steel (refer to Fig. 12). In this instance there is no fatigue limit as such and it can be seen that these materials will fail if subjected to an appropriate number of stress reversals, even at very small stresses. When materials have no fatigue limit an endurance limit together with a corresponding number of cycles is quoted instead. It follows that components made from such materials must be designed with a specific life in mind and removed from service at the appropriate time. The service fatigue lives of complete airframes or airframe members are typical examples of this philosophy.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M An S-N Curve for an Aluminium Alloy Fig. 12

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Non-metallic materials are also liable to failure by fatigue. As is the case with metals, the number of stress cycles, required to produce a fatigue failure, will increase as the maximum stress in the loading cycle decreases. There is, however, generally no fatigue limit for these materials and some form of endurance limit must be applied. The importance of fatigue strength can be illustrated by the fact that, in a high- cycle fatigue mode, a mere 10% improvement in fatigue strength can result in a 100-times life improvement.

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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.

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Nevertheless, fatigue cracking occurs frequently, and even the most sophisticated engineering product does not possess immunity from this mode of failure. Such failures are often due to unforeseen factors in design, environmental or operating conditions, material, and manufacturing processes. Two essential requirements for fatigue development in a material are: 

An applied stress fluctuation of sufficient magnitude (with or without an applied steady stress).



A sufficient number of cycles of that fluctuating stress.

The stress fluctuations may be separated by considerable time intervals, as experienced in aircraft cabin pressurisation, during each take-off (e.g. daily), or they may have a relatively short time interval, such as encountered during the aerodynamic buffeting/vibration of a wing panel. The former example would be considered to be low-cycle fatigue and the latter to be high-cycle fatigue. In practice, the level of the fluctuating stress, and the number of cycles to cause cracking of a given material, are affected by many other variables, such as stress concentration points (stress raisers), residual internal stresses, corrosion, surface finish, material imperfections etc.

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

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Resonant frequencies are undesirable (and in some cases could be disastrous), so it is important to ensure that, over their normal operating ranges, critical components are not vibrated at their natural frequencies and so avoid creating resonance. The resonant frequency, of a component, is governed by its mass and stiffness and, on certain critical parts, it is often necessary to do full-scale fatigue tests to confirm adequate fatigue life before putting the product into service.

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Fatigue Metallurgy Under the action of fatigue stresses, minute, local, plastic deformation on an atomic scale, takes place along slip planes within the material grains. If the fatigue stresses are continued, then micro cracks are formed within the grains, in the area of the highest local stress, (usually at or near the surface of the material). The micro cracks join together and propagate across the grain boundaries but not along them. A fatigue fracture generally develops in three stages (refer to Fig. 13):   

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

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Propagation (crack growth)

Ultimate (rapid) fracture

The Three Stages of Fracture Fig. 13

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The resultant fractured surface often has a characteristic appearance of:  An area, on which a series of curved, parallel, relatively smooth ridges are present and are centred around the starting point of the crack. These ridges are sometimes called conchoidal lines or beach marks or arrest lines.  A rougher, typically crystalline section, which is the final rapid fracture when the cross-section is no longer capable of carrying its normal, steady load.

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The arrest lines are, normally, formed when the loading is changed, or the loading is intermittent. However, in addition to these characteristic and informative marks, there are similar, but much finer lines (called ‘striations’), which literally show the position of the crack front after each cycle. These striations are obviously of great importance to metallurgists and failure investigators when attempting to estimate the crack initiation and/or propagation life. The striations are often so fine and indistinct that electron beam microscopes are required to count them.

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In normal circumstances, a great deal of energy is required to `weaken' the material sufficiently to initiate a fatigue crack, and it is not surprising, therefore, to find that the nucleation phase takes a relatively long time. However, once the initial crack is formed, the extremely high stress concentration (present at the crack front) is sufficient to cause the crack to propagate relatively quickly, and gaining in speed as the crack front not only increases in size, but also reduces the component cross-sectional area. A point is eventually reached (known as the 'critical crack length') at which the remaining cross-section is sufficiently reduced to cause a gross overloading situation, and a sudden fracture finally occurs. It is not unusual for the crack initiation phase to take 90% of the time to failure, with the propagation phase only taking the remaining 10%. This is one of the major reasons for operators of equipment being relatively unsuccessful in detecting fatigue cracks in components before a failure occurs. Fatigue Promoters

As fatigue cracks initiate at locations of highest stress and lowest local strength, the nucleation site will be:  dictated largely by geometry and the general stress distribution  located at or near the surface or

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

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a. Design Apart from general stressing, the geometry of a component has a considerable influence on its susceptibility to fatigue. A good designer will therefore minimise stress concentrations by:  avoiding rapid changes in section and

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 using generous blend radii or chamfers to eliminate sharp corners

b. Manufacture While the designer may specify adequate blend radii, the actual product may still be prone to fatigue failure if the manufacturing stage fails to achieve this sometimes-seemingly unimportant drawing requirement.

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Several other manufacturing-related causes of premature fatigue failure exist, the most common of which are:  Inherent material faults: e.g. cold shuts, pipe, porosity, slag inclusions etc.  Processing faults: e.g. bending, forging, grinding, shrinking, welding, etc.

 Production faults: e.g. incorrect heat-treatment, inadequate surface protection, poor drilling procedures, undue force used during assembly, etc  In-service damage: e.g. dents, impact marks, scratches, scores, tooling marks etc.

c. Environment One of the most potent environmental promoters of fatigue occurs when the component is operating in a corrosive medium. Steel (normally), has a welldefined fatigue limit on the S-N curve but, if a fatigue test is conducted in a corrosive environment, not only does the general fatigue strength drop appreciably, but the curve also resembles the aluminium alloy curve (e.g. the fatigue failure stress continues to fall as the number of cycles increases). Other environmental effects such as fretting and corrosion pitting, erosion or elevated temperatures will also adversely affect fatigue strength.

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

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

 Material.

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

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 Geometry.

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

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b) Modify the geometry to change the vibration frequency or introduce a damping feature, to reduce the vibration amplitudes. c) Improve the surface finish or put a compressive stress in the skin (e.g. shot peen or cold expand).

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a. Cold Expansion (Broaching) Most fatigue failures occur whilst a material is subject to a tensile, alternating stress. If the most fatigue-prone areas, such as spar fastener holes, have a compression stress applied (refer to Fig. 14), they are significantly more resistant to fatigue failure. The fastener hole is initially checked for defects (using, usually, an Eddy Current NDT procedure) and the surface finish is further improved by reaming (and checked once again). A tapered mandrel is then pulled through the hole, resulting in a localised area of residual (compressive) stress which will provide a neutral or, at least, a significantly reduced level of fatigue in the area around the fastener hole

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Area around hole pre-stressed in compression

Tapered Mandrel pulled through fastener hole

Cold Expansion of Fastener Hole Fig.14

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Do's and Dont's – Preventing Fatigue Failures DO  Be careful not to damage the surface finish of a component by mishandling.  Use the right tools for assembling press-fit components etc.  Maintain drawing sizes and tolerances.  Keep the correct procedures (e.g. don't overheat when welding).  Avoid contact or near contact of components that might cause fretting when touching. DON'T  Leave off protective coverings - plastic end caps etc.  Score the surface.  Leave sharp corners or ragged holes.  Force parts unnecessarily to make them fit.  Work metal unless it is in the correct heat-treated state.

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

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

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

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

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d. Intelligent Skins Development Modern developments in aircraft structures will allow the structures to be designed and built with a variety of sensors and systems embedded into the 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.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M i. Smart Structures The generic heading ‘Smart Structures’ actually covers three areas of development:



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



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



Intelligent Skins. Skin embedded with fibre optic sensors

Smart Structures perceived benefits include:



Self-diagnostic in the monitoring of structural integrity



Reduced life cycle costs



Reduced inspection costs



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



From a military point of view – an improvement in ‘Stealth’ characteristics.

A fully monitored and self-diagnostic system could: 

Assess structural integrity.



Pinpoint structural damage.

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Process flight history.

Composite laminates, containing embedded fibre optic sensors can be used for SHM, including fatigue monitoring and flight envelope exceedance monitoring and their advantages include:  Cover a greater area of structure   Not prone to electrical interference Less vulnerable to damage when embedded in the plies Increased knowledge of structural loads aids designers

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6.2 NON-FERROUS METALS Metal used on today’s aircraft contains no iron. The term that describes metals which have elements other than iron as their base is non-ferrous. Aluminium, copper, titanium and magnesium are some of the more common non-ferrous metals used in aircraft construction and repair.

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Aluminium and Its Alloys

Pure aluminium lacks sufficient strength to be used for aircraft construction. However, its strength increases when it is alloyed or mixed with other compatible metals.

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Example: Aluminium is mixed with copper or zinc, the resultant alloy is as strong as steel with only one third of the weight, and the corrosion resistance possessed by the aluminium carries over to the newly formed alloy.

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Aluminium is one of the most widely used metals in modern aircraft construction. It is vital to the aviation industry because of its high strength-to-weight ratio and its comparative ease of fabrication. Aluminium alloys, although strong are easily worked because they are malleable and ductile. They may be rolled into sheets as thin as 0.0017 inch or drawn into wire 0.004 inch in diameter. Most aluminium alloy sheet stock used in aircraft construction ranges from 0.016 to 0.096 inches in thickness, while some of the larger aircraft use sheet stock which may be as thick as 0.356 inch. The various types of aluminium may be divided into two general classes:

1. The casting alloys (those suitable for casting in sand, permanent mold or die castings) 2. The wrought alloys (those which may be shaped by rolling, drawing or forging)

Of these two, the wrought alloys are the most widely used for stringers, bulkheads, skin, rivets and extruded section.

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Aluminium Casting Alloys Aluminium casting alloys are divided into two basic groups: 1. The physical properties of the alloys are determined by the alloying elements and cannot be changed after the metal is cast 2. The alloying elements make it possible to heat treat the casting to produce the desired physical properties.

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The casting alloys are identified by a letter preceding the alloy number. When a letter precedes a number, it indicates a slight variation the composition of the original alloy.

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Example: In casting alloy 214, the addition of zinc to improve its qualities is indicated by the letter A in front of the number, thus creating the designation A214.

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When castings have been heat treated, the heat treatment and the composition of the casting is indicated by a letter T, followed by an alloying number. Example: In sand casting alloy which has several different compositions and tempers and is designated by 355-T6, Aluminium alloy castings are produced by one of three basic methods: 1. Sand mould 2. Permanent mould 3. Die cast

In casting aluminium, it must be remembered that in most cases different types of alloys must be used for different types of castings. Sand and permanent mould castings are parts produced by pouring molten metal into a previously prepared mold, allowing the metal to solidify or freeze ad then removing the part. The permanent mould process is a later development of the sand casting process, the major difference being in the material from which the moulds are made. The advantage of this process is that there are fewer openings (called porosity) than in sand casting. The sand and the binder, which is mixed with the sand to hold it together, give off a certain amount of gas which causes porosity in a sand casting. The permanent mould castings are used to obtain higher mechanical properties, better surfaces, or more accurate dimensions.

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Wrought Aluminium

Wrought aluminium and wrought aluminium alloys are divided into two general classes: non-heat-treatable alloys and heat-treatable alloys. Non-heat treatable alloys are those in which the mechanical properties are determined by the amount of cold-work introduced after the final annealing operation. The mechanical properties obtained by cold working are destroyed by any subsequent heating and cannot be restored except by additional cold working. For heat-treatable aluminium alloys the mechanical properties are obtained by heat treating to a suitable temperature, holding at that temperature long enough to allow the alloying constituent to enter into solid solution and then quenching to hold the constituent in solution. In the wrought form, commercially pure aluminium is known as 1100. It has a high degree of resistance to corrosion and is easily formed into intricate shapes. It is relatively low in strength, however and does not have the strength required for structural aircraft parts. The most widely used alloys in aircraft construction are hardened by heat treatment rather than by cold-work. The heat-treatable alloys will turn black due to the copper content, whereas the others will remain bright. In the case of clad material, the surface will remain bright, but there will be a dark area in the middle when viewed from the edge.

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Alclad Aluminium The terms “Alclad and Pureclad” are used to designate sheets that consist of an aluminium alloy core coated with a layer of aluminium to a depth of approximately 5 ½ percent on each side. The pure aluminium coating affords a dual protection for the core, preventing contact with any corrosive agents, and protecting the core electrolytically by preventing any attack caused by scratching or from other abrasion.

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Magnesium and Its Alloys

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Magnesium alloys are used for castings and in its wrought form is available in sheets, bars, tubing and extrusions. Magnesium is one of the lightest metals having sufficient strength and suitable characteristic for use in aircraft structures. Magnesium is highly susceptible to corrosion and tends to crack. The corrosion problem is minimized by treating the surface with chemicals that form an oxide film to prevent oxygen from reaching the metal. Another important step in minimizing corrosion is to always use hardware such as rivets, nuts, bolts, and screw than are made of compatible material.

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Magnesium burns readily in a dust or small particle form. Caution must be exercised when grinding and machining magnesium. If a fire should occur, extinguish it by something it with dry sand or some other dry material that excluded air from the metal and cools its surface. If water is used, it will only intensify the fire. Titanium and Its Alloys

Titanium and its alloys are lightweight metals with very high strength. Pure titanium weights 163 pounds per cubic inch which is about 50 % lighter than stainless steel. In addition to its light weight and high strength, titanium and its alloys have excellent corrosion resistance characteristic, particularly to the corrosive effects of salt water. Because of its high strength to weight ratio, titanium is now used extensively in the civilian aerospace industry. Although rare on commercial aircraft, modern jet transports now utilize alloys containing 10 to 15 percent titanium in structural areas.

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Nickel and Its Alloys As an aircraft engineer / technician, we need to be familiar with two nickel alloys. They are monel and inconel.

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Monel

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Monel contains about 68% nickel and 29% copper, with small amounts of iron and manganese. It can be welded and has very good machining characteristics. Certain types of monel, especially those containing small percentages of aluminium are heat-treatable to tensile strengths equivalent to steel. Monel works well in gears and parts that require high strength and toughness, as well as for parts in exhaust systems that require high strength and corrosion resistance at higher temperature. K- Monel

K- Monel is a non-ferrous alloy containing mainly nickel, copper and aluminium. It is produced by adding a small amount of aluminium to the Monel formula. It is corrosion resistant and capable of being hardened by heat treatment. K- Monel has been successfully used for gears and structural members in aircraft which are subjected to corrosive attacks. This alloy is non-magnetic at all temperature. K- Monel sheet has been successfully welded by both oxy-acetylene and electric arc welding.

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Inconel and its Alloys High strength, high temperature alloys containing approximately 80 percent nickel, 14 percent chromium and small amounts of iron and other elements are commonly referred to as inconel. Inconel alloys are frequently used in turbine engines because of their ability to maintain their strength and corrosion resistance under extremely high temperature conditions.

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Inconel and stainless steel are similar in appearance. Often, a test is used to differentiate between unknown metal samples. A common test involves applying one drop of cupric chloride and hydrochloric acid solution to the unknown metal and allowing it to remain for 2 minutes. At the end of the dwell period, a shiny spot indicates that the material is inconel, whereas a copper-colored spot identifies stainless steel.

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Brass Brass is a copper alloy containing zinc and small amounts of aluminium, iron, lead, manganese, nickel, and phosphorous. Brass with a zinc content of 30 to 35 percent is very ductile, but that containing 45 percent has relatively high strength.

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Muntz metal is a brass composed of 60 percent copper and 40 percent zinc.It has excellent corrosion–resistant qualities in salt water. This metal has an ultimate tensile strength. Its strength can be increased by head treatment. It is used in making bolts and nuts, as well as part that come in contact with salt water.

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Copper and Its Alloys Neither copper nor its alloys find much use as structural materials in aircraft construction. Due to its excellent electrical and thermal conductivity, however, copper is the primary metal used for electrical wiring.

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Copper is one of the most widely distributed. It is the only reddish – colored metal and is second only to silver in electrical conductivity. In aircraft, copper is used primarily in the electrical system for bus bars, bonding and as lock wire.

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Beryllium copper is one of the most successful of all the copper base alloys. This alloy contains about 97 percent copper, 2 percent beryllium and sufficient nickel to increase the percentage of elongation. The resistance of beryllium copper to fatigue and wear make it suitable for diaphragms, precision bearing bushings Bronze

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Bronze is a copper alloy that contains tin. A true bronze consists of up to 25 percent tin and along with brass, is used in bushing, bearings and fuel – metering valves. Bronzes with less than 25 percent tin are used in items such as tube fitting in aircraft.

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Titanium and Titanium Alloys Titanium falls between aluminium and stainless steel in term of elasticity, density and elevated temperature strength. It has melting point of from 2,730 º F to 3,155 º F. It is light, strong and resistant to stress–corrosion cracking. Titanium is approximately 60 percent heavier than aluminium and about 50 percent lighter than stainless steel.

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The ultimate yield strength of titanium drops rapidly above 800 º F. The absorption of oxygen and nitrogen from the atm temperature above 1000 º F makes the metal so brittle on long exposure. Titanium does have some merit for short–time exposure up to 3000 º F where strength is not important. Aircraft firewalls demand this requirement.

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Titanium is non – magnetic and has an electrical resistance comparable to that of stainless steel. In aircraft construction and repair, titanium is used for fuselage skins, engine shrouds, firewalls, longerons, frames, filting, air ducts and fasteners. Titanium is used for making compressor disks, spacer rings, compressor blades and vanes through bolts, turbine housing and miscellaneous hardware for turbine engines.

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Titanium in appearances is similar to stainless steel. One quick method to identify titanium is the spark test. Titanium Designations

The A - B – C classification of titanium alloys was established to provide a convenient and simple means of describing all titanium alloys. Titanium and titanium alloys possess three basic types of crystals:

1. A (alpha) – All around performance, good weldability; tough and strong in both cold and hot conditions, and resistant to oxidation. 2. B (beta) – Bend ability; excellent bent ductility; strong in both cold and hot conditions, but vulnerable to contamination. 3. C (combined alpha and beta for compromise performances) – Strong when cold and warm, but weak when hot; good bend ability; moderate contamination resistance; excellent forge ability.

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Heat-Treatment of Aluminium Alloys SAFETY PRECAUTIONS MUST BE OBEYED WHENEVER YOU ARE INVOLVED WITH HEAT-TREATMENTS. WARNING:BATHS, OVENS AND FURNACES ALL PRESENT DANGERS – FROM CORROSIVE AGENTS, HEAT AND ELECTROCUTION – EXERCISE EXTREME CAUTION WITH THESE METHODS AND WEAR ADEQUATE PROTECTIVE CLOTHING (APRONS, FACE MASKS, GOGGLES AND GLOVES) WHERE NECESSARY AND ENSURE THE CORRECT FIRE-FIGHTING APPLIANCES ARE AVAILABLE.

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Heat-treatment is a series of operations involving the heating and subsequent cooling of alloys in their solid state. Its purpose is to make the metal harder, stronger and more resistant to impact but it can also make the metal softer and more ductile for working into a required shape (bending etc.). One treatment cannot give all of these properties. Some treatments are achieved at the expense of others when, for example, a hardened material usually becomes more brittle.

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The heating and cooling cycles occur in most treatments and it is only the time and temperatures which differ. Aluminium alloys have two main heattreatments, which are referred to as solution heat-treatment and precipitation heat-treatment.

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The procedures for heat-treating aluminium alloys are critical if correct properties are to be obtained. Uniform heating is absolutely essential and two methods are used:  a muffle furnace  or a salt bath

The muffle furnace uses hot air, which circulates around an inner chamber in which the aluminium alloy is placed.

The salt bath employs molten mineral salts (water would evaporate long before the required temperatures were reached. The salts (usually nitrate of soda or similar) are solid at room temperature, but become liquid when they are electrically heated. Gradual heating of the bath is necessary to avoid spattering or spitting. The aluminium alloy (pre-dried, also to avoid spattering) can then be submerged within the heated liquid. Another precaution when using a salt bath is to avoid any adjacent flames or sparks, because the salts are inflammable. Accurate thermostatic control is vital, as narrow tolerances on temperatures are specified (typically plus or minus 5ºC). Quench tanks must be sited nearby the furnace or salt bath, to avoid delay between removing from the heating source and quenching. Most quench tanks contain cold water but hot water is sometimes specified (especially for heavy sections e.g. large forgings). Limits are also stipulated for the permissible period between heating and quenching which is known as the lag-time (typically 10 seconds max.). If these lag-times are exceeded, material properties or corrosion resistance may be adversely affected. If the cooling rate, during quenching, is too slow this may also affect the corrosion resistance.

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Thorough washing of the material is essential after salt bath heat-treatment to remove any salt residue. There is no limit to the number of times that heat-treatment may be carried out on normal aluminium/copper alloys but, if the material is clad with pure aluminium, for corrosion resistance (Alclad), then a maximum of three treatments is imposed.

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This is to limit the migration of copper, from the alloyed material, into the pure aluminium cladding, which would significantly reduce its corrosion resistance. Solution Treatment Solution treatment is sometimes called ‘re-crystallisation H.T’. This operation serves to distribute the copper uniformly throughout the aluminium (i.e. to create a solid solution). The heating may be achieved (as previously stated) in an oven or, more commonly (to obtain better overall heating), in a bath of special, molten salts. However, although the aluminium can accommodate 5% or so of copper in solid solution at high temperature, this condition is unstable at lower temperatures and, after the alloy has cooled to room temperature, most of the copper slowly comes out of solution and separates into local `islands' of copper aluminide.

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By cooling the alloyed metals very quickly (quenching), the copper becomes trapped 'in solution', making the aluminium very strong.

Age-Hardening The gradual formation of the copper alumide ‘islands’ (also referred to as ‘slip’), causes an increase in hardness and strength and these properties reach maximum values after several days (or weeks in some instances). Because of the time lapse involved, this gradual hardening is termed ‘age-hardening’. Although copper may be the major alloying element (in the ‘2000 series’ alloys) other elements, including magnesium and manganese can also be present. Although the aluminium/copper alloys are the most common age-hardened, high-strength metals, they are not unique. Aluminium, when alloyed with 5%-7% Zinc, is also able to be age-hardened. This is a more modern alloy than the aluminium/ copper type and is the highest-strength aluminium alloy in general use. This alloy is used in heavy loaded applications such as Main Spars, Landing Gear and Mainplane Attachment brackets etc.. Annealing Annealing, as with steel, serves to soften the aluminium alloy, to enable it to be worked without cracking. Even in this condition, ageing will gradually occur and 24 hours is the normal limit for working after annealing, although this can be extended if the material is stored under refrigerated conditions to slow the ageing process. A temperature of -5ºC will provide approximately 2 days’ delay while one of -20ºC will provide approximately 1 week’s delay in the agehardening process The maximum for refrigeration is approximately 150 hours at -20°C. Typical annealing procedure may be achieved by raising the temperature of the alloy to between 340°C and 410C. The alloy is then cooled slowly at about 10C per hour (rates will differ with each particular alloy), until it reaches a pre-determined temperature. At this point it is allowed to cool naturally.

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These, heat-treatable type, alloys must never be installed in an aircraft structure while in the annealed state, since material properties and corrosion resistance will be severely affected. Note: Alloys, in the annealed state, are very prone to corrosion.

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Precipitation Treatment Solution-treated aluminium alloys are comparatively soft, immediately following quenching although, with time, the metal gradually becomes harder and gains strength.

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

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We have already discussed how the natural ageing process can be drastically retarded (allowing the metal to be kept in a soft condition until required for use), by storing the alloys at sub-zero temperatures (refrigeration) for prescribed periods of time. Alternatively, following quenching, by re-heating the metal to a lower temperature than that employed for the solution treatment and allowing it to ‘soak’ at that heat for a period of time, the ageing process (and, thus, the hardening of the alloy) can be accelerated. This process is referred to as artificial ageing or precipitation treatment.

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Identification of Heat-Treated Aluminium Alloys Aluminium alloys that have been subjected to heat-treatment are usually identified by markings that indicate the heat-treatments involved. Three typical identification systems are those of the British Standards Institute (BS), the Ministry of Supply (MoS), and the American systems as can be seen in Table 5.

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Table 5 IDENTIFICATION MARKINGS OF HEAT-TREATED ALUMINIUM ALLOYS BS System M O OD T W WP

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

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MoS System A N W WP

American System T3 T4 T6 T8 T9

Meaning

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

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

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

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Apart from these systems, many other exist world-wide, but the British systems are, broadly, confined to three basic ones for light alloys.  British Standards for general engineering use BS 1470 -1475. In this series the prefix N is used to denote non-heat-treatable aluminium alloys and prefix H for the heat-treatable alloys.  British Standards for aerospace use: BS X LXX. (The "L" series) e.g. BS 3 L72 indicates the 3rd amendment to the basic L 72 spec. LM - indicates a cast material. The wrought materials are commonly abbreviated to L71, L72, L 73 etc. Examples of some of these aircraft BS codes are: a) L159 DURAL*

Solution-Treated - Artificially aged

b) L163 ALCLAD

Solution-Treated - Naturally aged

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*DURAL is, actually, a Trade name for an Al/Cu/Mg/Si/Mn alloy, originally manufactured by the Duren Aluminium Company (Germany), but it tends to be used as a generic name for similar alloys, regardless of source of manufacture.  D.T.D. Specifications: - these are material identification numbers issued by the Directorate of Technical Development (a Ministry Department) for specialised applications. i.e. when widespread use is not anticipated. If such a material finally becomes commonly used, a British Standards specification is compiled and issued.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – NON-FERROUS (DCAM 6.2 L1 & L2)

Hardness Tests The Brinell Hardness System

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The Brinell hardness system is one of the most widely used systems for indicating the hardness of metals and alloys. The Brinell hardness tester uses a hydraulic force to impress a spherical penetractor into the surface of a sample. The amount of force used is approximately 3,000 kilograms of steel and 500 kilograms for non-ferrous metals. The force is hydraulically applied by a hand pump and read on a pressure gauge. When the sample is removed from the tester, the diameter of the impression is measure with a special calibrated microscope. This diameter is converted into a Brinell number by using a chart.

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The Rockwell Hardness System

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The Rockwell hardness tester gives the same information the Brinell tester gives, except that it measures the depth to which the penetrator sinks into the material rather than the diameter of the impression. The Rockwell tester uses three types of penetrators: a) A conical diamond b) 1/16 inch ball c) 1/8 inch ball

There are also three major loads: a) 60 kilograms b) 100 kilograms c) 150 kilograms

The two most commonly used Rockwell scales are: i. ii.

B-scales for soft metals, which uses a 1/16 inch ball penetrator100kg major load C-scales for hard metals, which uses the conical diamond penetrator and a 150kg major load

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – NON-FERROUS (DCAM 6.2 L1 & L2)

Tensile Testing Tensile tests are usually carried out on wire, strip, or machined samples with either circular or rectangular cross section. Test pieces are screwed into or gripped in jaws and stretched by moving the grips apart at a constant rate while measuring the load and grip separation. In the tensile strength test, the material is pulled until it breaks.

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Impact Strength Testing

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Impact strength is measured by allowing a pendulum to strike a grooved machined test piece and measuring the energy absorbed in the break. Impact resistance measures the material toughness.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

6.3 COMPOSITE STRUCTURES Composite structures differ from metallic structures in several ways: i. ii. iii. iv.

Excellent elastic properties Ability to be customized in strength Damage tolerance characteristics Sensitivity to environmental factors

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Composites require a vastly different approach from metals with regards to their design, fabrication and assembly, quality control and maintenance. One main advantage to using a composite over a metal structure is its high strength-to-weight ratio. Weight reduction is a primary objective when designing structures using composite materials.

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Composite strength depends upon the types of fibres and bonding materials used, and how the part is engineered to distribute and withstand specific stresses. In aircraft construction, most currently produced composites consist of a reinforcing material to provide the structural strength, joined with a matrix material to serve as the bonding substance. The three main parts of a fibre-reinforced composite are:

a) The fibre b) The matrix c) The interface or boundary between the individual elements of the composite Fibres

Reinforcing fibres provide the primary structural strength to the composite structure when combined with a matrix. The following are the five most common types of reinforcing fibres used in aircraft construction.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

1. Fibreglass (Glass Cloth) Fibreglass is made from small strands of molten silica glass that are spun together and woven into cloth. Many different weaves of fibreglass are available. The disadvantages of fibreglass are that it weighs more and has less strength than most other composite fibres. Fibreglass is an excellent reinforcing fibre currently used in advance composite application.

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Two most common types of fibreglass are: a) b)

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S-glass E-glass

S-glass is produced from magnesia-aluminasilicate and is used where a very high tensile strength fibreglass is needed.

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E-glass known as “electric glass” because of its high resistivity to current flow, is produced from borosilicate glass and is the common type of fibreglass used for reinforcement. Fibreglass is usually a white gleaming cloth. The widespread availability of fibreglass and its low cost make it one of the most common reinforcing fibres utilised in aircraft non-structural composites.

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2. Aramid ( Kevlar ) Aramid is an organic aromatic-polymide polymer, commercially known as Kevlar. Aramid exhibits high tensile strength, exceptional flexibility, high tensile stiffness, low compressive properties and excellent toughness. The tensile strength of Kevlar composite material is approximately four times greater than alloy aluminium.

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Aramid fibres are non-conductive and produce no galvanic reaction with metals and the advantage is its strength-to-weight ratio, it is very light compared to other composite materials. Aramid-reinforced composite also demonstrate excellent vibration-damping characteristics in addition to a high degree of shelter and fatigue resistance. Aramid is ideal for use in aircraft parts that are subject to high stress and vibration.

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Example: Some advanced helicopter designs have made use of aramid materials to fabricate main rotor blades and hub assemblies. Flexibility of the aramid fabric allows the blade to bend and twist in flight, absorbing much of the stress. Aramid fibre is usually characterised by its yellow colour, and as with most reinforcing fibres.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

3. Carbon/Graphite Carbon fibres are produced in an inert atmosphere by the pyrolysis of organic fibres such as rayon, polyacrylonitrile and pitch. The term carbon is often interchangeable with the term graphite. Carbon fibres are typically carbonised at approximately 2400o F and composed of 93% to 95% carbon.

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Graphite fibres are produced at approximately 3450oF to 5450oF and are more than 99% carbon.

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Advantages to carbon/graphite materials are in their high compressive strength and degree of stiffness. Carbon fibre is cathodic while aluminium and steel are anodic.

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Carbon promotes galvanic corrosion when bonded to aluminium or steel, and special corrosion control techniques are needed to prevent this occurrence. Carbon/graphite materials are kept separate from aluminium components. Americans refer to carbon fibres as “graphite” while Europeans refer to it as carbon fibre. Carbon/graphite is a black fibre that is very strong, stiff and used primarily for rigid strength characteristic. Fibre composites are used to fabricate primary structural components such as the ribs and skin surfaces of the wings.

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4. Boron Boron fibres are made by depositing the element boron onto a thin filament of tungsten. The resulting fibre is approximately 0.004 inches in diameter, has excellent compressive strength and stiffness and extremely hard. Boron is not commonly used in civil aviation because it can be hazardous to work with, and is extremely expensive. Many civil aviation manufacturers are utilizing hybrid composite materials of aramid and carbon/graphite instead of boron.

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5. Ceramic Ceramic fibres are used where a high-temperature application is needed. This form of composite will retain most of its strength and flexibility as temperatures up to 2200oF. Tiles on the Space Shuttle are made of a special ceramic composite that dissipate heat quickly. Some firewalls are also made of ceramic-fibre composites.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

1. Warp The warp of threads in a section of fabric run the length of the fabric as it comes off the roll. Warp direction is designated as 0o. There are typically more threads woven into the warp direction than the fill direction, making it stronger in the warp direction. Warp is critical in fabricating or repairing composites, insertion of another colour or type of thread at periodic intervals identifies the warp direction.

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2. Weft

Weft or fill threads of the fabric are those that run perpendicular 90o to the warp fibres. The weft/fill threads interweave with the warp threads to create the reinforcing cloth. 3. Selvage Edge

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The selvage edge of the fabric is the tightly woven edge parallel to the warp direction, which prevents edge from unravelling. The selvage edge is removed before the fabric is utilised. The weave of the selvage edge is different from the body of the fabric and does not have the same strength characteristics as the rest of the fabric. 4. Bias

o The bias is the fibre orientation that runs at a 45 angle (diagonal) to the warp threads. The bias allows for manipulation of the fabric to form contoured shapes. Fabrics can often be stretched along the bias but seldom along the warp or fill.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Matrix Systems

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The function of the matrix in a composite is to hold the reinforcing fibres in a desired position. It also gives the composite strength and transfers external stresses to the fibres. A wide range of resin systems are used for the matrix position of fibre reinforced composites. Resin is an organic polymer used as a matrix to contain the reinforcing fibres in a composite material. Resin matrix systems are a type of plastic and include two general categories: a) Thermoplastic b) Thermosetting

Thermoplastic and thermosetting resins by themselves do not have sufficient strength for use in structural applications. Thermoplastic resins use heat to form the part into the desired shape. However, this shape is not necessary permanent. If a thermoplastic resin is reheated, it will soften and could easily change shape. Thermosetting resins use heat to form and irreversibly set the shape of the part. Thermosetting plastics, once cured, cannot be reformed even if they are reheated. Most structural airframe applications are constructed with thermosetting resins.

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Polyester Resins

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Polyester resins, an early thermosetting matrix formula, are mainly used with fibreglass composites to create non-structural applications such as fairings, spinners and aircraft trim. Polyester resins give fibreglass cohesiveness and rigidity. Polyester resin/fibreglass composites do not offer sufficient strength to fabricate primary structural members. As the polyester resin shrinks, it produces an increasingly tight grip on the embedded metal.

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Epoxy Resins

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Most of the newer aircraft composite matrix-formulas utilize epoxy resins, which are thermoset plastic resins. Epoxy resin matrices are two-part systems consisting of a resin and catalyst. The catalyst acts as a curing agent by initiating the chemical reaction of the hardening epoxy. Epoxy resin systems are well known for their outstanding adhesion, strength and resistance to moisture and chemicals. Not every type of epoxy resin is suitable for every type structure or repair. Make sure to use the proper resin called for in the manufacturer’s repair manual. Some of the properties of epoxy which make it useful for bonded structures are its low shrinkage percentage, high strength-to-weight ratio and ability to adhere to an almost endless variety of materials. Epoxies may be used in place of polyester resins for almost any application. They also have a long shelf life. Unmixed, epoxies generally kept for almost a o year at 72 F.

Adhesives

Resins come in different forms. Resins used for laminating are generally thinner, to allow proper saturation of the reinforcing fibres. Others are used for bonding and are typically known as adhesive. Adhesive resins and catalysts are available either in pre-mixed quantities or in separate containers.

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Pre-impregnated Materials

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Process of pre-preg

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Pre-preg carbon fibre

Pre-impregnated fabrics, commonly known as “pre-preg”, are fabrics that have the resin system already saturated into the fabric. Fabrics are pre-impregnated with the proper amount and weight of a resin matrix to eliminate the mixing and application details such as proper mix ratios and application procedure. One limitation to pre-impregnated materials is that they must be stored in a freezer to prevent the resin from curing. Pre-preg fabrics cannot be left out of the cold for prolonged periods, and must warmed slightly before use to achieve better workability. A disadvantage associated with pre-impregnated materials is that they are usually purchased in full roll quantities. The roll may exceed its shelf life before used. Although the material may appear to be good condition, it cannot be used for aircraft application once the shelf life has expired.

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Methods of Curing Composite matrix systems cure by chemical reaction. Some matrix systems can cure at room temperature while others require heat to achieve maximum strength. Failure to follow the proper curing requirements and improper use of the curing equipment can cause defects in the repair.

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Room Temperature Cure

Some types of composite repairs may be cured at room temperature 65o – 80oF over a time period of 8 – 24 hours depending on the type of resin used. In some cases, room temperature curing can be accelerated by applying low heat 140o – 160oF. Full cure strength is not usually achieved for five to seven days. Room temperature cures are used on non-structural or lightly loaded parts.

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Heat Curing

Most advanced composites utilised resins that require high temperatures during the curing process in order to develop full strength. The repair of parts that use theses types of resins must also cure at high heat settings 250o – 750oF to restore the original strength. The amount of heat applied must be controlled by monitoring the surface face temperature of the repair, overheating can cause severe damage. Heat curing can be accomplished using several different methods: 1. 2. 3. 4. 5.

Heat lamps are not recommended due to the uncontrolled heating of the part. Heat lamps may localize the heat in one spot causing uneven curing. Heat guns must be controlled with the temperature monitor. Heat guns can produce heat up to 750oF when left on continuously. Excessive heat can evaporate resins, leaving dry areas in the part. Oven curing offers controlled and uniform heating of all repair surfaces. Some evens incorporate vacuum ports to provide pressure while curing. Autoclaves are customarily used in the manufacturing of composites, rather than in repair. Hot patch bonding utilizes a flexible silicon heating-blanket that incorporates a temperature control. This is the preferred method of curing, due to the controlled even heating of the part.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Composite Inspection Composite inspection techniques and non-destructive testing (NDT) methods typically involve the use of multiple methods to accurately determine the airworthiness of the structure. Many metal inspection and NDT methods transfer to composite applications. Composite structures require ongoing inspection intervals along with non-scheduled damage inspection and testing.

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When a composite structure is damaged, it must first be thoroughly inspected to determine the extent of the damage, which often extends beyond the immediate apparent defects. Proper inspection and testing methods help determine the classification of damage which is, whether the part must be replaced. The manufacturers’ structural repair manual outlines inspection procedures damage classification factors and recommended repair methods.

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Some of the more common composite inspection and testing methods are visual inspection, tap testing and ultra sonic testing along with several other more advanced NDT methods.

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Visual Inspection

Visual inspection is the most frequently used inspection method in aviation. Pilots, ground creq and maintenance technicians visually inspect the aircraft on a daily basis. Method of inspection is generally used to detect resin-rich areas, resin starvation, edge delamination, fibre break-out, cracks, blistering and other types of surface irregularities. A strong light and magnifying glass are useful tools for visual inspection. Shining strong light through the structure, called backlighting helps in the identification of cracked or broken fibres, and in some cases delamination. Backlighting does not detect entrapped water. In addition, to properly inspect a composite using the backlight method, you must strip the surface of all paint. Tap Test

The tap test is one of the simplest methods used to detect damages in bonded parts. The laminated part is tapped with a coin or small metallic object, such as a tap hammer to detect delamination. The tap test is an acoustic test, one in which you listen for sound differences in the part and is not the most accurate test method. The tap test detects delaminations close to the surface in addition to transitions to different internal structures. A properly prepared, undamaged laminated area produces a sharp, even pitch as compared to a delaminated area, which produces a dull sound. However changes in the thickness of the part, reinforcements, fasteners and previous repairs may give false reading when using the tap test. Tap testing will not indicate delamination well below the surface in thick parts.

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Ultrasonic Inspection

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An ultrasonic tester is useful for detecting internal damage such as delaminations, core crush and other subsurface defects. Two common methods of ultrasonic testing include the pulse echo and through transmission methods. In the pulse echo method, the test generates ultrasonic pulses, sends them through the part and receives the return echo. The echo patterns are displayed on an oscilloscope. The “through transmission” method uses two transducers. One transducer emits ultrasonic waves through the part and the other receives them.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Radiography

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X-ray View of Same Riveted Sheet of Metal

Radiography or X-ray inspection is used to detect differences in the thickness or physical density when compared to the surrounding material of a composite. Radiography also detects entrapped water inside honeycomb core cells. In addition to detecting the actual defect, it can also detect the extent and size of the damage. Radiography or X-ray inspection will also detect foreign objects in the composite structure if the objects density is different from the composite structure. Composite Repair

The newer advanced composites use stronger fabrics and resin matrices, which cannot be repaired in the same way as fibreglass. To repair an advanced composite structure using the materials and techniques traditionally used for fibreglass repairs may not result in an airworthy repair. Depending on the manufacturer of the aircraft, classification of damage is usually placed in one of the categories: 1. 2. 3.

Negligible damage – may be corrected by a simple procedure with no flight restrictions Repairable damage – damage to the skin, bond or core that cannot be repaired without placing restriction on the aircraft or structure. Non-repairable damage – a composite structure that is damaged beyond limits must be replaced unless a structurally sound repair can be designed by a structural engineer.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

6.3.2 Wooden Structures Introduction The first airplane built by the Wright brothers was made of wood. Wood was used on early aircraft because of its availability and relatively high strength to weight ratio. The cost of the additional hand labour needed for wood construction and maintenance, caused wood aircraft to become almost entirely superseded by those of all metal construction.

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Refer to Figure 20

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 20: Wooden Aircraft

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Wood Wood is not as strong as steel or aluminium, but the construction can be designed that the necessary strength is achieved with corresponding savings in weight.

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Many designers prefer to use wooden spars in acrobatic aircraft because wood will better withstand the bending loads imposed during aerobatics. Unlike metal, wood does not weaken from fatigue.

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Types of Wood

Wood and adhesive materials used in aircraft repair should meet aircraft quality standards and be purchased from reputable distributors to ensure such quality.

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Sitka spruce is the reference wood used for aircraft structures because of its: • • •

uniformity strength excellent shock resistance qualities.

Refer to Figure 21 and Table 9

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 21: Sitka Spruce

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Species of Wood Sitka Spruce Douglas Ft

White Pine

Strength Properties, Compared Remarks to Spruce 100% Excellent for all causes. Considered as standard for this table. Exceeds spruce May be used as substitute for spruce in same size. Difficult to work with hand tools. Gluing satisfactory. 85% - 96% Excellent working qualities and uniform in properties, but somewhat low in hardness and shock-resisting capacity. Gluing satisfactory. Slightly exceeds spruce Less uniform in texture than spruce. May be used as direct substitute for spruce. Gluing satisfactory.

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Table 9: Wood Properties

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Construction of Wooden Airframe Structure Woodworking is a skill that is easily learned by the novice who usually has a basic knowledge of wood construction and some of the necessary tools.

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Figure 22: Wood Structure

Figure 23 : Wing wood Structure

Strong, rigid, light weight truss or framework wooden structures have been in use since the 1920s and are probably the easiest structural type to build. Wood is used in fabricating spars, building ribs, floorboards, instrument panels, wing tip bows, longerons and stringers, leading edges, etc. Wood is easily formed into shapes making it the obvious choice for wing tip bows, leading edges, and wing walkways. The easiest wing to build is rectangular with a constant aerofoil section, constant thickness and constant chord, commonly known as a plank wing.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Types of Defects in Wood Material (Refer to Figure 24) Following are several examples of wood defects: a) Checks A lengthwise separation or crack of the wood that extends along the wood grain. It develops during drying and is commonly caused by differences in radial and tangential shrinkage or because of uneven shrinkage of the tissues in adjacent portions of the wood.

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b) Shakes A separation or crack along the grain, the greater part of which may occur at the common boundary of two rings or within growth rings

c) Heartwood The inner core of a woody stem or log, extending from the pith to the sap, which is usually darker in colour. This part of the wood contains dead cells that no longer participate in the life processes of the tree

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d) Knot That portion of a branch or limb that is embedded in the wood of a tree trunk, or that has been surrounded by subsequent stem growth.

Figure 24: Several Wood Defects

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Aircraft Adhesives / Glues The adhesive used in aircraft structural repair plays a critical role in the overall finished strength of the structure. The maintenance technician must only use those types of adhesives that meet the performance requirements necessary for use in aircraft structures. Not every type of glue is appropriate for use in all aircraft repair situations. Because of its importance, use each type of glue in strict accordance with the aircraft and adhesive manufacturer's instructions.

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Types of Adhesives

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1. Casein Glue

Most older airplanes were glued with casein glue, which was a powdered glue made from milk. Casein glue deteriorates over the years after it is exposed to moisture in the air and to wide variations in temperature. Many of the more modern adhesives are incompatible with casein glue.

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2. Plastic resin glue

This type of glue usually comes in a powdered form. Mix it with water and apply it to one side of the joint. Apply a hardener to the other side of the joint, clamp the two sides together and the adhesive will begin to set. Plastic resin glue rapidly deteriorates in hot, moist and under cyclic stresses, making it obsolete for all aircraft structural repairs. 3. Resorcinol glue

It's a two-part synthetic resin glue consisting of a resin and a hardener and is the most water-resistant of the glues used. The glue is ready for use as soon as the appropriate amount of hardener and resin has been thoroughly mixed. Resorcinol adhesive are one of the most common types of glue used in aircraft wood structure repair. 4. Phenol-formaldehyde glue

It's the most commonly used in the manufacturing of aircraft-grade plywood. This glue requires high curing temperatures and pressures making it impractical for use in the field.

5. Epoxy resins These are two-part synthetic resins that generally consist of a resin and a hardener mixed together in specific quantities. Epoxies have excellent working properties.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Inspection of Wood Structure Most wood damage is caused by conditions such as moisture, temperature, and sunlight. Because wood is an organic material, it is subject to mildew and rot unless protected from moisture. Keep wood airplanes in well ventilated hangars and take special care to ensure that all of the drain and ventilation holes remain open. If a ventilation hole becomes obstructed, changes in air temperature will cause moisture to condense inside the structure, which will cause the wood to deteriorate.

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When inspecting a wood structure aircraft, move it into a dry, well-ventilated hangar. One of the first steps is to check the moisture content of the wood using a moisture meter. If the moisture content is high, dry the wood structures before inspecting further.

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Wooden structures of the aircraft need to be dry to be able to effectively determine the condition of the bonded joints. The following are several inspection methods and associated equipment employed for inspecting wooden structures. 1. Moisture Metering Use to determine the moisture content of wood structure. Wood that is too wet or too dry may compromise the strength and integrity of the structure. A moisture meter reads the moisture content through a probe that is inserted into a wooden member. When water is ingress in wood it is recognizes by grayish stain 2. Tapping

The wood structure may be inspected for structural integrity by tapping the suspect area with a light plastic hammer or screwdriver handle. Tapping should produce a sharp, solid noise from a solid piece of wood. If the wood area sounds hollow or feels soft, inspect further. 3. Probing

If soft, hollow wood is found during the tap test, probe the suspect area with a sharp metal tool to determine whether the wood is solid. Ideally, the wood structure should feel firm and solid when probed.

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4. Prying Use prying to determine whether a bonded joint shows signs of separation. Light prying is sufficient to check the integrity of a joint. If there is any movement between the wood members of the joint, a failure of the bond is confirmed.

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5. Smelling

Smell is a good indicator of musty or moldy areas. When removing the inspection panels, be aware of any odors that may indicate damage to the wood structure. Odor is an essential indicator of possible wood deterioration. Musty and moldy odors reveal the existence of moisture and possible wood rot.

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6. Visual Inspection

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Visual inspection techniques are used to determine any visible signs of damage. Both internal and external visual examinations are imperative to a complete inspection of the wood structure. Repair of Wooden Structure

The basic criterion for any aircraft repair is that the repaired structure must not only be as strong as the original structure, but the rigidity of the structure and the aerodynamic shape must also be equivalent. Materials used for the repair of a wooden structure should be the same as the original unless they have become obsolete. If substitutions are made, they must produce a repair that meets the basic requirements of the manufacturer and the authority.

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Plywood Skin Repair Aircraft that incorporate plywood skins normally carry a large amount of stress from the flight load. Therefore, repairs to plywood skins are made in strict accordance with the recommendations of the aircraft manufacturer. There are several types of plywood patches repair

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a) Splayed Patch (Refer to Figure 25)

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Small holes in thin plywood skin may be repaired by a splayed patch. This type of patch is used if the skin is less than or equal to 1/10 inch thick and the hole can be cleaned out to a diameter of less than 15 thickness (15T)

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 25: Splayed patch

b) Surface Patch c) Plug Patch d) Scarfed Patch

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

Fabric Orientation When working with composite fibres, it is important to understand the construction and orientation of the fabric because all design, manufacturing and repair work begins with the orientation of the fabric. Some of the terms used to describe fibre orientation are: 1. 2. 3. 4.

Warp Weft / Fill Selvage edge Bias

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Fabric Covering It is important to understand the construction and orientation of fabric material because all design, manufacturing and repair work begins with the orientation of the fabric. Fabric structure relies on the proper placement and use of the reinforcing fabric to produce a strong covering Some of the terms used to describe fabric orientation are warp, weft / fill, selvage edge and bias.

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1. Warp

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The warp of threads in a section of fabric run the length of the fabric as it comes off the roll. Warp direction is designated as 0º. There are typically more threads woven into the warp direction than the fill direction, making it stronger in the warp direction. The warp is critical in creating or repairing fabric coverings. The fabric must be applied with the warp parallel to the direction of flight.

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2. Weft / Fill

The Weft / Fill threads of the fabric are those that run perpendicular 90 º to the warp fibres. The weft threads interweave with the warp threads to create the reinforcing cloth. 3. Selvage Edge

The selvage Edge of the fabric is the tightly woven edge parallel to the warp direction, which prevents edge from unravelling. The selvage edge is removed before the fabric is utilized. The weave of the selvage edge is different from the body of the fabric and does not have the same strength characteristic as the rest of the fabric. 4. Bias

The Bias is the fibre orientation that runs at a 45º angle (diagonal) to the warp threads. The bias allows for manipulation of the fabric to form contoured shapes. Fabric can often be stretched along the bias, seldom along the warp or fill.

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Finishing Materials Several finish materials that increase the durability and appearance of fabric are used in covering processes. These items provide additional rigidity of the fabric, which helps to transfer the aerodynamic lift provided by the covering into the structure of the aircraft. Inspection hole and drainage grommets, as well as tapes and lacing cords, are vital components to a quality fabric-covered structure.

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Reinforcing Tape

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Reinforcing tape is a flat woven cotton material that is available in ¼ - inch, 3/8 - inch and ½ -inch widths, with a strength of 150 pounds per half-inch of width. This tape is used under rib-lacing cord or other fabric-attaching devices from pulling through the fabric covering. Reinforcing tapes made from polyester are also available in the same widths as cotton tape. The polyester is less susceptible to deccus from moisture and mildew and has more strength than the fibre tape. Surface Tape

Surface tape is made of the same material as the covering fabric and is used over all seams, ribs, around corners, along the trailing edges, around the tips and along the trailing edge of all surfaces. The purpose of the tape is to blend the covering around contours and irregularities to make for smoother surface finish. In addition, the tape aids to prevent the airstreams during flight. Surface tapes are available in a bias cut or straight cut. Straight cut tape has a weave that runs parallel to its edges and is primarily used over flat surfaces such as on top of wing ribs. Bias cut tapes are constructed so that the weave of the fabric runs at a 45 º angle to the edge. The bias weave provides for better contouring around curves such as those found on the rudder or wing tip bows.

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Machine-Sewing Threads

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YARN

Machine sewing threads are used primarily to sew lengths of fabric to form large blankets or to form an envelope to slip over wing or other surface. These threads are available in grade–A cotton or polyester. Hand-Sewing Thread

Cotton thread is used for hand sewing stitches. This thread is generally supplied without any coating but should be lightly waxed beeswax before being used. Polyester hand sewing thread is commonly uncoated with multiple plies and has a tensile strength of over 15 pounds. Drainage Grommets and Inspection Rings

Drainage grommets are small doughnut-shaped plastic, aluminium or brass rings that are installed in numerous locations on the aircraft. Typical installation position include the lowest point on the bottom of the wing and fail surfaces, toward the rear of each rib bay and on the fuselage fabric at the lowest point of each compartment. Grommets are usually installed when the second coat of dope is applied to fabric, while the dope is still wet. When all the finishing coats have cured, the centre of the grommet is cut out with a sharp knife blade to allow any moisture within the structure to drain out and to ventilate the inside of the structure to minimize condensation. Larger inspection hole grommets, or inspection rings are installed on the fabric over any location where access to the interior structure may be needed.

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Organic Fabric Materials

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A common organic fabric covering material is grade-A cotton. This material meets the Aeronautical Material Specification.

The cloth has between 80 and 84 threads per inch in both the warp (the direction along the length of the fabric) and the weft directions (the direction along the width) and weighs about four ounces per square yard. In the process of manufacturing grade- A cotton fabric, the natural material is mercerized by dipping the threads in a hot caustic soda solution to give them sheen and to increase their strength. Cotton intermediate-grade fabric has a much finer weave than grade A cotton, with up to 94 threads per inch allowed in both the warp and fill directions . A very fine-weave cotton fabric called glider fabric has up to 110 threads per inch in warp and fill and meets specification. This fabric is designed for use on glider and sailplanes. Irish linen, produced in the British Isles, is another organic fabric. Since this fabric was originally milled by the British, it was designed to meet British specification. This fabric is stronger than grade-A cotton, with strength.

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Inorganic Fabric Materials

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Man-made inorganic fabric that is produced from synthetic polyester has quickly become one of the most popular aircraft covering materials. Polyester fibres, woven into cloth with different weights, are sold under trade names such as Ceconite, Polyfiber and Superflite. The fibres used to make the material have been passed through rollers and are woven so that the number of fibres in the warp direction is equal to the number in the fill direction. When the material is finished it is delivered in an unshrunk condition. Once heat is applied during the installation process, the unshrunk fabric will constrict back to its original length and size. Other inorganic fibre-covering systems use fibreglass filaments woven into cloth which will not decay with moisture or mildew and has virtually unlimited life. Fibreglass cloth has previously been approved as reinforcement over in sound condition, but treated fibreglass has become an approved direct replacement for grade-A cotton.

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Determining Fabric Strength One of the most important duties that an aircraft technician must perform is to determine the airworthiness of the fabric covering by checking its strength. Due to the expense and time involved, most aircraft are re-covered only when the strength of the fabric drops below the minimum airworthy value. The strength of the fabric is a major factor in the airworthiness of an airplane, its condition is determined during each 100-hour, annual or other required airworthiness inspection.

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There are a variety of methods available to determine fabric-covering strength. Some of these methods can be done in the field using a simple testing apparatus.

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Seyboth Tester

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To determine fabric strength, a Seyboth tester is often used by maintenance technicians working in the field. These testers are sometimes called a “punch tester” because of their method of operation. These tools provide a direct indication of the strength of the fabric. This instrument, a spring-loaded housing, holds a shaft, which has a flared point at one end with a hardened steel tip in its centre. The opposite end of the shaft is marked with red, yellow and green bands. When pressure is applied to the tip, the bands become exposed at the top of the housing. To use the tester, hold it vertically over the covering surface and press straight into it until the tip penetrates the fabric. The point on the instrument must break the fabric and enter far enough to allow the shaft face to make full contact. A small amount of pressure moves the red band out of the housing to indicate that the fabric is weak. The yellow band indicates that the fabric is stronger, and the green band indicates the condition of good quality fabric. After the test is complete, cover the hole in the fabric with a small circular patch. Since the Seyboth tester punches holes in the fabric during each test, another type of tester that does not leave holes is the Maule tester.

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Maule Test Instrument

A Maule tester is similar to the Seyboth tester in that it measures the amount of pressure applied directly to the fabric. This tester consists of a tubular housing containing a calibrated spring. When pressed against the fabric, pressure is measured and indicated on a scale. If fabric fails, the Maule tester penetrates the fabric prior to reaching the specified point on the strength scale. It the fabric has adequate strength, the tester will not penetrate the fabric and a repair is not required.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

REPAIRS TO FABRIC COVERINGS If the fabric has been damaged extensively, it is usually impractical and uneconomical to make satisfactory repairs by sewing and patching. The extent and location of damage to the fabric that may be repaired will be detailed in the repair section of the aircraft manual concerned, but extensive damage is often made good by replacing complete fabric panels. However, the replacement of large fabric panels, particularly on one side of a component, may lead to distortion of the structure and it may be advisable to completely re--cover the component.

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Before attempting any repair to the fabric covering, the cause of the damage should be ascertained. The internal structure should be inspected for loose objects such as stones, remains of birds, insects, etc, and any structural damage made good. Using thinners, all dope should be removed from the fabric surrounding the damaged area before any stitching is carried out, since doped fabric will tear if any tension is applied to the repair stitches.

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If the fabric has been damaged extensively, it is usually impractical and uneconomical to make satisfactory repairs by sewing and patching. The extent and location of damage to the fabric that may be repaired will be detailed in the repair section of the aircraft manual concerned, but extensive damage is often made good by replacing complete fabric panels. However, the replacement of large fabric panels, particularly on one side of a component, may lead to distortion of the structure and it may be advisable to completely re--cover the component. Before attempting any repair to the fabric covering, the cause of the damage should be ascertained. The internal structure should be inspected for loose objects such as stones, remains of birds, insects, etc, and any structural damage made good. Using thinners, all dope should be removed from the fabric surrounding the damaged area before any stitching is carried out, since doped fabric will tear if any tension is applied to the repair stitches. 

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Herring--Bone Stitch. The herring--bone stitch (also known as the ’ladder’ or ’baseball’ stitch) should be used for repairing straight cuts or tears which have sound edges. The stitches should be made as shown opposite, with a lock knot every 150 mm (6 in). o

There should be a minimum of two stitches to the centimetre (four stitches to the inch) and the stitches should be 6 mm (0.25 in) from the edge of the cut or tear.

o

After the stitching has been completed, 25 mm (1 in) wide serrated tape should be doped over the stitching. A square or rectangular fabric patch should then be doped over the whole repair, ensuring that the edges of the patch are parallel to the warp and weft of the fabric covering and that they overlap the repair by 37 mm (1.5 in). The original doping scheme should then be restored.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Herring-bone (Baseball) Stitch

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Repairs with Woods Frames. On some aircraft, repairs to cuts and tears with jagged edges, which cannot be stitched as described in the previous paragraphs, can be repaired by using the Woods frame method described for inspection panels previously. Repairs of up to 50 mm (2 in) square may be made, provided they are clear of seams or attachments by a distance of not less than 50 mm (2 in). The affected area should be cleaned with thinners or acetone and repaired in the following manner. o The Woods frame should be doped into position surrounding the damaged fabric and, if the frame is of the square type, the edges should be parallel to the weft and warp of the covering. When the dope has dried, the damaged portion of the fabric should be cut out and the aperture covered by a fabric patch. o If Woods frames are not readily available they can be made from cellulose sheet 0.8 mm (0.030 in) thick with minimum frame width of 25 mm (1 in); in the case of the square type of frame the minimum comer radii should be 12 mm (0.5 in). In some special cases, aircraft manufacturers use 2 mm plywood complying with British Standard V3 for the manufacture of the frames, in which case it is important to chamfer the outer edges of the frame to blend with the aerofoil contour.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Repair of L-Shaped Tear

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2) 

Repair by Darning. Irregular holes or jagged tears in fabric may be repaired by darning provided the hole is not more than 50 mm (2 in) wide at any point. The stitches should follow the lines of the warp and weft and should be closely spaced as shown in the adjacent figure. The whole repair should be covered with a serrated fabric patch in the usual way, with an overlap of 37 mm (1.5 in) from the start of the dam.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Repair by Darning

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2) 

Repair by Insertion. For damage over 100 nun (4 in) square, insertion repairs are generally used, either of the two methods described below being suitable. o Normal Insertion Repair. The damaged area of the fabric should be cut out to form a square or rectangular hole with the edges parallel to the weft and warp. The comers of the hole should then be cut diagonally, to allow a 12 mm (0.5 in) wide edge to be folded under the fabric and this should be held in position with tacking or hemming stitches. o The patch should be made 25 mm (1 in) larger than the cut--out area and its edges should be folded under for 12 mm (0.5 in) and tacked in position in a manner similar to that described in the previous paragraph. In this condition the size of the insertion patch should be similar to, or slightly smaller than, that of the cut--out area. o The insertion patch should be held in position inside the cut--out area with a few tacking stitches and then sewn in position using a herring-bone stitch of not less than two stitches to the centimetre (four stitches to the inch), as shown in the figure opposite. A 25 mm (1 in) wide tape should then be doped over the seams. o For small repairs a square or rectangular cover patch, with frayed or serrated edges, should be doped in position ensuring that the patch overlaps the edge of the tape by 31 mm (1.25 in). Where the size of the insertion patch is more than 225 mm (9 in) square, a 75 mm (3 in) wide fabric serrated tape is often used; the tape should be mitred at the corners and doped in position. The original finish should then be restored.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Normal Insertion Repair

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o

Alternative Insertion Repair. An alternative repair is shown opposite. This consists of cutting away the damaged fabric as described previously but, in this case, the edges of the aperture as well as the edges of the insertion patch are turned upwards. The insertion patch is attached to the fabric cover by stitching along the folded--up edges as near to the contour of the component as practicable (i.e. about 1 mm (0.0625 in) above the surface) using the boot stitch described previously (Stage 1 of the figure opposite). The edges are then doped down (Stage 2 of the figure opposite) and the repair covered with a doped--on fabric patch.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Alternative Insertion Repair

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o

Boot Stitch. A single, well--waxed No.18 linen thread to BS F34 should be used for the boot stitch. The stitches should be made as shown in the figure adjacent and the ends of both threads tied together in a lock knot every 150 mm (6 in), and at the end of a seam.

Boot Stitch

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DOPING INTRODUCTION Fabric has been used from the early days of the aeroplane as a covering for fuselages and aerofoils. It still continues to provide good service for light aircraft but must be protected from deterioration by the application of a dope film. Natural fabrics, such as cotton or linen, deteriorate in use as a result of the effects of sunlight, mildew and atmospheric pollution. Man--made fibres resist some of these agents better than natural fabrics but still require protection. The dope film then achieves the following functions:   

Tautening of natural fabrics Waterproofing Airproofing Lightproofing.

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MATERIALS

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The basic film consists of dope but other materials are used in its application, as described in the following paragraphs. 

Dopes. Dope consists of a number of resins dissolved in a solvent to permit application by brush or spray. This formulation is then modified with plasticisers and pigments to add flexibility and the required colour (see opposite). There are two types of dope in use, namely, cellulose nitrate and cellulose acetate butyrate. The former is usually known simply as nitrate dope and the latter as butyrate or CAB dope. The main difference between the two types of dope is the film base. In nitrate dope a special cotton is dissolved in nitric acid, whilst in butyrate dope cellulose fibres are dissolved in acetic acid and mixed with butyl alcohols. The plasticisers in the two dopes are also different, as are the resin balance and solvent balances. Dope must be stored under suitable conditions, and has a tendency to become acid with age; if old dope is used for refinishing an aircraft it will quickly rot the fabric. Only fresh dope should be used, preferably buying it for the job in hand.



Dope--Proof Paints. Due to the nature of the solvents used in dope, many paints will be attacked and softened by it. Dope--proof paint is therefore used to coat structure which will be in contact with the doped fabric. In the case of wooden structure, spar varnish provides a good dope--resistant finish and an epoxy primer is suitable for metal structures.



Aluminium Dope. To make the fabric lightproof and so prevent damage from ultra--violet radiation, an aluminium dope is used. This is usually supplied ready mixed but can be prepared by mixing aluminium paste or powder in clear dope but it is essential that the materials are obtained from an approved supplier and mixed in accordance with the manufacturer’s instructions.

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Thinners. Dopes are formulated in such a way that the solid constituents are suspended in the appropriate solvents. It will normally be necessary to thin or reduce the dope to make it suitable for spraying. It is important that only the thinners recommended by the manufacturer of the dope is used. The amount of thinners is determined from the manufacturer’s recommendations and is modified by experience to take account of the equipment used and the atmospheric conditions.

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The viscosity can be measured by using a viscosity cup which contains a small hole in the bottom. In use, the cup is dipped into the dope and the flow of fluid is timed from when the cup is lifted from the container to the first break in the flow. In this way subsequent batches of dope can be mixed to exactly the same viscosity as the first batch. It is important that nitrate and butyrate dopes are mixed only with their own specialised thinners. A retarder, or anti--blush thinners, is a special type of thinners with slow-- drying solvents. By drying more slowly they prevent the temperature drop and consequent moisture condensation that cause blushing in a dope finish. In use, the retarder replaces some of the standard thinners and can be used in a ratio of up to one part retarder to four parts of thinners. The use of more retarder than this is unlikely to achieve the desired result. 

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Cleaning Agent. Methyl--ethyl--ketone (MEK) is an important, relatively low cost, solvent similar to acetone. It is widely used as a cleaning agent to remove wax and din and to prepare surfaces for painting or re--doping. It is also useful as a solvent for cleaning spray guns and other equipment.

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Fungicides. Since natural fabrics can be attacked by various forms of mildew and fungus, it may be necessary to provide protection for cottons and linens when doping. This is achieved by having a fungicide added to the first coat of dope. The dope is usually supplied ready mixed but can be prepared by using a fungicidal paste obtained from an approved supplier. If the latter course is necessary, the fungicidal paste should be mixed with the clear dope in accordance with the manufacturer’s instructions; all fungicides are poisonous and therefore, standard precautions should be taken to prevent any ill effects. Since mildew or mould form on the inside of the fabric, it is important to ensure that this first coat of dope completely penetrates the fabric.



Tack Rags. A tack rag is a rag slightly dampened with thinners and is used to wipe a surface after it has been sanded to prepare it for the application of the next coat. Proprietary cloths are also available.



Sandpaper. Sanding is carried out using wet--or--dry paper. This is a waterproof sandpaper that will remain flexible and not clog. The grades most likely to be used are 280, 360 and 600, the last mentioned being the finest grade.



Drainage Eyelets and Inspection Rings. Openings in the fabric cover for drain holes and inspection panels are always reinforced with eyelets or grommets and inspection rings. These are made from cellulose nitrate sheet and are doped into position.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

SAFETY PRECAUTIONS The storage and use of dopes is covered by various Government regulations made under the Factories Act. 

The hazard with the use of dopes comes about because of the flammability of the solvents that are used. The solvents have a low flash point and the vapour produced is heavier than air. Accumulations of vapour are readily ignited producing a serious fire which can spread very rapidly.



One of the most common causes of ignition is a spark produced by the discharge of static electricity. For example, during the course of doping, the fumes from the solvents will accumulate inside the structure. When the dope has dried, subsequent dry sanding and dusting will build up a static charge on the surface. If the operator is wearing rubber soled shoes he will be at the same electrical potential as the surface and nothing will happen. Should the charge on the operator now be lost through his touching some metal part of the spray shop, for example and he then touches some metal part of the structure being doped the static charge will jump to earth creating a spark and igniting the fumes. The best way to prevent this type of problem is to eliminate the static charge altogether by grounding the structure being doped. A wire connected from the structure to a clean metal part of the spray shop will do the job satisfactorily. Clothing that is made of synthetic fibres will build up a static charge more readily than that made from cotton. Leather soled shoes will allow any static charge to be dissipated to ground. When spraying nitrate dope ensure that the spray gun, the operator and the structure being doped are all grounded together.

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

The standard of housekeeping in the spray shop is an important aspect of safety. If the floor becomes contaminated with dried nitrate dope overspray, subsequent sweeping will produce a static charge with the attendant risk of ignition and possible explosion. To clean the floor, it should be doused well with water and then swept whilst it is still wet. Since dopes will not be the only materials used in a spray shop, it should be noted that spontaneous combustion can be the result of a mixing of dope and zinc chromate oversprays.



The fumes created during the spraying process are hazardous to health as well as being a fire risk. Proper operator protection must be provided as recommended in the dope manufacturer’s technical literature. At the first sign of any irritation of the skin or eyes, difficulty in breathing or a dry cough, the operator should stop work and seek medical advice.



Electrical equipment to be used in the spray shop must be of such a nature that it cannot ignite the vapours that will be present. Lead Lamps must be of the explosion--proof variety and dopes must not be mixed using stirrers driven by portable electric drills.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

WORKING CONDITIONS 

In order to accomplish a proper dope job, it is important to control both the temperature and humidity of the air in the spray shop. In addition to this it is necessary to maintain sufficient air flow through the shop to remove the heavy vapours caused by atomisation and evaporation of the solvents used.



To maintain a suitable air flow through the spray shop it is necessary to install a fan at floor level since the vapours produced are heavier than air. The fan must be explosion proof, as must be all other electrical equipment installed in the area. The rate of air flow is dictated by the size of the spray shop and is the subject of various Government regulations. The discharge of the vapours may also be the subject of further requirements and the advice of the Factory Inspectorate should be sought. The air inlet to the spray shop should preferably be in an adjoining room, or at least behind a suitable baffle, in order to reduce draughts to a minimum. If the inlet is in a separate room then the air temperature can be raised to that required before entering the spray shop.

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

Many problems associated with doping can be traced to incorrect temperatures of the air or the dope. If the dope has been left overnight in a cold place then it will take many hours to bring it to the room temperature. Overnight heating of the spray shop is the most satisfactory method to prepare for doping since it usually results in more uniform temperatures throughout the shop. Rapid heating tends to result in stratified heating with the ceiling being considerably hotter than the floor level. Air temperature should be maintained between approximately 21° and 26°C (70° to 79°F) for best results. If the temperature is too low the rapid evaporation of the solvents will lower the temperature of the surface to the point where moisture will condense and be trapped in the finish. Too high a temperature causes very rapid drying of the dope which can result in pin holes and blisters. The only satisfactory way to operate is to constantly monitor and control the air temperature as necessary.



In addition to the proper control of air temperature, the humidity of the air must also be controlled. The desirable range of air humidity is 45 to 50. Satisfactory work can be produced with air humidity as high as 70 or as low as 20, depending upon other variables such as temperature and air flow, but the control of the dope application at extremes is always more difficult.



Humidity should be measured with a hygrometer and although direct reading instruments are available, the wet and dry bulb type is still the most common. In this instrument two thermometers are mounted side by side, the bulb of one being kept wet by water evaporating through a wick. To take a reading of humidity, both thermometers should be read and the difference between them noted; the wet bulb thermometer will be lower. After finding the dry bulb reading in the table opposite, a reading should be taken across to the column headed with the depression of the wet bulb. The relative humidity as a percentage is given at the intersection of the two lines. Example. Assuming a dry bulb reading of 17°C and a wet bulb reading of 14°C, the depression of the wet bulb, that is the amount by which the reading of the wet bulb is reduced below that of the dry bulb, is 3°C. Reading across from 17°C in the dry bulb column to the depression column headed 3°C indicates a relative humidity of 72.

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In order to produce a satisfactory dope film, it is vitally important that all brushes, spray equipment and containers should be scrupulously clean. It is important that oil and water traps in the air lines are properly cleaned and that air reservoirs are drained of accumulated moisture. Pressure pots and spray guns should be thoroughly cleaned with thinners before the dope hardens. If passages have become obstructed with dried dope, the equipment should be dismantled and the parts soaked in MEK or a similar solvent. Packings and seals should never be soaked in solvents or they will harden and become useless.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

PREPARATION PRIOR TO DOPING Before the component is moved into the spray shop, normal housekeeping tasks should be carried out. All dirt, dust and dried overspray should be removed, bearing in mind the safety precautions stated previously. Then the working conditions of temperature and humidity should be achieved with the dope and other materials being brought to the correct temperature. 

o

The structure has been painted with dope--proof paint where required

o

Correct and secure attachment of the fabric to the structure

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

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An inspection should be made of the fabric--covered component to verify the following points:-

o

Correct allowance for tautening of the cover where this is of a natural fabric such as cotton or linen. If the cover is too slack, no amount of doping will rectify this. If it is too tight, a lightweight structure, such as a control surface, could easily be distorted

o

All dust has been removed from the fabric

o

The fabric has reached the temperature of the air in the spray shop.

o

Plastics components, such as windows and windscreens, are adequately protected against solvent attack; newspaper is not satisfactory for this purpose.

With the dope at the correct temperature, it should be mixed and then thinned to a suitable consistency for brush or spray application as appropriate. Whilst the dope is in storage the solid materials tend to settle and the purpose of mixing is to bring these materials back into suspension. To mix any dope satisfactorily, half the contents of the tin should be poured into a clean tin of the same size. The remaining material should be stirred until all the solid material is in suspension, paying particular attention to the bottom of the tin. The contents of the first tin should then be poured into the second tin and a check made that all pigment has been loosened from the bottom. Finally, the dope from one tin should be poured into the other and back again, until it is thoroughly mixed.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

APPLICATION TO NATURAL FABRIC The best looking and most durable film is produced by using multiple coats of a dope that is low in solids. A large number of thin coats, however, requires a great deal of time and modern dope schemes tend to use fewer, but thicker, coats than the earlier schemes. The dope scheme is a schedule listing the number and order of coats of each type of dope. Typical examples of schemes detailed in British Standard BS X26 are given in the tables here and opposite. The standard aircraft doping scheme is 752, but 751 is used on light structures that would be distorted by overtautening and 753 is used where an extra taut cover is required. 

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Priming Coats. This name is given to the first coats applied to the raw fabric. The first coat of dope provides the foundation for all the subsequent coats and as such its mechanical attachment to the fabric is very important. This mechanical attachment is formed by the dope encapsulating the fibres of the fabric. Nitrate dope has much better properties with regard to encapsulating the fibres and is therefore preferred for the first coat. The dope should be thinned by 25 to 50 and applied by brush. The dope should be worked into the fabric to ensure adequate penetration, but not to the point where it drips through to the opposite surface. Since organic fabrics are subject to attack by mildew, a fungicide should be added to the dope used for this first coat.

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When applying the first coat of dope to the wings, the entire wing should first be doped on both sides aft of the front spar. The dope should be allowed to shrink the fabric before doping the leading edge. In this way the fabric will tauten evenly and adjust itself over the leading edge cap without forming wrinkles.



After the dope has dried for a minimum of 1 hour, the tapes, drainage eyelets or grommets and inspection panel rings may be applied. A heavy coat of nitrate dope should be brushed on where required and the tape laid into it, working it down to the surface and rubbing out any air pockets as the tape is laid. A further coat of clear dope is brushed over the top of the tapes. Drainage eyelets or grommets and inspection rings are attached in a similar fashion at this time. To ensure the best adhesion, eyelets or grommets and rings may be soaked in dope thinners for no more than two minutes to soften them. Inspection rings are best reinforced with a circular pinked-- edge patch, a little larger than the ring, doped over the top. The holes in eyelets or grommets and rings are opened with a sharp, pointed knife after doping is complete. The taping is followed by another coat of clear dope which may be butyrate and may be applied by spray gun.



Filling Coats. When the first butyrate coat has fully dried, the fabric will feel rough due to the short fibre ends (the nap) standing up. This nap should be very lightly sanded off, using dry sandpaper, to leave a smooth finish. The surface should then be rinsed clean with water and dried thoroughly. Two full wet cross--coats of butyrate dope should now follow; a cross--coat is a coat of dope sprayed on in one direction and then covered with a second coat at right angles to it before the first coat dries. These in turn should be followed with one good cross--coat of aluminium dope after lightly sanding the clear dope to encourage adhesion. The aluminium coat is in its turn lightly wet sanded to produce a smooth surface and the residue rinsed off with water. Once the aluminium coat has dried, it should be checked for continuity by shining a light inside the structure. The film should be completely lightproof.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2) 

Finishing Coats. The finishing coats of pigmented butyrate dope may now be sprayed on. The number of coats will be determined as a balance between quality and cost but should not be less than three. A high gloss finish is obtained by lightly sanding each coat when dry and spraying multiple thin coats rather than several thick coats. The use of a retarder in the colour coats will allow the dope to flow out and form a smoother film. The final coat should be allowed to dry for at least a month before it is polished with rubbing compound and then waxed. The surface should be waxed at least once a year with a hard wax to reduce the possibility of oxidation of the finish.

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NOTE: A tolerance of +/- 20% is permissable on any of the weights given in these tables.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

APPLICATION TO POLYESTER-FIBRE FABRIC Polyester--fibre fabrics are being increasingly widely used for covering aircraft because of their long life and resistance to deterioration. For this reason it is extremely important that the dope film is of the highest quality so that its life will match that of the fabric. 

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Priming Coats. Tautening of the fabric cover is not a function of the dope film where synthetic fabrics are used, although all dopes will tauten to some extent. Polyester--fibre fabrics are heat shrunk when the structure is covered. The most notable difference in doping a synthetic cover is the difficulty, when compared with natural fabrics, of obtaining a good mechanical bond between the dope and the fibres of the material. Unlike natural fibres the polyester filaments are not wet by the dope and the security of attachment depends upon them being totally encapsulated by the first coat of dope. The first coat must be nitrate dope thinned in the ratio of two or three parts of dope to one part of thinners. This coat is then brushed into the fabric in order to completely encapsulate every fibre. The dope should form a wet film on the inside of the cover but it should not be so wet that it drips through to the opposite side of the structure. The initial coat should be followed by two more brush coats of nitrate dope thinned to an easy brushing consistency.

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Certain additives are approved by the material manufacturer for use with the first coat for improving adhesion to the fabric. However, since polyester is not organic, there is no need for a fungicide to be added to the first coat of dope. 

Filling Coats. Taping and attaching of drainage eyelets or grommets and inspection rings follows the same procedure as for natural fabrics. The priming coats should be followed by spraying two full--bodied cross--coats of clear butyrate dope. After these coats have completely dried they should be lightly sanded (400 grit) and cleaned thoroughly with a tack rag. One full cross--coat of aluminium dope should then be sprayed on and lightly wet sanded when dry, the residue being rinsed off with water. This coat should be tested to verify that it is lightproof by shining a light inside the structure.



Finishing Coats. The finishing coats should now be applied in the same manner as for natural fabrics. It should be noted that with a properly finished polyester cover the weave of the fabric will still show through the dope film. Because the fibres are continually moving, any attempt to completely hide them will result in a finish that does not have sufficient flexibility to resist cracking.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

APPLICATION TO GLASS-FIBRE FABRIC Glass--fibre fabric has a loose weave which tends to make it difficult to apply to aircraft structures. To overcome this problem it is pre--treated with butyrate dope and the covering and doping must be carried out in accordance with the manufacturer’s installation instructions. 

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Priming Coats. Nitrate dope must not be used under any circumstances with this type of fabric. The first coat of clear butyrate dope is sprayed on with the dope being thinned only enough to permit proper atomisation. The atomising pressure must be set to the lowest possible that will permit proper atomisation without the dope being blown through the fabric. The coat should be heavy enough to thoroughly wet the fabric and soften the dope in the fabric, but must not be so heavy that it causes the dope to run on the reverse side of the fabric. If the dope is allowed to run in this way an orange peel finish will develop and the fabric will not tauten properly.

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After the first coat has dried, further coats of butyrate dope should be sprayed on, each a little heavier than the one before it, until the weave fills and the fabric tautens; this may take as many as five coats. Tapes, drainage eyelets or grommets and inspection rings are applied in a coat of butyrate dope.

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

Filling Coats. Once the fabric is taut and the weave has been filled, two full--bodied brush coats of clear butyrate dope should be applied and allowed to dry thoroughly. The film should then be very carefully sanded, making sure that it is not sanded through to the fabric. Whilst the fabric is not damaged by ultra--violet radiation, the clear dope can deteriorate as a result of exposure and therefore, a coat of aluminium dope should be sprayed on for protection and lightly wet--sanded smooth. After the aluminium dope has been sanded, the residue should be removed by washing with water and then the surface thoroughly dried.



Finishing Coats. The application of the finishing coats is carried out in the same manner as for natural fabrics. Several thin, wet coats of coloured butyrate dope will allow the surface to flow out to a glossy finish.

DOPING PROBLEMS

The production of a doped finish that is both sound and attractive is dependent upon a great deal of care and attention being paid to detail at each stage of the finishing process. In spite of this, problems do occur and the following paragraphs detail some common ones and their possible causes. 

Adhesion. There are two basic areas in which adhesion may be poor; between the fabric and the first coat of dope and between the aluminium coat and subsequent coats. Adhesion to the fabric, particularly polyester fabric, is largely dependent upon the technique used to ensure the encapsulation of the fibres. Adhesion to the aluminium coat may be impaired if too much aluminium powder was used or if the surface was not thoroughly cleaned after sanding. The use of a tack rag to finally clean a surface before applying the next coat is always recommended.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2) 

Blushing is a white or greyish cast that forms on a doped surface. If the humidity of the air is too high, or if the solvents evaporate too quickly, the temperature of the surface drops below the dew--point of the air and moisture condenses on the surface. This water causes the nitrocellulose to precipitate out. Moisture in the spray system or on the surface can also cause blushing. Blushing can be controlled by reducing the humidity in the air (raising the temperature by several degrees may achieve this) or by using a retarder in the place of some of the thinners. A blushed area can be salvaged by spraying another coat over the area using a retarder instead of some of the thinners; the solvents attack the surface and cause it to flow out.



Bubbles or Blisters are caused by the surface of the dope drying before all the solvents have had time to evaporate. This may happen if a heavy coat of dope is applied over a previous coat that had not fully dried.



Dull Finish. The gloss of butyrate dope may be improved by the addition of up to 20 retarder in the last coat. Excessive dullness may be caused by holding the spray gun too far from the surface so that the dope settles as a semi--dry mist. Small dull spots may be due to a porous surface under the area.



Fisheyes. These are isolated areas which have not dried due to contamination of the surface with oil, wax or a silicone product. Cleanliness is important, especially when refinishing a repair. All wax should be removed using a suitable solvent before attempting to re--dope the surface.



Orange Peel. This is caused by insufficient thinning of the dope or holding the spray gun too far from the surface. It can also be caused by too high an atomising pressure, use of thinners that is too fast drying or by a cold, damp draught over the surface.



Pinholes. These are smaller versions of a blister. Apart from the causes listed in the ’Bubbles or Blister’ paragraph, they can be caused by water or oil in the spray system. An air temperature that is too high can also be a cause.



Roping. This is a condition in which the surface dries as the dope is being brushed, resulting in an uneven surface. This is common when the dope is cold and has not been brought up to the temperature of the spray shop. When applying dope with a brush, it should not be overbrushed. The brush should be filled with dope then stroked across the surface and lifted off. The pressure applied to the brush should be sufficient to ensure the proper penetration of the dope.



Rough Finish. Dirt and dust on the surface, insufficient sanding and too low a working temperature can all cause a rough finish.



Runs and Sags. This type of defect is caused by too thick a coat, especially on vertical surfaces. This can be the result of incorrectly adjusted spray equipment or incorrect technique.



Wet Areas. This is a larger version of the defect described in the ’Fisheyes’ paragraph.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE AIRCRAFT MATERIALS – COMPOSITE AND NON-METALLIC (DCAM 6.3 L2)

GENERAL CONSIDERATIONS 

The weight of the dope applied to the fabric is an indication that the scheme has been correctly applied. In the BS X26 doping schemes the weight per unit area is given and should be checked by doping a test panel at the same time as the structure. The fabric is weighed before doping and then again after doping, the difference being the weight of the dope film. United States Military Specifications call for a minimum dope weight of 161 g/m2 (4.75 oz/yd2). A tolerance of +/-20% may be applied to the weights given in BS X26.



When an aircraft is re--covered and re--doped it is essential that it is re-- weighed and a new Weight Schedule raised



After the re--covering, repair and doping of control surfaces it is essential that the static balance of each surface is checked against the manufacturer’s requirements. Addition of weight aft of the hinge line without correction of the static balance is likely to cause flutter of the control surface.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM 6.4 L1 & L3)

6.4 TYPES OF CORROSION Corrosion is a very general term and may appear in a variety of forms, depending on the metal involved and the corrosion-producing agents present. Oxidation

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One of the simpler forms of corrosion is “dry” corrosion or, as it is most generally known, oxidation.

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When a metal such as aluminium is exposed to a gas containing oxygen a chemical reaction takes place on the surface between the metal and the gas. Two aluminium atoms join three oxygen atoms to form aluminium oxide (AL2 O3 ). If the metal is iron or steel, two atoms of iron join three atoms of oxygen to form iron oxide, or rust (Fe2 O3 )

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The best way to protect iron from dry corrosion is to keep oxygen from coming into contact with its surface. This is done temporarily by covering the surface with oil or grease or permanently with a coat of paint. Aluminium alloy can be protected from oxidation by the formation of an oxide film on its surface. The protection afforded by an aluminium oxide coating is the principal reason for cladding (Alclad) aluminium alloy used in structural applications.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM 6.4 L1 & L3)

Uniform Surface Corrosion Where an area of unprotected metal is exposed to an atmosphere containing battery fumes, exhaust gases, or industrial contaminants, a uniform over the entire surface occurs. This dulling of the surface is caused by microscopic amounts of the metal being converted into corrosion salts. If these deposits are not removed and the surface protected against further action, the surface become so rough that corrosion pits form.

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Corrosion sometimes spreads under the surface and cannot be recognised by either roughening of the surface or by a powdery deposit. A common type of uniform surface corrosion is caused by the reaction of metallic surface with atmospheric contaminants. Reactive compounds from exhaust gases, as well as fumes from storage batteries, frequently cause uniform surface corrosion.

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The amount of damage caused by uniform surface corrosion is ordinarily determined by comparing the thickness of the corroded metal with that of an undamaged specimen.

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Corrosion

Corrosion and its control are of primary importance to all aircraft operators. Corrosion weakens primary structural members, which must then be replaced or reinforced in order to sustain flight loads. Corrosion is a natural phenomenon which attacks metal by chemical or electrochemical action and converts it into a metallic compound, such as an oxide, hydroxide or sulphate. Substances that cause corrosion are called corrosive agents. Water or water vapour containing salt combine with oxygen in the atmosphere to produce the most prominent corrosive agents. Additional corrosive agents include acids, alkalis and salt. The appearance of corrosion varies with various metals. Examples:  

On aluminium alloys and magnesium it appears as surface pitting and etching, often combined with a grey or white powdery deposits. On copper and copper alloys corrosion forms a greenish film and on steel a reddish rush.

There are two general classifications of corrosion, chemical and electro chemical, both types involve two simultaneous changes. The metal that is attacked or oxidized suffers an anodic change, and the corrosive agent is reduced and suffers a cathodic change.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM 6.4 L1 & L3)

Chemical Corrosion Pure chemical corrosion results from direct exposure of bare surface to caustic resulting liquid or gaseous agents. The most common agents causing direct chemical corrosion include:

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1. Spilled battery acid or fumes from batteries 2. Residual flux deposits resulting from inadequate cleaned, welded, brazed or soldered joints. 3. Entrapped caustic cleaning solutions.

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Electrochemical Corrosion

Electrochemical corrosion is similar to the electrolytic reaction that takes place in dry cell battery. When the number of electrons matches the number of protons in an atom, the atom is said to be electrically balanced. If there are more or fewer electrons than protons, the atom is said to be charged and is called ion, but if there are more protons than electrons it is a positive ion.

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Metals are arranged to show the relative ease with which they ionize in what is called the electrochemical series.

The earlier a metal appears in the series, the more easily it gives up electrons. In other words, a metal that gives up electrons is known as an anodic metal and corrodes easily. On the other hand, metals that appear later in the series do not give up electrons easily and are called cathodic metals. Pitting Corrosion

The most common type of corrosion on aluminium and magnesium is pitting. Pitting first appears as a white powdery deposit. It starts on the surface of a material and then extends vertically into the material. This type of corrosion is dangerous because of the vertical extension, which decrease the material strength. They penetrate deeply into the metal and cause damage completely out of proportion to the amount of metal consumed.

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Galvanic Corrosion

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This common type of corrosion occurs any time two dissimilar metals make electrical contact in the presence of an electrolyte. Example:

Galvanic corrosion can take place where dissimilar metal skins are riveted together, or where aluminium inspection plates are attached to the structure with steel screws. When metals of the same galvanic grouping are joined together, they show little tendency for galvanic corrosion. But metals of one group corrode when they are held in contact with those in another group.

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Filiform Corrosion

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Filiform corrosion is a special form of oxygen concentration cell corrosion or crevice corrosion which occurs on metal surface having an organic coating system. It is recognized by its fine threadlike lines under a polyurethane enamel finish. Filiform corrosion often results when the wash primer used on a metal has not been properly cured. A wash primer is a two-part metal preparation material in which phosphoric acid converts the surface of the metal into a phosphate film that protects the metal from corrosion, and provides an excellent bend for paint. Filiform corrosion shows itself as a puffiness under the paint film and is first noticed around rivet heads and along the lap joint of skins. When the paint film is broken, you will notice that the puffiness was caused by the growth of the powdery salts of corrosion. There is no cure for filiform corrosion short of stripping all the paint, removing the corrosion, treating the metal’s surface and refinishing the aircraft. Filiform can be prevented by storing aircraft in an environment with relative humidity below 70 percent, using coating systems having a low rate of diffusion for oxygen and water vapours, and by washing aircraft to remove acidic contaminants from the surface, such as those created by pollutant in the air.

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Intergranular Corrosion This type of corrosion is an attack along the grain boundaries of an alloy and commonly results from a lack of uniformity in the alloy structure. Aluminium alloys and some stainless steels are particularly susceptible to this form of electro-chemical attack. The lack of uniformity is caused by a change that occurs in the alloy during heating and cooling.

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Intergranular corrosion may exist without visible surface evidence. Very severe intergranular corrosion may sometimes cause the surface of a metal to “exfoliate”. This is a lifting or flaking of the metal at the surface due to delamination of the grain boundaries caused by the pressure of corrosion residual product build up.

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Intergranular corrosion occurs within the metal itself, rather than on the surface; therefore it is quite difficult to detect without ultrasonic or eddy-current equipment.

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Stress Corrosion

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Stress corrosion occurs when mental is subjected to a tensile stress in the presence of a corrosive environment. The stresses in the metal can come from improper quenching after heat treatment, or from an interference fit of a fastener. Stress corrosion cracking is found in most metal systems; however it is particularly characteristic of aluminium, copper, certain stainless steel and high-strength alloy steels. Since stress corrosion can occur only in the presence of tensile stresses, one method for preventing this type of corrosion in some heat-treated aluminium alloy parts is to shot-peen the surface to provide a uniform compressive stress on the surface. Common locations for stress corrosion to form are between rivets in a stressed skin, around pressed-in bushings and tapered pipe fittings. If stress corrosion is severe enough; it may be visible through careful visual inspection. However, dye penetrate inspection is required to find the actual extent of the crack.

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Fretting Corrosion

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When two surfaces fit tightly together but can move relative to one another, corrosion occurs. This type of corrosion is the result of the abrasive wear cause by the two surfaces rubbing against each other. This rubbing, known as fretting, prevents the formation of protective oxide film, exposing active metal to the atmosphere. When this type of corrosion makes its appearance on the surface, the damage is usually done and the parts must be replaced. Fretting corrosion occurs around rivets in a skin and is indicated by dark deposits around around the rivet heads streaming out behind, giving the appearance of rivet smoking. Rivets showing this sign of fretting must be drilled out and replaced.

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Organic Growths

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Water which condenses in fuel tanks produces relatively minor corrosion problems. Microbial growth occurs at the interface of water and fuel, where the fungus feeds on the fuel. The fungus typically attaches itself to the bottom of the tank and looks like a brown deposit the tank coating when the tank is dry. The fungus growth may start again when water and fuel are present. This water contains microscopic animal and plant life called microbes. These organic bodies live in the water and feed on the hydrocarbon fuel. The dark insides of the fuel tank promote their growth, and in very short periods of time these tiny creatures multiply and form a scum inside the tank. This scum can grow to cover the entire bottom of a tank and hold water in contact with the tank structure. If the scum forms along the edge of the sealant in an integral fuel tank, the sealant can pull away from the structure, causing a leak and an expensive resealing operation. It is virtually impossible to prevent the formation of this scum as long as microbes are allowed to live in fuel. The most successful solution to the problem has been to use an additive in fuel which kills these organic growths and prevents the formation of the corrosion-forming scum.

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MICROBIAL CORROSION Microbial attack includes actions of bacteria, fungi or moulds. Micro-organisms occur nearly everywhere. Those organisms causing the greatest corrosion problems are bacteria and fungi.

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Bacteria may be either aerobic or anaerobic. Aerobic bacteria require oxygen to live. They accelerate corrosion by oxidizing sulphur to produce sulphuric acid. Bacteria living adjacent to metals may promote corrosion by depleting the oxygen supply or by releasing metabolic products. Anaerobic bacteria, on the other hand, can survive only when free oxygen is not present. The metabolism of these bacteria requires them to obtain part of their sustenance by oxidizing inorganic compounds, such as iron, sulphur, hydrogen and carbon monoxide. The resultant chemical reactions cause corrosion.

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Fungi are micro-organisms that feed on organic materials. While low humidity does not kill microbes, it slows their growth to prevent corrosion damage. Ideal growth conditions for most micro-organisms are temperatures 68-104°F (20--40°C) and relative humidity85--100%.

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It was once thought that fungal attack could be prevented by applying moisture-proofing coatings to nutrient material or by drying the interiors of compartments with desiccants. However, some moisture-proofing coatings are attacked by mould, bacteria or other microbes, especially if the surfaces on which they are used are contaminated. Microbial growth occurs at the interface of water and fuel, where the fungus feeds on the fuel. Organic acids, alcohols and esters are produced by growth of the fungus. These by-products provide even better growing conditions for the fungus. The fungus typically attaches itself to the bottom of the tank and looks like a brown deposit on the tank coating when the tank is dry. The fungus growth may start again when water and fuel are present. The spore form of some micro-organisms can exist for long periods while dry, and become active when moisture is present. When desiccants become saturated and unable to absorb moisture passing into the affected area, micro-organisms can begin to grow. Dirt, dust and other airborne contaminants are the least-recognized contributors to microbial attack. Unnoticed, small amounts of airborne debris may be sufficient to promote fungal growth. Microbial corrosion can be minimized with a maintenance programme which includes programmed draining of water from fuel tank traps, followed by inspection for milky white products that indicate microbial growth is present, tank inspections, total removal of microbial growth and application of biocide with effected soak periods.

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Factors Affecting Corrosion Many factors will affect the cause, type, speed of attack, and seriousness of metal corrosion. Some are beyond the control of the aircraft designer or maintenance engineer while some of them can be controlled.

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Climatic

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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.

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 Moisture laden atmosphere as against a dry atmosphere. The USA store hundreds of aircraft in a desert (dry) atmosphere for emergency war use.

 Temperature considerations i.e. Hot climate against cold climate. High temperatures will increase the rate of corrosion (all chemical reactions occur faster at higher temperatures). The worst conditions would exist in a hot, wet, maritime environment. Size and Type of Metal

Some metals corrode more easily than others. Magnesium corrodes readily, whilst Titanium is extremely corrosion-resistant because it oxidises readily. Thick structural sections are also more susceptible than thin sections, because variations in physical characteristics are greater. Such sections are also likely to have been cold worked and are, therefore, more susceptible to stress corrosion. Corrosive Agents Foreign materials, that may adhere to metal surfaces, and, consequently result in corrosion, can include:  Soil and atmospheric dust

 Oil, grease and engine exhaust residues

 Salt water and salt moisture condensation

 Spilled battery acids and caustic cleaning solutions  Welding and brazing flux residues

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Common Metals and Corrosion Products One of the problems involved in corrosion control, is the recognition of corrosion products whenever they occur. The following brief descriptions are of typical corrosion products, common to materials used in aircraft construction.

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Iron and Steel

The most common, and easily-recognisable, form of corrosion is red rust. The initial oxide film, formed on freshly exposed steel, is very thin and invisible. In the presence of water, or in a damp atmosphere, especially if sulphur dioxide (industrial atmosphere) or salt (marine environment) is present, thick layers of hydrated oxide develop. These layers vary in colour from brown to black. Rust promotes further corrosion by retaining salts and water. Mill scale (a type of oxide formed at high temperatures), also promotes rusting, by forming an electrolytic cell with the underlying steel. Heavy deposits of rust can be removed only by abrasive blasting or by immersion in rust-removing solutions.

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Surface rust can develop on steel nuts, bolts and other fasteners and may not adversely affect the operational integrity of the equipment. Its appearance is an indication that adequate maintenance procedures have not been followed. Aluminium Alloys

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

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

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

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Cadmium can penetrate the surface of titanium alloys and embrittle them at all temperatures above ambient (as can Lead, Tin and Zinc at temperatures higher than approximately 120°C)). Embrittlement can occur if the cadmium is plated onto the titanium or if cadmium-plated steel parts (and cadmiumcontaminated spanners) are used with titanium. Great care must be taken to ensure that these conditions never occur if at all possible. Copper Alloys

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

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

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Nickel and Chromium Electroplated nickel is used as a heat-resistant coating, while chromium is used for its wear-resistance. Both metals protect steel only by excluding the corrosive atmosphere. The degree of protection is proportional to the thickness of the coating. Once the underlying steel is exposed (through loss of the coating, due to abrasion or other damage), then the coatings actually accelerate the rusting, due to the fact that the steel is more anodic than the protective coating.

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

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Corrosion Removal

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General treatments for corrosion removal include:

 Cleaning and stripping of the protective coating in the corroded area.  Removal of as much of the corrosion products as possible.  Neutralisation of the remaining residue.  Checking if damage is within limits

 Restoration of protective surface films

 Application of temporary or permanent coatings or paint finishes. Cleaning and Paint Removal.

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

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DCAM PART 66 CAT B2 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM REF 6.4 L1 & L3)  Brush the area with stripper, to a depth of approximately 0.8 mm – 1.6 mm (0.03 in – 0.06 in). Ensure that the brush is only used for paint stripping.  Allow stripper to remain on the surface long enough for the paint to wrinkle. This may take from 10 minutes to several hours.  Re-apply the stripper to those areas which have not stripped. Non-metallic scrapers may be used.  Remove the loosened paint and residual stripper by washing and scrubbing the surface with water and a broom or brush. Water spray may assist, or the use of steam cleaning equipment may be necessary. Note. Strippers can damage composite resins and plastics, so every effort should be made to 'mask' these vulnerable areas.

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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 will suggest possible further corrosive attack on more critical areas. The most practical means of controlling the corrosion of steel is the complete removal of corrosion products by mechanical means. Abrasive papers, power buffers, wire brushes and steel wool are all acceptable methods of removing rust on lightly stressed areas. Residual rust usually remains in pits and crevices. Some (dilute) phosphoric acid solutions may be used to neutralise oxidation and to convert active rust to phosphates, but they are not particularly effective on installed components. High-Stressed Steel Components

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

Corrosion attack, on aluminium surfaces, gives obvious indications, since the products are white and voluminous. Even in its early stages, aluminium corrosion is evident as general etching, pitting or roughness. Aluminium alloys form a smooth surface oxidation, which provides a hard shell, that, in turn, may form a barrier to corrosive elements. This must not be confused with the more serious forms of corrosion.

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DCAM PART 66 CAT B2 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM REF 6.4 L1 & L3) General surface attack penetrates slowly, but is speeded up in the presence of dissolved salts. Considerable attack can take place before serious loss of strength occurs. Three forms of attack, which are particularly serious, are:  Penetrating pit-type corrosion through the walls of tubing.  Stress corrosion cracking under sustained stress.  Intergranular attack characteristic of certain improperly heat treated alloys.

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Treatment involves mechanical or chemical removal of as much of the corrosion products as possible and the inhibition of residual materials by chemical means. This, again, should be followed by restoration of permanent surface coatings.

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Alclad

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WARNING: USE ONLY APPROVED PAINT STRIPPERS IN THE VICINITY OF REDUX BONDED JOINTS. CERTAIN PAINT STRIPPERS WILL ATTACK AND DEGRADE RESINS. USE ADEQUATE PERSONAL PROTECTIVE EQUIPMENT WHEN WORKING WITH CHEMICALS. USE ONLY THE APPROVED FLUIDS FOR REMOVING CORROSION PRODUCTS. INCORRECT COMPOUNDS WILL CAUSE SERIOUS DAMAGE TO METALS.

Obviously great care must be taken, not to remove too much of the protective aluminium layer by mechanical methods, as the core alloy metal may be exposed, therefore, where heavy corrosion is found, on clad aluminium alloys, it must be removed by chemical methods wherever possible. Corrosion-free areas must be masked off and the appropriate remover (usually a phosphoric-acid based fluid) applied, normally with the use of a stiff bristled brush, to the corroded surface, until all corrosion products have been removed. Copious amounts of clean water should, next, be used to flood the area and remove all traces of the acid, then the surface should be dried thoroughly. Note: A method of checking that the protective aluminium coating remains intact is by the application of one drop of diluted caustic soda to the cleaned area. If the alclad has been removed, the alumium alloy core will show as a black stain, whereas, if the cladding is intact, the caustic soda will cause a white stain. The acid must be neutralised and the area thoroughly washed and dried before a protective coating (usually Alocrom 1200 or similar) is applied to the surface. Further surface protection may be given by a coat of suitable primer, followed by the approved top coat of paint.

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DCAM PART 66 CAT B2 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM REF 6.4 L1 & L3) Magnesium Alloys The corrosion products are removed from magnesium alloys by the use of chromic/sulphuric acid solutions (not the phosphoric acid types), brushed well into the affected areas. Clean, cold water is employed to flush the solution away and the dried area can, again, be protected, by the use of Alocrom 1200 or a similar, approved, compound.

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Acid Spillage

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

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Aircraft batteries, of the lead/acid type, give off acidic fumes and battery bays should be well ventilated, while surfaces in the area should be treated with antiacid paint. Vigilance is required of everyone working in the vicinity of batteries, to detect (as early as possible) the signs of acid spillage. The correct procedure to be taken, in the event of an acid spillage, is as follows:

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 Mop up as much of the spilled acid using wet rags or paper wipes. Try not to spread the acid.

 If possible, flood the area with large quantities of clean water, taking care that electrical equipment is suitably protected from the water.  If flooding is not practical, neutralise the area with a 10% (by weight) solution of bicarbonate of soda (sodium bicarbonate) with water.  Wash the area using this mixture and rinse with cold water.

 Test the area, using universal indicating paper (or litmus paper),to check if acid has been cleaned up.

 Dry the area completely and examine the area for signs of damaged paint or plated finish and signs of corrosion, especially where the paint may have been damaged.  Remove corrosion, repair damage and restore surface protection as appropriate. Alkali Spillage

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

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Mercury Spillage WARNING: MERCURY (AND ITS VAPOUR) IS EXTREMELY TOXIC. INSTANCES OF MERCURY POISONING MUST, BY LAW, BE REPORTED TO THE HEALTH AND SAFETY EXECUTIVE. ALL SAFETY PRECAUTIONS RELATING TO THE SAFE HANDLING OF MERCURY MUST BE STRICTLY FOLLOWED.

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Mercury contamination is far more serious than any of the battery spillages and prompt action is required to ensure the integrity of the aircraft structure.

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While contamination from mercury is extremely rare on passenger aircraft, sources of mercury spillage result from the breakage of (or leakage from) containers, instruments, switches and certain test equipment. The spilled mercury can, quickly, separate into small globules, which have the capability of flowing (hence its name ‘Quick Silver’) into the tiniest of crevices, to create damage.

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Mercury can rapidly attack bare light alloys (it forms an amalgam with metals), causing intergranular penetration and embrittlement which can start cracks and accelerate powder propagation, resulting in a potentially catastrophic weakening of the aircraft structure. Signs of mercury attack on aluminium alloys are greyish powder, whiskery growths, or fuzzy deposits. If mercury corrosion is found, or suspected, then it must be assumed that intergranular penetration has occurred and the structural strength is impaired. The metal in that area should be removed and the area repaired in accordance with manufacturer’s instructions. Ensure that toxic vapour precautions are observed at all times during the following operation:  Do not move aircraft after finding spillage. This may prevent spreading.

 Remove spillage carefully by one of the following mechanical methods:

 Capillary brush method (using nickel-plated carbon fibre brushes).  Heavy-duty vacuum with collector trap.

 Adhesive tape, pressed (carefully) onto globules may pick them up  Foam collector pads (also pressed, carefully, onto globules).

 Alternative, chemical methods, of mercury recovery entail the use of:  Calcium polysulphide paste.

 Brushes, made from bare strands of fine copper wire

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DCAM PART 66 CAT B2 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM REF 6.4 L1 & L3)  Neutralise the spillage area, using ‘Flowers of Sulphur’.  Try to remove evidence of corrosion.  The area should be further checked, using radiography, to establish that all globules have been removed and to check extent of corrosion damage.  Examine area for corrosion using a magnifier. Any parts found contaminated should be removed and replaced. Note: Twist drills (which may be used to separate riveted panels, in an attempt to clean contaminated surfaces) must be discarded after use.

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

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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.

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Electro-Plating

There are two categories of electro-plating, which consist of:

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

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

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

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Surface Conversion Coatings These are produced by chemical action. The treatment changes the immediate surface layer into a film of metal oxide, which has better corrosion resistance than the metal. Among those widely used on aircraft are:  Anodising of aluminium alloys, by an electrolytic process, which thickens the natural, oxide film on the aluminium. The film is hard and inert.

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 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.

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Other surface conversion coatings are produced for special purposes, notably the phosphating of steel. There are numerous proprietary processes, each known by its trade name (e.g. Bonderising, Parkerising, or Walterising).

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Locations of Corrosion in Aircraft

Certain locations in aircraft are more prone to corrosion than others. The rate of deterioration varies widely with aircraft design, build, operational use and environment. External surfaces are open to inspection and are usually protected by paint. Magnesium and aluminium alloy surfaces are particularly susceptible to corrosion along rivet lines, lap joints, fasteners, faying surfaces and where protective coatings have been damaged or neglected. Exhaust Areas

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

The protective finish, on engine frontal areas, is abraded by dust and eroded by rain. Heat-exchanger cores and cooling fins may also be vulnerable to corrosion.

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DCAM PART 66 CAT B2 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM REF 6.4 L1 & L3) Special attention should be given, particularly in a corrosive environment, to obstructions and crevices in the path of cooling air. These must be treated as soon as is practical. Landing Gear Landing gear bays are exposed to flying debris, such as water and gravel, and require frequent cleaning and touching-up. Careful inspection should be given to crevices, ribs and lower-skin surfaces, where debris can lodge. Landing gear assemblies should be examined, paying particular attention to magnesium alloy wheels, paintwork, bearings, exposed switches and electrical equipment.

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Frequent cleaning, water-dispersing treatment and re-lubrication will be required, whilst ensuring that bearings are not contaminated, either with the cleaning water or with the water-dispersing fluids, used when re-lubricating. Bilge and Water Entrapment Areas

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

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

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

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DCAM PART 66 CAT B2 MODULE 6 MATERIALS AND HARDWARE CORROSION (DCAM REF 6.4 L1 & L3) Aluminium Alloy Skins The most vulnerable skins are those which have been integrally machined, usually in main-plane structures. Due to the alloys and to the manufacturing processes used, they can be susceptible to intergranular and exfoliation corrosion. Small bumps or raised areas under the paint sometimes indicate exfoliation of the actual metal. Treatment requires removal of all exfoliated metal followed by blending and restoration of the finish.

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Spot-Welded Skins and Sandwich Constructions

Corrosive agents may become trapped between the metal layers of spot-welded skins and moisture, entering the seams, may set up electrolytic corrosion that eventually corrodes the spot-welds, or causes the skin to bulge. Generally, spot-welding is not considered good practice on aircraft structures.

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

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Electrical Equipment

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

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

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

6.5 FASTENERS The installation of fasteners is one of the usual procedures used to attached components or assemblies to aircraft structures. Screw Threads

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A screw thread is the ridge produced by forming on a cylindrical or conical surface, a continuous helical groove of uniform section. Refer to Figure 43

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 43: Screw Threads

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Screw Nomenclature Objective: At the end of this lesson the student will be able to explain screw threads. Screw threads are identified by their profile. Some examples are: 1. Metric Thread

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2. Metric Fine Thread

3. Whitworth Thread

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4. Trapezoidal Thread

5. Buttress Thread

6.

Round Thread

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Thread Forms, Dimensions and Tolerances for Standard Threads used in Aircraft Objective: At the end of this lesson the student will be able to identify the thread form.

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Screw Thread Terminology

1. Major Diameter  The diameter of in imaginary cylinder, coaxial with the screw line and touching the CRESTS of an external or ROOTS of an internal thread.

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2. Minor Diameter  The diameter of an imaginary, coaxial with the screw centre line and touching the ROOTS of an external or the CRESTS of an internal thread.

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3. Effective Diameter The diameter of an imaginary cylinder, coaxial with screw centre line, the diameter being such that where it intersects the thread form, the distance between the intersections on the adjacent flanks equals 1/2 pitch. 4. Pitch The distance measured parallel to the screw axis between corresponding points on consecutive thread forms.

5. Lead The axial distance advanced by the screw in one revolution, the lead equals the pitch for single start threads. On multiple start threads the lead is the same multiplied by the pitch as the number of starts. 6. Flank The surface of the thread which connects the root and the crest.

7. Flank Angle The angle between the flank of the thread and a line drawn perpendicular to the thread centre line (axis). 8. Thread Angle This is the included angle between the flanks of the thread form twice the flank angle.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 9. Length of Engagement The axial distance over which two mating threads are designed to make contact. 10. Root That part of the surface of the thread which connects adjacent flanks at the bottom of the thread.

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11. Crest That part of the surface of the thread which connects adjacent flanks at the top of the thread.

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12. Truncation A truncation thread is one having flat crests. 13. Depth of Thread The distance between the root and the crest, measured at right angles to the centre line (axis).

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Refer to Figures 44 and 45

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 44: Screw Threads

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 45: Screw Threads

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Coarse Pitch • • •

Few threads per inch. Has stronger threads, but a slightly weaker core diameter than a fine pitch thread of the same crest diameter. Owing to its greater lead, a coarse pitch thread gives a more rapid action.

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Fine Pitch • • •

Many threads per inch. This type of thread normally used on fastening devices of aircraft construction. Advantages are; the fine pitch thread gives a stronger core diameter, tighter grip, finer adjustment, and is more resistant slackening under vibration.

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Refer to Figure 46

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 46: Coarse Pitch and Fine Pitch

Effective Diameter Tolerance • This is derived from a three-part formula, which takes account of diameter, pitch and length of engagement Major Diameter Tolerance • With external threads the tolerance on major diameter is derived solely from a formula based on pitch.

Minor Diameter Tolerance • The minor diameter tolerance on external threads is related directly to the effective diameter tolerance.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 1. Unified Thread  The unified system or screw thread introduced by Canada, USA and the UK was provided to give a common standard between the three countries. The range of threads include:a) U.N.C. Unified coarse pitch thread with progressive pitch sizes (i.e. pitch varies with the diameter). b) U.N.F. Unified fine pitch thread with progressive pitch sizes.   c) U.N.E.F. Unified extra fine pitch with progressive pitch sizes.   d) U.N. Unified thread with constant pitch (regardless of diameter).

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e) U.N.S. Unified thread or special, pitch/diameter not included above

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) f) U.N.J. A recent addition to the unified series and is designed for increases fatigue strength where stress levels are high. It features an enlarged root radius on the external thread and is particularly useful for aircraft applications. Sizes are quoted in fractions of an inch above 1/4". Size below 1/4" are designated by a number related to its size, followed by a number indicating the threads per inch (T.P.I) E.g. 4-40 UNC Refer to Figure 50 

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 50: 4-40 UNC

Notes:

The symbol used for visual identification of Unified thread is the circle; usually three circles with their borders touching. This symbol is marked on the part or adjacent to the component, so that it will be visible after assembly

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Types of Threads 2. Whitworth (BS 84.1956). This is a form of thread designed by Sir Joseph Whitworth in the middle of the 19th century and represents one of the first attempts to standardise threads and make them inter-changeable.

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It is a symmetrical ‘V’ shaped thread form with threads of 55°, rounded equally at crests and roots. They have been standardizes by 3 different series of threads, all of Whitworth form. These are: • British Standard Whitworth (B.S.W.) • British Standard Fine (B.S.F.) • British Standard Pipe (B.S.P.)

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In the B.S.W. and B.S.F. series, the range of diameters are much the same, but as its name implies, the pitches in the B.S.F. range are finer for the same diameters in the B.S.W. range. The diameters referred to are the major diameters, measured over the crests of the thread. In the case of the B.S.P. threads, the nominal diameter is the internal diameter of the pipe with which it is associated, so that the major diameter is always larger than the nominal diameter. The general relationship between these threads is illustrated in Figure 47, where the relative pitches, together with the major diameters in the B.S.P. range are given for a representative selection of nominal sizes Refer to Figure 47

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 47: Whitworth Tables and Specifications

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 3. British Association (B.A.) (BS93.1951)  The B.A. was introduced by the British Association when the need was felt for a series of small threads for use in the scientific instrument industry. It has since been widely applied in British Engineering products of all classes. Today, its greatest uses are in the field of radio and other electrical equipment. The threads are dimensioned in metric units • Each thread is designated with a number between 0 and 25. • The largest is 0 BA (0 BA = 6mm). • The major diameter of the 25 BA is 0.25 mm (0.010"). The form of the BA thread is illustrated below and differs from the Whitworth in thread angle of 47 ½ °

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Refer to Figure 48

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 48: British Association

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 4. Metric Thread (BS 1095)  The metric thread is very widely used on the European continent. The thread angle is 60° and is truncated. Any metric pitch can be associated with any diameter as desired, but there are various standard series in force in different countries. The series in general use in Britain is tabulated in Specification BS 1095. Generally speaking, variations in standard between countries only effect threads at the extremes of the range, below 6 mm and above 80 mm diameter.

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Refer to Figure 49

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 49: Metric Thread

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 5. American National Thread  The American Thread is another 60° angle thread which is truncated, (same as Metric Thread) There are 2 standard series; the Coarse (A.N.C.) and Fine (A.N.F.). These correspond to the B.S.W. and B.S.F. series, most sizes of the A.N.C. thread having the same pitches as the B.S.W. range, although the A.N.F. series is size for size, finer than the B.S.F. range.

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These American thread sizes are expressed in inches and fractions from 1/4" upwards. Sizes below 1/4" are designated by numbers from 12 and below, the form of the thread remaining the same. No.1 is the finest thread in the coarse series (0-073" diameter. 64 T.P.I.) No. 0 is the finest in the fine series (0.060", 80 T.P.I.) The designating number corresponds to the same basic major diameter in both series.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Aircraft Fasteners Identification Most items of aircraft hardware are identified by their specification number of trade name. Threaded fasteners are usually identified by: 1. AN (Air Force-Navy) 2. NAS (National Aircraft Standard) 3. MS (Military Standard)

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Quick-release fasteners are usually identified by factory trade names and size designations. Various types of fastening devices allow quick dismantling or replacement of aircraft parts that must be taken apart and put back together at frequent intervals. Bolts and screws are two types of fastening devices which give the required security of attachment and rigidity. Bolts are used where great strength is required and screws are used where strength is not the deciding factor.

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Bolts and screws are similar in many ways. They are both used for fastening or holding and each has a head on one end and screw threads on the other. A bolt has a fairly short threaded section and a comparatively long grip length or unthreaded portion, where as a screw has a longer threaded section and may have no clearly defined grip length. A bolt assembly is generally tightened by turning the nut on the bolt, but the head of the bolt may or may not be designed for turning. A screw is always tightened by turning its head.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Thread Type and Fits Aircraft bolts, screws and nuts are threaded in either the: 1. 2. 3. 4.

American National Coarse (NC) American National Fine (NF) American Standard Unified Coarse (UNC) American Standard Unified Fine (UNF)

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The difference between the American National series and the American Standard Unified series is the American National series has more threads per inch than the American Standard Unified series.

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Example:

On a one inch diameter bolt, the NF thread specifies 14 threads per inch ( 1-14NF), while the UNF thread specifies 12 threads per inch (1-12 UNF).Both thread types are designated by the number of times the threads rotate (number of turns) around a 1-inch length of given diameter bolt or screw. Threads are also designated by class of fit from one to five. The class of thread indicates the tolerance allowed in manufacturing. 1. 2. 3. 4. 5.

Class 1 thread is a loose fit Class 2 thread is a free fit Class 3 thread is a medium fit Class 4 thread is a close fit Class 5 thread is a tight fit

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) A Class 1 fit allows you to turn the nut all the way down using only fingers. Wing nuts are a good example of Class 1 fit. A Class 4 and 5 fit requires a wrench to turn a nut down from start to finish. Aircraft bolts are usually fine threaded with a Class 3 fit. Whereas screws are typically a Class 2 or 3 fit. Bolts and nuts are also produced with right-hand and left-hand threads. A right-hand thread tightens when turn clockwise, a left-hand thread tightens when turned counter clockwise.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Bolts The bolts, used in the construction of aerospace components and structures, have evolved into a bewildering range of materials, shapes and sizes, all of which are dictated by the applications for which the items have been designed Standards and systems have been established, to provide identification of the many different forms of threaded devices, in order to ensure that only the correct items are installed in the relevant locations.

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

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British Bolts

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

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

UNF.

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

Unified National Fine.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Identification of BS Unified Bolts British Standard Unified (BS Unified) bolts are identified by the use of an alpha-numeric code, which provides information relating to the type, material, surface finish, length, diameter and any other important characteristics of the threaded device Table 9 shows a (very small) selection of aircraft standard bolts and screws with a (shortened) description of the type of device and the materials from which it is made.

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Reference to the table shows that the code A102 signifies a hexagonal-headed bolt which is made of high-tensile steel, while the code A175 represents a 100° countersunk-headed bolt, made from an aluminium alloy. Table 9 Examples of Code Numbers for Unified Threads Description Material Hex. Headed Bolt HTS. Hex. Headed Bolt SS Hex. Close Tolerance. Bolt HTS Shear Bolt HTS 100º Countersunk. Head. Bolt SS 100º Countersunk. Head. Bolt Al Al 100º Countersunk. Head. Screw HTS Pan Head. Screw HTS

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Standard No. A102 A104 A111 A112 A174 A175 A204 A205

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

Three contiguous (touching) circles marked in a convenient position (machine items). Note: Due to the difficulty in applying the identifying marks to individual items, it is planned to merely mark the packets in which the threaded devices are marketed, so that some, or all, of the identification marks will not be seen on the items (particularly screws). Great care must, therefore, be taken to ensure that the items being used are correctly identified and to the approved standard.



A shallow recess in the head of a bolt, equal to the nominal diameter of the thread (cold forged items).



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

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Further numbers and letters are added to the identifying code, to provide information relating to the length (usually of the plain shank or gripping portion) and to the diameter of the items. The length is given by a number, which signifies increments of tenths of an inch, so that a 5 would represent a bolt with a plain shank of 0.5 in, while the number 12 would signify the plain shank as being 1.2 in long Reference to Table 10, will show how the diameter of an item is designated by the addition of another letter to the system, so that a bolt, with the code marking of A102 9 E, would signify a Unified-threaded, hexagon-headed bolt, made from high-tensile steel, with a plain shank length of 0.9 in, and a diameter of ¼ in.

Code Y Z A B C D E G

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

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Note: In the earlier UK system (which may be encountered on older, or home-constructed, light aircraft), bolts more than ¼ inch diameter are normally BSF, whilst bolts less than ¼ inch diameter (and most screws) are BA. Both of these items also use a number to represent their nominal length and a letter code (as can be seen in Table 11) to identify their diameter.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Other bolts of this era may have nicks at the corners of the head (High Tensile Steel) or a raised ring on the bolt head (Cold Rolled) to assist differentiation of their particular designations. Table 11 EXAMPLES OF BA AND BSF BOLT AND SCREW CODES Code Size Code Size A 6 BA P 9/16" BSF B 4 BA Q 5/8” BSF C 2 BA S 3/4" BSF E 1/4” BSF U 7/8" BSF G 5/16" BSF W 1" BSF J 3/ 8" BSF X 12 BA L 7/16" BSF Y 10 BA N 1/2" BSF Z 8 BA

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) American Bolts American aircraft bolts and nuts are threaded in the NC (American National Coarse), the NF (American National Fine), the UNC (Unified National Coarse), and the UNF (Unified National Fine) thread series. The item is often coded to give the diameter of the threaded portion and the number of threads per inch (tpi).

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Aircraft bolts may be made from HTS, Corrosion-Resistant Steel or Aluminium Alloy. Head types may be hexagonal, clevis, eyebolt, internal wrenching and countersunk (refer to Fig. 43) and head markings may be used to indicate other features such as close tolerance, aluminium alloy, CRS or other types of steel.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Examples of Aircraft Bolts Fig 43

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Identification of AN Standard Bolts While there are several different US Standards, there is only need to discuss one type for the purpose of these course notes, as the others are very similar. AN bolts come in three head styles, Hexagon Head, Clevis and Eyebolts and Table 12 provides an indication of the various code numbers in use.

AN No.

3 – 20

Table 12 EXAMPLES OF AN STANDARD BOLTS (EARLY SERIES) Type Material Process Thread Size

Bolt, hex. Head

Steel

Cadmium Plated Nil Anodised

No. 10 to 1¼”

Thread Type

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UNF

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No. 6 to

1”

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21 – 36

Bolt, Clevis

Steel

Cadmium Plated

UNF

42 – 36

Bolt, Eye

Steel

Cadmium Plated

No. 10 to 9/16”

UNF

73 – 81

Bolt, hex. Drilled head

Steel

Cadmium Plated

No. 10 to ¾”

UNF or UNC

173 – 186

Bolt, close tolerance

Steel

Cadmium Plated thread & head

No. 10 to 1”

UNF

Note: The later series uses a different number system

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) For identification purposes the AN number is used to indicate the type of bolt and its diameter. In addition a code is used to indicate the material, length and presence of a split pin or locking wire hole as follows: 

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

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 

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

Drilled Shank Steel

CRS

Steel, Close Tolerance

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Diameter

Aluminium Alloy

Drilled Head, (Except AN 73 –81)

Drilled Head, AN 73 -81

Aluminium Alloy, Close Tolerance

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CRS, Close Tolerance

Grip

Head Markings for AN Bolts Fig. 44

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 

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.

Drilled head only

= H added before dash No. (replacing dash) A added e.g. AN25H15A after dash No.

Drilled head and shank

e.g. AN24 – 15A

= H added before dash No.

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e.g. AN25H15

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

f T o g y n r i r a t e e e i n r i p g o n r E P S A M  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

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Special-to-Type Bolts The hexagon headed aircraft bolt AN3 – AN20 (refer to Fig.45), is an all purpose structural bolt used for applications involving tension or shear loads where a light drive fit is permissible.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Eye Bolt

Clevis Bolt

Special-to-Type Bolts Fig. 45

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Alloy steel bolts, smaller than 3/16” diameter, and aluminium alloy bolts smaller than ¼” are not used on primary structure. Other bolts may be used as follows: 

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



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



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



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

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Metric Bolts The identification of a Metric bolt is by the use of the diameter in millimetres, immediately after the capital letter ‘M’. In this way, M6 represents a 6 mmdiameter bolt. The length is also shown in millimetres, so the bolt M6 -15 will be a 6 mm- diameter bolt, which is 15 mm long. The basic terminology, for identifying bolts of the Metric system, involves the nominal length, the grip length and diameter (refer to Fig. 46).

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Length

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Diameter

Metric Bolt Terminology Fig. 46

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Nuts

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

All nuts used in aircraft construction must have some sort of locking device to prevent them from loosening and falling off. Many nuts are held on a bolt by passing a cotter pin through a hole in the bolt shank and through slots or castellation in the nut. Aircraft nuts are made in a variety of shaped and sizes. They are made of cadmium-plated, carbon steel, stainless steel or anodized 2024 -T aluminium alloys. There are two basic types of nuts, self-locking and non self-locking. A self-locking nut locks onto a bolt on its own while a non self-locking nut relies on either a cotter pin, check nut or lock washer to hold it in place.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) AN310 Castle Nut These fine-thread nuts are designed to fit on a standard airframe bolt with a Class 3 fit, and are used when the bolt is subjected to either shear or tensile loads. The size of a nut is indicated in the part code by a dash number which denotes the size of the bolt it fits. Example: AN310-6 nut fits an AN6 bolt which has a diameter of 3/8 inch.

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Castle nuts are available in cadmium-plated nickel steel, corrosion resistant and 2024 aluminium alloy.

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AN 315 Plain Nut

The AN315 plain nut has no castellation and, therefore, cannot be held in place using a cotter pin. There fine-thread nuts have no locking provisions, a springtype lock washer must be used in combination with the nut. The lock washer applies a spring force to prevent the nut from shaking loose.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M

An315 nuts are used with either tensile or shear loads and made of either nickel steel, corrosion-resistant steel and aluminium alloy. Plain nuts are made with both right and left-hand threads. AN350 Wing Nut

Wing nuts are used when it is necessary to remove a part frequently without the use of tools. Aircraft wing nuts are made of either cadmium-plated steel or brass and are available in sizes. All of these nuts have national fine threads that produce a Class 2 Fit. Nuts for machine screw sizes are designated by the series number. Nut used on bolts have a bolt size given in 1/16 inch increments. Self-Locking Nut

Self-locking nuts, or lock nuts, employ a locking device in their design to keep them from coming loose. There are several different types of lock nuts, you must be certain that the proper locknut is used in a given application. Failure to do so could result in failure of the locking provision. The two general types of self-locking nuts used in aviation are the fiber, or nylon types and the all metal type. Self-locking nuts are used on aircraft to provide tight connections which will not shake loose under severe vibration. Do not use self-locking nuts at joint which subject either the nut or bolt to rotation.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Anchor Nuts

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Anchor nuts are permanently mounted nut plates that enable inspection plates and access doors to be easily removed and installed. To make the installation of an access door easier where there are a great numbers of screws, a floating anchor nut is often used. With a floating anchor nut it fits loosely into a small bracket which is riveted to the skin. Since the nut is free to move within the bracket it aligns itself with a screw. To speed the production of aircraft, ganged anchor nuts are installed around inspection plate opening. These are floating-type anchor nuts that are installed in channel that is riveted to the structure.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Screws

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Screws are probably the most commonly used threaded fastener in aircraft. They differ from bolts in that they are generally made of lower strength materials. Screws are typically installed with a loose-fitting thread, and the head shapes are made to engage a screwdriver or wrench. Some screws have a clearly defined grip length while others are threaded along their length. There are three basic classifications of screw used in aircraft construction:

1. Machine screws – which are the most widely used. 2. Structural screws – which have the same strength as bolts. 3. Self-tapping screws – which are typically used to join light weight materials.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Machine Screws Machine screws (refer to Fig. 49) are used extensively for attaching fairings, inspection plates, fluid line clamps and other light structural parts. The main difference between aircraft bolts and machine screws, is that the threads of a machine screw usually run the length of the shank, whereas bolts usually have an unthreaded grip length.

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The most common machine screw used in aviation is the fillister-head screw, which can be wire-locked using the drilled hole in the head. The flat-head (countersunk-head) screw is available with single or cross-point slotted heads. The round-head screw and the truss-head (mushroom-head) screw, provide good holding properties on thin metal sheets.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Studs Types and Uses, Insertion and Removal Objectives:

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At the end of this lesson the student will be able to define 'stud' types and uses, insertion and removal.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A Figure 56: Types of Studs M For Training Purposes Only

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 1. Standard Stud (Refer to Figure 57) This is a plain or parallel type. The diameter of the unthreaded portion is the same as the major diameter of the screw thread at both ends, with the threads merging smoothly into the plain part.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 57: Standard Stud

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2. Waisted Stud (Refer to Figure 58)

This is a weight-saving stud. The plain shank of the stud is reduced to the minor diameter of the screw thread. This will lighten the stud without losing any of its strength.

Figure 58: Waisted Stud

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 3. Stepped Stud (Refer to Figure 59) The purpose of this stud is two fold. First its larger diameter end provides a stronger anchorage, which is particularly useful when the stud is located in soft metal. Second, it is used as a replacement stud and fitted in a stud housing that, because of damage, has been re-drilled and tapped with a larger diameter thread.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 59: Stepped Stud

4. Shouldered Stud (Refer to Figure 60)

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The projecting shoulder of this stud gives extra support and resistance to any side pressure.

Figure 60: Shouldered Stud

5. Fitting Studs

Studs may be inserted by using a stud box and a spanner, or by fitting locknuts to the stud and using a spanner on the upper nut.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Removing Studs Studs that have broken off flush or below the surface may be removed in a variety of ways depending upon the size and the equipment available. Removing Studs

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Ezy-out (Refer to Figure 61)

a) A hole is drilled centrally in the stud. Ezy-out is entered in the hole and turned anti-clockwise with a tap wrench. Ezy-outs are supplied in sets of various diameters.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 61 : Ezy-out

b) Drill a hole of approximately half the stud diameter down the centre of the stud. Drive in a square drift so that square edges cut into the stud and unscrew by using a spanner. c) Drill and tap the stud with a smaller opposite hand thread. Then, unscrew the stud by screwing in an opposite hand bolt. d) Drill the stud out with its normal tapping size drill and re-tap the thread. Then, carefully picking out the old loose threads.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Self Tapping Screws and Dowels Objective: At the end of this lesson the student will be able to define self-tapping screws and dowels.

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Self-tapping screws are used to secure thin gauge sheet metal where nuts and bolts are impracticable. They are screwed into a hole of the correct diameter and form their own thread. The screw head may be slotted or cruciform; • four types are in common used.

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Examples of self-tapping screws are Parker Kalon (P-K), Barber and Calon (B-K). Refer to Figure 62

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 62: Self Tapping Screws, Dowels

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 6.5.3: Locking Devices Locking: Parts and Method Objective:

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At the end of this lesson the student will be able to recognize locking device and wire-locking methods.

Corrosion resisting steel and heat resisting nickel alloy are the materials normally used in the wire recommended for wire locking. The following techniques are equally effective:

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1. Double twisted method. (Refer to Figure 63)

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 63: Single Strand

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 2. Single strand with 2.5 twists at originating end and closing end. (Refer to Figure 64)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 64: Double Twisted

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Washers Washers provide a bearing surface area for nuts, and act as spacers or shims to obtain the proper grip length for a bolt and nut assembly. Washers are also used to adjust the position of castellated nuts with respect to drilled cotter pin holes in bolts as well as apply tension between a nut and material surface to prevent the nut from vibrating loose. The three most common types of washers used in airframe repair are: 1. Plain washers 2. Lock washers 3. Special washers

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Plain Washers

Plain washer provides a smooth surface between a nut and the material being clamped. These washers are made of cadmium-plated steel, commercial brass, corrosion-resistant steel and aluminium alloy 2024. Plain washer should be used under lock washers to prevent damage to the surface material.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Aluminium and aluminium alloy washers may be used under bolt threads or nuts on aluminium alloy or magnesium structures where corrosion caused by dissimilar metals is a factor. It is common practice to use a cadmium-plated steel washer under a nut bearing directly against a structure as this washer will resists the cutting action of a nut better than an aluminium alloy washer. Lock Washers

Lock washers are made of steel and are twisted so that when a nut is tightened against it, the spring action of the washer creates a strong friction force between the bolt threads and those in the nut. Two types of lock washers are used in aircraft construction. The most common is the AN935 split lock washer. The second type of lock washer is the thinner AN936 shake proof lock washer which is available with both internal and external teeth.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 2. Spring Washers Supplied as: • single coils of square section spring with sharp corners or, • double coil of flat spring, which can be re-used if it is still springy and retains it's sharp corners

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Refer to Figure 66 and Figure 67

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 66: Spring Washers

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Figure 67: Spring Washers (Single and Double Coil)

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 3. Shake Proof Washer Spring steel washer, which have slanting serration on their internal or external edge Refer to Figure 68

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 68: Shake proof washer

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 4. Tab Washers Thin metal washers with two tabs and a projection. It is not permissible to straighten the tab of a tab-washer and re-use it. Refer to Figure 69

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 69: Tab Washers

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 5. Locking Plate A thin metal plates, fitted around the nut or bolt after it has been fully tightened. Can be re-use provided they are still a good fit on the nut or bolt. Refer to Figure 70

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 70 : Locking Plate

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 6. Circlips These are either spring plate or spring wire rings that are spring into grooves. They may be fitted internally or externally. Wire type circlips are used once only. Refer to Figure 71

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 71 : Circlips 

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 7. Dzus Fasteners

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Pins The main types of pins used in aircraft structures are the roll pin, clevis pin, cotter pin and taper pin. Pins are used in shear applications and for safe tying.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M ROLL PIN

Roll Pin

CLEVIS PIN

TAPER PIN

Roll pins are often used to provide a pivot for a joint where the pin is not likely to be removed. A roll pin is made of flat spring steel that is rolled into a cylinder but the two ends are not joined. This allows the pin to compress when it is pressed into a hole and create a spring action that holds the pin tight against the edge of the hole. To remove a roll pin, it must be driven from a hole with a proper size pin punch.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Clevis Pin Clevis, or flat-head, pins are used for hinge pins in some aircraft control systems. They are made of cadmium-plated steel and have grip lengths in 1/16 inch increments. When installing a clevis pin place the head in the up position, place a plain washer over the opposite end, and insert a cotter pin through the hole to lock the pin in place.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Cotter Pins

Castellated nuts are locked into drilled bolts by passing a cotter pin through the hole and nut castellation and then spreading the ends of the cotter pin. They are made of either cadmium-plated carbon steel or corrosion-resistant steel. There are two methods of securing cotter pins that are generally acceptable. First method: One leg of the cotter pin is bent up over the end of the bolt, and the other leg is bent down over one of the flats of the nut. Second method: The cotter pin is rotated 90 degrees and the legs wrapped around the castellations.

It is important to note that nuts should never be overtorqued to make the hole in the bolt align with the castellations. If the castellations in the nut fail to align with the drilled bolt hole, add washer under the nut until a cotter pin can be inserted. Taper Pin

Both the plain and threaded taper pin are used in aircraft structures to make a joint that is designed to carry loads. This type of pin does not allow any loose motion or play. The plain taper pin is forced into a hole that has been reamed with standard taper pin reamer and is held in place by traction. Classified by diameter of the small end and length

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Pal-nuts. The capped washer type Palnut shown below has a washer base which can span large holes or odd shaped holes. It can also be provided with a scalloped washer design called a Style DF and can also be provided with in integral sealer. The spring steel provides enough flexibility to provide a self locking feature when properly tightened. The threads, "teeth", of the Palnut provides enough pressure on the threads of the bolt to damage the plating if not properly installed.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Keys Introduction The name key is given to a specially--shaped piece of metal that is used to transmit a drive at considerable mechanical power from a shaft to a hub, or vice versa, when the mating surfaces are otherwise smooth.

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The key is a solid piece of metal that is wedged between the parts, or fitted into matching recesses (or keyways) in the shaft and hub. A key is of rectangular or square section, uniform in width, and of either uniform or tapered thickness. It is produced in many variations, depending upon the situation and the load to be transmitted. In general, keys are used only in circumstances which do not call for frequent separation of the parts.

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Taper Keys

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The agreed engineering standard is for a taper of 1 part in 100 on the thickness, with the tapering surface of the key matching the recess (or keyway) cut into the bore of the hub. Several types of taper key are in common use: 

Hollow Saddle Key. This type of key is hollowed (shaped) to fit the radius of the shaft. When driven into position, its taper provides a friction drive between hub and shaft that is capable of transmitting a moderate load. There is no keyway cut into the shaft and, therefore, hollow keys are not suitable for heavy loads.



Flat Saddle Key. This rectangular or square-section key is driven into a keyway in the hub and bears upon a flat on the shaft. It provides a more positive drive than that achieved by the hollow saddle key.



Gib--Headed Key. This taper key is fitted into keyways which are machined partly in the shaft and partly in the hub. An important feature of fitting keys into these keyways is that the keyways must be perfectly aligned before fitting the key. With this in mind, it may be necessary to use a slave key when assembling the parts together; never rely upon the key to align the keyways as it is driven in. These keys and their keyways are capable of transmitting a much greater driving load than are the saddle-type keys. The head of the Gib--headed key is used as a means of removing the key when it is not possible to drift the key out from the opposite side.



Feather Key. This type of key is used when axial movement is required between the hub and the shaft. An example of the use of a feather key is when a gear or a pulley must slide along a shaft whilst continuing to transmit drive. The keyway in the hub is cut to allow the key the minimum side and top clearance needed to provide a sliding fit.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 

Woodruff Key. This key is made in the shape of a segment of a parallel--sided disc --similar to the capital letter D. It fits into a shaped cavity in the shaft which conforms closely to the profile of the key, and into a uniform keyway in the hub to provide a push fit on the sides with clearance along the top flat face of the key. The advantage of the Woodruff key is that it is suitable for fitting to either parallel or taper shafts.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Keys

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Aircraft Rivets The most common technique of joining sheets of aluminium is riveting. A rivet is a metal pin with a formed head on one end. A rivet is inserted into a drilled hole, and its shank is then deformed by a hand or pneumatic tool. Aluminium solid rivets are mainly used with high-strength, aluminium alloy components and assemblies. Steel, titanium or monel fastener are used for joints in high-temperature areas.

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Before the Federal Aviation Administration issues a Type Certificate for an aircraft the manufacturer must demonstrate that the aircraft conforms to all airworthiness requirements. These requirements pertain not only to performance, but to structural strength and integrity as well. To meet these requirements, each individual aircraft produced from a given must meet the same standards.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

SOLID RIVETS

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BLIND RIVETS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Rivet Codes Rivets are given part codes that indicate their size, head style and alloy material. Two systems are in use today, the Air Force-Navy, or AN system and the Military Standards. The first component of a rivet part number denotes the numbering system used, this can either be AN or MS.

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The second part of the code is a three-digit number that describes the style of rivet head.

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The two most common rivet head styles are the universal head, represented by the code 470, and the countersunk head, which is represented by the code 426. The head designation is a one or two-digit letter code representing the alloy material used in the rivet.

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After the alloy code, the shank diameter is indicated in 1/32 inch increments and the length in increments of 1/16 inch. Example:

The rivet has a diameter of 4/32 inch and is 5/16 of an inch long. The length of universal head (A470) rivet is measured from the bottom of the manufactured head to the end of the shank. The length of a countersunk rivet (AN 426) is measured from the top of manufactured head to the end of the shank.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Rivet Alloys Most aircraft rivets are made of an aluminium alloy. The type of alloy is identified by a letter in the rivet in the rivet code and by a mark on the rivet head itself. 1100 Aluminium

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Rivets made of pure aluminium have no identifying marks on their manufactured head, and are designated by the letter A in the rivet code. As this type of rivet is made out of commercially pure aluminium, the rivet lacks sufficient strength for structural application.

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1100 rivets are restricted to non-structural assemblies such as fairings, engine baffles and furnishings. The 1100 rivet is driven cold, and therefore, its shear strength increases slightly as a result of cold working. 2117 Aluminium Alloy (AD)

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The rivet alloy 2117-T3 is the most widely used for manufacturing and maintenance of modern aircraft. Rivets made of this alloy have a dimple in the centres of the head and are represented by the letters AD in the part codes. As AD rivets are so common and require no head treatment, they are often referred to as “Field Rivets” The main advantage for using 2117-73 for rivets is its high strength and shock resistance characteristics. The alloy 2117-T3 is classified as head-treated aluminium alloy, but does not require re-heat-treatment before driving. 5056 Aluminium Alloy (B)

Some aircraft parts are made of magnesium. If aluminium rivets were used on these parts, dissimilar metal corrosion could result. For this reason, magnesium structures are riveted with 5056 rivets which contain 5 percent magnesium.These rivets are identified by a raised cross on their heads and the letter B in a rivet code.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) 2017 Aluminium Alloy (D) 2017 aluminium alloy is extremely hard. Rivets made of this alloy are often referred to as D rivets and were widely used for aircraft construction. The introduction of jet engines places greater demands for structural strength on aircraft materials and fasteners. The aluminium industry modified 2017 alloy to produce a new version of 2017 aluminium called the crack free rivet alloy.

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D-rivets are identified by raised dot in the centre of their head and the letter D in rivet codes. Because D-rivets are so hard they must be heat treated before they can be used.

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When aluminium alloy is quenched after heat treatment it does not harden immediately. It remains soft for several hours and gradually becomes hard and gain full strength. Rivets made of 2017 can be kept in this annealed condition by removing them from a quench bath and immediately storing them in a freezer.

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D- rivets are often referred to as icebox rivets. These rivets become hard when they warm up to room temperature and may be reheat-treated as many times as necessary without impairing their strength. 2024 Aluminium Alloy (DD)

DD-rivets are identified by two raised dashes on their head. DD-rivets are also called icebox rivets and must be stored at cold temperatures until they are ready to be driven. The length of time the rivets remain soft enough to drive is determined by the storage temperature (below 32˚F); the rivets will remain soft enough to drive for two weeks. Icebox rivets attain one half their maximum strength in approximately 1 hour after driving and full strength in about 4 days. 2024-T rivets exposed to room temperature for a period exceeding 10 minutes. Once an icebox rivet has been taken from the refrigerator, it should not be mixed with the rivets still in cold storage.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Corrosion-Resistant Steel (F) Rivets Stainless steel rivets are used for fastening corrosion-resistant steel sheets in applications such as firewalls and exhaust shrouds. They have no marking on their head. Monel (M) Rivet

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Monel rivets are identified by two recessed dimples in their heads. They are used in place of corrosion-resistant steel rivets when their somewhat lower shear strength is not a detriment.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Special (Blind) Rivets There are many places on the aircraft where access to both sides of riveted structure or structural part is impossible, or where limited space will not permit the use of a bucking bar. In the attachment of many non-structural parts such as aircraft interior, furnishing, flooring and de-icing boots, the full strength of solid shank rivets is not necessary.

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Special (blind) rivets have been designed which can be bucked from the front. They are sometimes lighter than solid-shank rivets, yet amply strong for their intended use. These rivets are produced by several manufacturers and have unique characteristic that require special installation tools, special installation procedures and special removal procedures.

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Two classes of mechanically expanded rivets: 1) Non-structural

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a) Self-plugging (friction lock) rivets b) Pull-thru rivets

2) Mechanical lock, flush fracturing, self-plugging rivets.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M TYPES OF SPECIAL RIVETS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Self-Plugging (Friction Lock) Rivets The self-plugging (friction lock) blind rivets are manufactured by several companies. Self-plugging (friction lock) rivets are fabricated in two parts: 1. A rivet head with a hollow shank or sleeve, and a stem that extends through the hollow shank. 2. A protruding head and a countersunk head self-plugging rivet produced by one company.

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The self-plugging (friction lock) rivets are fabricated from several materials. Rivets are available in the following material combinations: 1. 2. 3.

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Factors to consider in the selection of the correct rivet for installation are: 1. 2. 3. 4.

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Stem 2017 aluminium alloy and sleeve 2117 aluminium alloy Stem 2017 aluminium alloy and sleeve 5056 aluminium alloy Stem steel and sleeve steel.

Installation location Composition of the material being riveted Thickness of the material being riveted Strength desired

If the rivet is to be installed on an aerodynamically smooth surface, or if clearance for an assembly is needed, countersunk head rivets should be selected. In other areas where clearance or smoothness is not a factor, the protruding head type rivet may be utilized.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Pull-Thru Rivets Pull-thru blind rivets are fabricated in two parts: a rivet head with hollow shank or sleeve, and a stem that extends through the hollow shank. Pull-thru rivets are fabricated in two common head styles: 1. 2.

Protruding head or universal head A 100 º countersunk head

Factors to consider in the selection of the correct rivet for installation are: 1. 2. 3. 4.

Installation location Composition of the material being riveted Thickness of the material being riveted Strength desired

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The thickness of the material being riveted determines the overall length of the shank of the rivet. The shank of the rivet should extend beyond the material thickness approximately 3/64 inch to 1/8 inch before the stem is pulled. Self-Plugging (Mechanical Lock) Rivets

Self-plugging (Mechanical lock) rivets are similar to self-plugging (Friction lock) rivets, except for the manner in which the stem is retained in the rivet sleeve. This type of rivet has a positive mechanical locking collar to resist vibration that cause the friction lock rivets to loosen and possibly fall out.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) CherryLocks

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The Cherry mechanical-lock rivet, often called the bulbed CherryLock. The CherryLock rivet is an improvement over the friction-Lock rivet because its center stem is locked into place with a lock ring. This results in shear and bearing strengths that are high enough to allow CherryLocks to be used as replacement for solid shank rivets. CherryLock rivets are available with two head styles, 100 degree countersunk and universal. CherryLock are available with diameters of 1/8, 5/32 and 3/16 inch, with an oversize of 1/64 inch for each standard size. The rivet or shell portion of a CherryLock may be constructed of 2017 aluminium alloy, 5056 aluminium alloy, Monel or stainless steel. CherryLock rivets require a special pulling tool for each different size and head shape. Always use the proper rivet length selection gauge and follow the manufacturer’s installation recommendations.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) HI-Shear Rivets (Special Rivet)

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One of the first special fasteners used by the aerospace industry was the HI-Shear rivet. The HI- Shear rivet has the same strength characteristics as a standard AN bolt. The only difference between the two is that a bolt is secured by a nut and a HI-Shear rivet is secured by a crushed collar. The Hi-Shear rivet is installed with an interference fit, where the side wall clearance is reamed to a tolerance determined by the aircraft manufactures. AHiShear rivet has to be tapped into its hole before the locking collar is swaged on. Hi-Shear rivets are made in two head styles, flat and countersunk. The HI-Shear rivet is designed especially to absorb high shear loads. The HI-Shear rivet is made from steel alloy having the same tensile strength as an equal size AN bolt. The lower portion of its shank has a specially milled groove with a sharp edge that retains and finishes the collar as it is swaged into the locked position.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Hi- Loks

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HI-Lok bolts are manufactured in several different alloys such as titanium, stainless steel, steel and aluminum. They possess sufficient strength to withstand bearing and shearing loads and are available with flat and countersunk heads. A conventional HI-Lok has a straight shank with standard threads. Although wrenching lock nuts are usually used, the threads are compatible with standard AN bolts and nuts. To install a HI-Lok, the hole is first drilled with an interference fit. The HI-Lok is then tapped into the hole and a shear collar is installed. A HI-Lok retaining collar is installed using either specially prepared tools or a simple Allen and box end wrench. Once the collar is tightened to the appropriate torque value, the wrenching device shears off, leaving only the locking collar. The Hi-Lok bolts are used for blind attachment of such accessories as fairings, access door covers, door and window frames.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Olympic-Loks

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Olympic-lok blind fasteners are light weight, mechanically-locking spindle-type blind rivets. They come with a lock ring stowed on the head. When the Olympic-lok is installed, the ring slips down the stem and locks the centre stem to the outer shell. These blind fasteners require a specially designed set of installation tools. Olympic-lok rivets are made with three head styles. 1. 2. 3.

Universal 100 degree flush 100 degree flush shear

Rivets diameters of 1/8, 5/32 and 3/16 inch are available in eight different alloy combinations of 2017-T4, A-286, 5056 and Monel. When Olympic-loks were first introduced, they were advertised as an inexpensive blind fastening system. The price of each rivet is less than the other types of mechanical locking blind rivets, and only three installation tools are required. The installation tools fit both countersunk and universal heads in the same size range.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) CherryMax

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The CherryMax rivet is economical to use and strong enough to replace solid rivets, size for size. The economic advantage of the CherryMax system is that one size puller can be used for the installation of all sizes of Cherrymax rivets. A CherryMax rivet is composed of five main parts: 1. 2. 3. 4. 5.

a pulling stem a driving anvil a safe-lock locking collar a rivet sleeve a bulbed blind head

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) CherryMax rivets are available in both universal and countersunk head styles, the rivet sleeve is made from 5056 monel, and inconel 600. The stems are made from alloy steel, CRES and inconel X-750,. The ultimate shear strength of CherryMax rivets ranges from 50 KSI to 75 KSI. ( Kip (K) 1000pounds / sq inch )The CherryMax rivets can be used at temperatures from 250oF to 1400oF. They are available in diameters of 1/8, 5/32, 3/16 and 1/4 inches and are also made with an oversize diameter for each standard diameter listed.

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Removal of Mechanical-Lock Rivets

To remove mechanical-lock rivets, first file a flat spot on the rivet’s centre stem. Once this is done, a centre punch is used to punch out the stem so the lock ring can be drilled out.

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With the lock ring removed, tap out the remaining stem, drill to the depth of the manufactured head, and tap out the remaining shank. All brands of mechanical-lock blind rivets are removed using the same technique.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Lockbolts Lockbolts are manufactured by several companies and conform to Military Standards. These standards describe the size of a lockbolt’s head in relation to its shank diameter, as well as the alloy used. Lockbolts are used to assemble two materials permanently. They are lightweight and are strong as standard bolts. They are three types of lockbolts used in aviation, they are:

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1. 2. 3.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M the pull-type lockbolt the blind-type lockbolt the stump-type lockbolt

The pull-type lockbolt has a pulling stem on which a pneumatic installation gun fits. The gun pulls the material together and then drives a locking collar into the grooves of the lockbolt. Once secure, the gun fractures the pulling pin at its break point. The blind-type lockbolt is similar to most other types of blind fastener. To install a blind lockbolt, it is placed into a blind hole an installation gun is placed over the pulling stem. As the gun pulls the stem, a blind head forms and pulls the materials together. Once the materials are pulled tightly together, a locking collar locks the bolt in place and the pulling stem is broken off. The stump-type lockblolt is installed in places where there is not enough room to use the standard pulling tool. The stump-type lockbolt is installed using and installation tool similar to that use to install Hi-shear rivets.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Lockbolts are available for both shear and tension applications. With shear lockbolts, the head is kept thin and they are only two grooves provided for the locking collar. For tension lockbolts, the head is thicker and four or five grooves are provided to allow for higher tension values. The locking collars used on both shear and tension lockbolts are colour coded for easy identification.

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Issue 1 Revision 0 Jan 2011 Page 174

DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Hi-Lites

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The Hi-Lite fastener is similar to the Hi-Lok except that it is made from lighter materials and has a shorter transition from the threaded section to the shank. The elimination of material between the threads and shank give an additional weight saving with no loss of strength. Hi-Lite’s main advantage is its excellent strength to weight ratio. Hi-Lites are available in an assortment of diameters ranging from 3/16 to 3/8 inch. They are installed either with a Hi-Lok locking collar or by swaged collar like the Lockbolt. The shank diameter is not reduced by stretch torquing.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

CherryBucks

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The CherryBuck is one-piece special fastener that combines two titanium alloys which are bonded together to form a strong structural fastener. The head and upper part of the shank of CherryBuck is composed of 6AL-4V alloy while Ti-Cb alloy is used in the lower shank. When driven, the lower part of the shank forms a bucktail. An important advantage of the CherryBuck is the fact that it is a one piece fastener. Since there is only one piece, CherryBucks can safely be installed in jet engine intakes with no danger of foreign object damage.

Taper-Lok

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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Taper-Loks are the strongest special fastener used in aircraft construction. Because of its tapered shape, the Taper-Lok exerts a force on the conical walls of a hole, much like a cork in a wine bottle. To a certain extent, a Taper-Lok mimics the action of a driven solid shank rivet, in that it completely fills the hole. However, a Taper-Lok does this without the shank swelling. When a washer nut draws the Taper-Lok into its hole, the fastener pushes outward and creates a tremendous force against the tapered walls of the hole. This creates radial compression around the shank and vertical compression lines as the metals are squeezed together. The combination of the these forces generate strength unequalled by any other type of fastener.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Hi-Tigue

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The Hi-Tigue fastener has a bead that encircles the bottom of its shank and is a further advancement in special fastener design. This bead preloads the hole it fills, resulting in increased joint strength. During installation, the bead presses against the side wall of the hole, exerting a radial force which strengthens the surrounding area. Hi-Tigue fasteners are produced in aluminium, titanium and stainless steel alloys. The collars are also composed of compatible metal alloys and are available in two types, sealing and non-sealing. As with Hi-Loks, Hi-Tigue fasteners can be installed using an Allen key and open end wrench.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Jo-Bolts

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The Jo-Bolt is a high-strength, blind structural fastener that is used on difficult riveting jobs’ when access to one side of the work is impossible. They are used in close-tolerance holes or where Jo-Bolts may be required for weight-saving advantages. The hole for a Jo-Bolt is drilled, reamed and countersunk before the Jo-Bolt is inserted and held tightly in place by a nose adapter of either a hand tool or power tool. A wrench adapter then grips the bolt’s driving flat and screw it up through the nut. As the bolt pulls up, it forces a sleeve up over the tapered outside of the nut and forms a blind head on the inside of the work. When driving is complete, the driving flat of the bolt breaks off.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Removal of Special Fasteners Special fasteners that are locked into place with a crushable collar are easily removed by splitting the collar with a small cape chisel. After the collar is split, knock away the two halves and tap the fastener from the hole. The removal techniques of certain special fastener are basically the same as those used for solid shank rivets. However, in some cases, the manufacturer may recommend that a special tool be used. Removal of Taper-Loks, Hi-Loks, Hi-Tigue, and Hi-Lites requires the removal of the washnut or locking collar. Both are removed by turning them with the proper size box end wrench or a pair of vise-grips. After removal, a mallet is used to tap the remaining fastener out of its hole. To remove a Jo-Bolt, begin by drilling through the nut head with a pilot bit followed by a bit of the same size as the bolt shank. Once the nut head is removed, a punch is used to punch out the remaining portion of the nut and bolt.

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Issue 1 Revision 0 Jan 2011 Page 181

DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2) Threaded Rivets – Rivnuts

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Rivnuts, are tubular rivets internally threaded and counterbored and used with matching screws. They are applied blind, and they are used where nut plates cannot be installed. An example of such a location is the leading edge on wings where deicing boots are attached. Rivnuts are made in two head styles: flat and countersunk heads with open or closed ends. The keyed Rivnut is used as a nutplate, and Rivnuts keys are used for blind riveting where torque loads are not imposed. Closed-end Rivnuts are used when a sealed installation is required. To install a Rivnut, a hole is drilled in the skin to accommodate the Rivnut, and a special cutter is used to cut a small notch in the circumference of the hole. This notch locks the Rivnut into the skin to prevent it from turning when it is used as a nut. A Rivnut of the proper grip length is then screwed onto the puller and inserted into the hole with its key aligned with the keyway cut in the hole. When the handle of the puller is squeezed, the hollow shank of the Rivnut upsets and grips the skin. The tool is then unscrewed from the Rivnut, leaving a threaded hole that accepts machine screws for attaching.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE FASTERNERS (DCAM 6.5 L2)

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Issue 1 Revision 0 Jan 2011 Page 183

DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

6.6 PIPES AND UNIONS Rigid Fluid Lines

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Many fluid lines used in early aircraft were made of copper tubing. Copper tubing proved troublesome because it became hard and brittle from the vibration encountered during flight and eventually failed. When working on an aircraft that has copper tubing, the tubing should be annealed each time it is removed. Copper lines must be regularty inspected for cracks, hardness and general condition. Aircraft plumbing lines usually are made of metal tubing and fitting or flexible hose. Metal tubing is widely used in aircraft for fuel, oil, coolant, oxygen and hydraulic lines. Aluminium alloy or corrosion resistant steel tubing has replaced copper tubing. The high fatigue factor of copper tubing is the chief reason for its replacement. Inspection of copper tubing for cracks, hardness, brittleness and general condition should be accomplished at regular intervals to reduce failure. Corrosion resistant steel tubing does not have to be annealed for flaring or forming, in fact, the flared section is somewhat strengthened by the cold working and strain hardening during flaring process. Corrosion resistant steel is also used in areas that are subject to physical damage from dirt, debris, and corrosion caused by moisture, exhaust fumes and salt air. Repairs or replacement to aircraft tubing must be made of the same size and materials that are the same as the original or an approved. All tubing is pressure tested prior to initial installation. One way to ensure that a replacement is made of the same material is to compare the code markings on the replacement tube to those on the original.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Identification of Materials Before making repairs to any aircraft plumbing, it is important to make accurate identification of plumbing materials. The size of rigid tubing is determined by its outside diameter in increments of 1/16 inch. A tube diameter is typically printed on all rigid tubing.

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Another important size designation is wall thickness, since this determines a tube’s strength. The outside diameter, wall thickness is generally printed on the tube in thousandths of an inch.

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It is difficult to determine whether a material is carbon steel or stainless steel or whether it is 1100, 3003, 5052-0 or 2024-7 aluminium alloy. It may be necessary to test sample of the material for hardness by filing or scratching with a scriber.

The magnet test is the simplest method for distinguishing between the annealed austenitic and the ferrite stainless steel. The austenitic types are nonmagnetic unless heavily cold worked, whereas the straight chromium carbon and low alloy steel are strongly magnetic.

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Fabricating Rigid Tubing

It is necessary to replace a rigid fluid line, to obtain a replacement tube assembly from the aircraft manufacturer or fabricate a replacement in the shop. When cutting a new piece of tubing, always cut it approximately 10 percent longer than the tube being replaced. This provides a margin of safety for minor variations in bending. A tube cutter is most often used on soft metal tubing such as copper, aluminium or aluminium alloy. However, they are not suitable for stainless-steel tubing because they tend to work harden the tube. After the tube has been cut and deburred, blow it out with compressed air to remove metal chips that could become imbedded in the tube. When carrying out a steel ball test on rigid tubing, the ball should be 80% of the diameter of the rigid tube

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Tube Bending

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Some applications require rigid lines with complex bends and curves. When duplicating these lines, it must be able to produce bends that are 75 percent of the original tube diameter and free of kinks. Any deformation in a bend affects the flow of fluid. To help reduce the chance of making a bad bend, there are several charts that illustrate standard bend radii for different size tubes. The information on these charts should be adhered to closely. Tube forming consists of four processes: 1. Cutting

When cutting tubing, it is important to produce a square end, free of burrs. Tubing may be cut with a tube cutter or hacksaw. The cutter can be used with any soft metal tubing, such as copper, aluminium or aluminium alloy. A new piece of tubing should be cut approximately 10 percent longer than the tube to be replaced to provide for minor variations in bending. After cutting the tubing, carefully remove any burrs from inside and outside the tube. When performing the deburring operation use extreme care that the wall thickness of the end of tubing is not reduced or fractured. Very slight damage of this type can lead to fractured flares or defective flares which will not seal properly. If a tube cutter is not available, or if tubing of hard material is to be cut, use a fine-tooth hacksaw, preferably one having 32 teeth per inch.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

2. Tube Bending

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The objective in tube bending is to obtain a smooth bend without flattening the tube. Tubing under one-fourth inch in diameter usually can be bent without the use of a bending tool. For larger sizes, use a hand tube bender. Bend the tubing carefully to avoid excessive flattening, kinking, or wrinkling. Tubing with flattened, wrinkled or irregular bends should not be installed. The radius blocks are so constructed that the radius at bend will vary with the tubing diameter. The radius of bend is usually stamped on the block. 3. Tube Flaring

Two kinds of flares are generally used in aircraft plumbing systems, the single flare and the double flare. Flares are frequently subjected to extremely high pressures, therefore, the flare on the tubing must be properly shaped or the connection will leak or fail. If a flare is not made properly, flaws cannot be corrected by applying additional torque when tightening the fitting. The flare and tubing must be free from cracks, dents, nicks, scratches, or any other defects. The flaring tool used for aircraft tubing has male and female dies ground to produce a flare of 35 º to 37 º. Under no circumstances is it permissible to use an automotive type flaring tool which produces a flare of 45 º.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

a.

Single Flare

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M IMPACT-TYPE FLARING TOOL

ROLL-TYPE FLARING TOOL

Roll-type flaring tools are quite popular in aviation maintenance shops because they are entirely self-contained and produce a good flare. A typical roll-type tool can flare tubing from 1/8 to ¾ inch outside diameter. The flaring cone is then turned into the end of the tube and rollers in the cone burnish the metal as it expands into the die. When the flare is formed, the handle is reversed to release the dies, and the tube is removed from the tool. Single flares must be made to certain tolerances. Both the diameter and the radius of the flare must be within specified ranges to ensure a durable, leak-free connection.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

b.

Double Flare

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Soft aluminium tubing with an outside diameter of 3/8 inch or smaller can be double-flared to provide a stronger connection. A double flare is smoother and more concentric than a single flare and therefore, provides a better seal. A double flare is more durable and resistant to the shearing effect of torque. The double-flare is a piece of tubing out of which the flaring is the same manner as a single-flare, remove all burrs. Insert the tubing into the flaring die to the depth allowed by the stop pin and then clamp the die. Insert the upsetting tool into the die and with a few blows of a hammer as possible, upset the tubing. Once the flare is started, insert the flaring tool and strike it with a hammer to fold the metal down into the tubing and form the double flare.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Flared Tube Fitting Flared fitting are identified by either AN (Army / navy) or MS (Military Standard) number. Since AC (Air Corps) fitting are still used in some older aircraft, it is important to be able to identity the differences in fittings.

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Example:

An AN (Army/Navy) fitting has a shoulder between the end of the threads and the flare cone. Another difference between AC (Air Corps) and (Army/Navy) fittings includes the sleeve design. The AN (Army/Navy) sleeve is noticeable longer than AC (Air Corps) sleeve of the same size. Flared-tube fittings are made of aluminium alloy, steel or copper base alloys. For identification purposes, all AN (Army / Navy) steel fittings are coloured black, and all AN (Army / navy) aluminium alloy fittings are coloured blue. The AN 819 (Army /Navy) aluminium bronze sleeves are cadmium plated and are not coloured. AN (Army/Navy) fitting come in a variety of shapes and sizes, each with a specific use.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Flareless Fitting The heavy wall tubing used in some high-pressure systems is difficult to flare. For these applications, the flareless fitting is designed to provide leak-free attachments without flares.

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Although the use of flareless fittings eliminates the need to flare the tube, a step referred to as presetting is necessary prior to installation of a new flareless tube assembly. Presetting is the process of applying enough pressure to the sleeve to cause it to cut into the outside of the tube. The MS (Military Standard) flareless-tube fittings are finding wide application in aircraft plumbing systems. This type fitting eliminates all tube flaring, yet provides a safe, strong, dependable tube connection. The fitting consists of three parts a body, a sleeve and a nut. The body has a counterbore shoulder, against which the end of the tube rests. The angle of the counterbore causes the cutting edge of the sleeve to cut into the outside of the tube when the two are joined. To preset a flareless fitting, lubricate a nut and sleeve, sometime called a ferrule, and slip them over the end of a tube. Next, screw the nut onto the presetting tool, making sure the tube is square against the bottom of the tool. The final tightening depends upon the tubing.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Rigid Tubing Installation

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Before installing a line assembly in an aircraft inspect the line carefully for nicks, scratches, dents and ensure all nuts and sleeves are snugly mated and securely fitted by proper flaring of the tubing. The line assembly should be clean and free of all foreign matter. Never apply sealing compound or anti-seize to a fittings sealing surface since these surfaces depend on metal-to metal-contact to seal. Before securing a line assembly in place, be sure that it is properly aligned. Since rigid line expands and shifts when pressurized, an installation that is under tension is undesirable.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Never pull an assembly into alignment by tightening the nut. Over tightening a fitting may damage the sealing surface, or weaken the flare and sleeve junction, fitting should always be installed to the specified torque using a torque wrench. After all connections are made, the system should be pressure tested. If a connection leaks, some manufacturers allow the nut to be tightened an additional 1/6 turn.

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Over-tightening a flareless tube nut drives the cutting edge of the sleeve deeply into the tube, causing the tube to be weakened to the point where normal inflight vibration could cause the tube to shear. After inspection (if no discrepancies are found), reassemble the connections and repeat the pressure test procedures.

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CAUTION: Do not in any case tighten the nut beyond 1/3 turn (two flats on the hex nut) this is the maximum the fitting may be tightened without the possibility of permanently damaging the sleeve and nut.

When carrying out pressure test on a flexible hose, the testing pressure should be 11/2 times the maximum operating pressure of the hose

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Common faults are: 1. 2. 3. 4. 5. 6. 7.

Flare distorted into nut threads Sleeve cracked Flare cracked or split Flare out of round Inside of flare rough or scratched Fitting cone rough or scratched Threads of nut union dirty, damaged or broken

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Some manufacturers service instructions will specify wrench torque values for flareless tubing installations.

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Support Clamps

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Support clamps are used to secure fluid lines to the aircraft structure or to assemblies in the engine nacelle. These clamps prevent chafing and reduce stress. The two clamps most commonly used are the rubber-cushioned clamp and the plain clamp. The rubber-cushioned clamp secures lines which are subject to vibration. The clamp’s rubber cushion reduces the transmission of vibrations to the line and prevents chafing. Areas subject to contamination by fuel or phosphate ester type hydraulic fluid, cushioned clamps utilizing Teflon are used. Although these do not provide the same level of cushion, they are highly resistant to deterioration. The plain clamp is used in areas that are not subject to vibration and typically consists of a metal band formed into a circle. Identification of Fluid Lines

Large aircraft contain plumbing systems for many different types of fluid. It is important that each line be clearly identified. This is generally accomplished by marking tubing with colour bands, symbols or writing. The symbols are generally printed on one-inch wide tape or decals and secured at regular intervals along a line. On lines four inches or larger in diameter or those subject to extreme temperature, steel tags are used instead of marking tape. In areas where there is possibility that tape, decals, or tags may be drawn into the induction system, paint is used.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

In addition to color bands, some lines carrying fuel are marked with the word “FLAM”. This identifies the lines as carrying a flammable fluid. Lines carrying fluids that are physically dangerous such as oxygen, nitrogen, or Freon are marked “PHDAN”. Additional markings are sometimes provided to identify a line’s function. These include PRESSURE, RETURN, DRAIN and VENT.

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Tapes and decals are placed on both ends of a line and at least once in each compartment through which the line runs. Identification markers are places immediately adjacent to each valve, regulator, filter or other accessory within a line.

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Flexible Fluid Lines

Flexible fluid lines are used extensively on aircraft to connect stationary parts to moving parts and in areas of high vibration. Aircraft systems operate with different fluids under a wide range of pressures. To identify the type of hose that is compatible with each fluid and strong enough to contain its pressure.

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Flexible hose construction generally consists of an inner liner covered with layers of reinforcement to provide strength, and outer cover to protect from physical damage. The materials and manufacturing process of each layer determine the suitability of a specific hose for a particular application. The Inner Lines

The inner lines of a flexible hose carries the fluid and, therefore, must have a minimum porosity and be chemically compatible with the material being carried. The liner must be smooth to offer the least resistance to flow, and remain flexible throughout an entire of operating temperatures. There are basically four different synthetic compounds used in the construction of the inner liner: 1. NEOPRENE is a form of synthetic rubber that is abrasion resistant and is used with petroleum based fluids. 2. BUNA-N is a synthetic rubber compound that is used to carry petroleum-based products. BUNA –N is better suited to carry petroleum products than neoprene. 3. BUTYL is a synthetic rubber compound made petroleum raw materials. BUTYL is excellent as an inner liner for fluid lines carrying hydraulic fluids such as skydrol. 4. TEFLON is the Dupont trade name for Tetrafluoroethylene resin. Teflon has an extremely broad operating temperature range (-65 ºF to + 450 ºF) and is compatible with nearly every liquid used.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Reinforcement Layers The reinforcement layers placed over an inner liner determine the strength of a hose. Common reinforcement layers are made of cotton, rayon, polyester fabric, carbon-steel wire, or a stainless steel wire braid. Hose has a tendency to increase in diameter and decrease in length when pressure is applied, the design of the reinforcement is critical.

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A protective outer cover, usually made of rubber-impregnated fabric or stainless steel braid, is put over the reinforcement to protect the hose from physical damage. In areas of high heat the outer cover is often designed as an integral fire-sleeve to provide extra protection. The outer cover of almost all aircraft flexible hose is marked with a lay line, which consists of a yellow, red, or white stripe running the length of the hose.

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A stripe, the information needed to identify the hose MIL – SPEC number, the manufacturer’s name or symbol, the dash number representing the hose size and in some cases, the manufacturer’s part number along with the year and quarter the hose was manufactured. When a hose is installed property, the lay lines runs straight with no twists.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Types of Flexible Hose Aircraft hoses are manufactured to meet a variety of applications. The types of hose are normally classified by the amount of pressure they are designed to withstand. These include low-pressure, medium-pressure, and high-pressured.

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Low-Pressure Hose

Low-pressure rubber hoses have a seamless inner tube and reinforcement made of a single layer of cotton braid. An outer cover of ribbed or smooth rubber is used to protect the reinforcement from physical abrasion

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The types of hose are normally classified by the amount of pressure they are designed to withstand under normal operating conditions. Low – pressure, any pressure below 250 p.s.i – fabric braid reinforcement.

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Medium-Pressure Hose

A medium-pressure hose is used with fluid pressures up to 3,000 p.s.i.; maximum operating pressure varies with its diameter. Medium-pressure hoses have a seamless inner liner with one layer of cotton braid and one layer of stainless-steel reinforcement. If the hose is used with skydrol or hydraulic fluid, the inner liner is made of synthetic Butyl rubber and the outer braid is coloured green with SKYDROL written on it. High-Pressure Hose

All high-pressure hose has a maximum operating pressure of at least 3000 p.s.i and uses a synthetic rubber liner. This inner liner is wrapped with two or more steel braids as reinforcement. Most high-pressure hose is black with yellow lay line. A hose designed to carry Skydrol has a Butyl rubber inner liner and a green outer cover with a white lay line.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

Flexible Hose Installation Before installing a hose assembly, verify that the aircraft manufacturer specifies a flexible hose is appropriate. Check for proper type and length, physical damage and cleanliness. Ensure that the hose assembly dates are within limits. Part number and assembly date of hose assemblies are found on the hose identification tag. It is important that the lay line be straight when the hose is installed.

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Any spiralling is an indication that the hose is twisted and is under an undue amount of strain when there is pressure in the line. Hose must be supported at least 24 inches, closer support are desirable. A flexible hose must never be stretched tightly between two fittings.

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When carrying out a steel ball test on rigid tubing, the ball diameter should be 90% of the diameter of the rigid tube

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The minimum bend radius for flexible hose is determined by the type of hose being used its size. Bends that are too sharp reduce the bursting pressure of flexible hose. Flexible hose must be protected from wear caused by abrasion or extreme heat. Example: If a fluid line must pass near a hot exhaust manifold, the line must be protected with a suitable fire shield.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE PIPES AND UNIONS (DCAM 6.6 L2)

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PIPELINE INDENTIFIES BY SYMBOLS, WORDS AND COLOURS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

6.7 SPRINGS

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These are usually made of metal, but can be made of composite material. A spring is designed to perform in an elastic fashion, i.e. to deform under a load or force and return to its original size after the removal of the load or force. In general if a force is applied to a spring it will deform and if the force is double the deformation is doubled.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

Hooke’s Law Up to the elastic limit the strain (change in length) of an elastic body is proportional to the applied stress (force). Springs are designed to:   

Absorb energy - to convert say kinetic energy to strain energy as in some shock absorbers. Apply a definite force - e.g. a valve spring to close a valve. Provide a comparator - the spring on a spring balance. Provide an elastic pivot or guide.

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Terms Used

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Free Length: This is the length of the spring without any load applied. When checking this length it should be within the limits as laid down in the appropriate maintenance manual. Pitch: This is the distance between the centre of one coil of the spring and its adjacent coil - without any load applied. Coil Distance: This is the distance between two adjacent coils - without any load applied Wire Diameter: The diameter of the wire from which the coils are made.

Outside Coil Diameter: The outside diameter of the unloaded spring (OCD). Inside Coil Diameter: The inside diameter of the unloaded spring (ICD). Mean Coil Diameter: The average between the OCD and the ICD.

Tip Thickness: The thickness of the ground section of the end of the spring.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

Compression Springs

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These are coil springs and may be right hand or left hand wound. The coil section may be of round or square cross section and the coil diameter is usually large compared to its free length. Usually has ground ends. Tension Springs

Again, these are coil springs and may be left hand or right hand wound. The coil diameter is usually smaller compared to its free length and the coils are usually of round cross section. The ends of the spring are finished in such a way as to provide for end attachment. The spring may be finished with a single hooked end or the coil diameter may be reduced locally to accept a ball ended hook.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

Flexural Springs

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These are designed to provide springiness in any direction.

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Torsion Springs

These are similar in construction to a compression spring, but are designed to rotate about its own longitudinal axis to provide for torsional movement.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

Springs may be designed in several different forms: 1) 2) 3) 4)

Helical - very common. Beam spring - Absorbs a great amount of energy but has limited movement. Leaf spring - Similar in principle to a beam spring except that it is thinner and is usually built up of several leaves. Special - e.g. special cupped spring washers - one placed on top of another over a central guide pin - to make up a stack of any length.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M BEAM SPRINGS

HELICAL SPRINGS

LEAF SPRINGS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE SPRINGS (DCAM 6.7 L2)

Materials The materials that springs are made of must exhibit the property of elasticity. In general materials can include:     

Carbon steel - hardened and tempered. Alloy steels. Nimonic alloys. Titanium alloys. Composites - rare.

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Maintenance

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In most cases springs are checked for serviceability and any unserviceability is usually rectified by replacement. Checks include: An inspection for corrosion, damage, wear, broken coils and distortion. Checking for correct free length of coil springs. Compression springs can be checked using a vernier calliper and tension springs are normally in their fully closed state unloaded. Check for "springiness". This may require a special process using masses and checking the extension/change in length with each added mass. A graph is plotted of mass against change in length from which the elasticity of the spring is ascertained. The spring should return to its free length condition when unloaded.

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PART 66 CAT B1.1 MODULE 6 AIRCRAFT MATERIALS BEARINGS (EASA 6.7 L2)

6.8 BEARINGS A bearing is any surface which supports and reduces friction between two moving parts. Typical areas where bearings are used in an aircraft engine include the main journals, crankpins, connecting rod ends, and accessory drive shafts.

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A good bearing must be composed of material that is strong enough to with stand the pressure imposed on it, while allowing rotation or movement between two parts with a minimum of friction and wear.

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For a bearing to provide efficient and quiet operation, it must hold two parts in a nearly fixed position with very close tolerances. Bearings must be able to withstand radial loads, thrust loads or combination of the two.

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There are two ways in which bearing surface move in relation to each other. One is by the sliding movement of one metal against another and the second is for one surface to roll over another. The three different types of bearing typically used in aircraft reciprocating engines include the plain bearing, the ball bearing and the roller bearing. Plain Bearing

Plain bearings are generally used for crankshaft main bearing, cam ring and camshaft bearings, connecting rod end bearings, and accessory drive shaft bearings. These bearings are typically subject to radial load only, however, flange-type plain bearing are often used as thrust bearings in opposed reciprocating engines. Plain bearings are usually made of nonferrous material, having no iron metals such as silver, bronze, aluminium and various alloys of copper, tin or lead. One type of plain bearing consists of thin shells of silver-plated steel, with lead-in plated over the silver on the inside surface only. Smaller bearings such as those used to support various accessory drive shafts, are called bushings.

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PART 66 CAT B1.1 MODULE 6 AIRCRAFT MATERIALS BEARINGS (EASA 6.7 L2)

Ball Bearings

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A ball bearing assembly consists of grooved inner and outer races, one or more sets of polished steel balls and a bearing retainer. The balls of a ball bearing are held in place and kept evenly spaced by the bearing retainer, while the inner and outer bearing races provide a smooth surface for the balls to roll over. Some races have a deep groove that matches the curvature of the balls to provide more support and allow a bearing to carry high radial loads. Because the balls of a ball bearing offer such a small contact area, ball bearings have the least amount of rolling friction. Because of their construction, ball bearings are well suited to withstand thrust loads and are, therefore, used as thrust bearings in large radial and gas turbine engines. Many of these bearings are prelubricated and sealed to provide trouble- free operation between overhauls. It a sealed ball bearing must be removed or replaced, it is important that used the proper tools to avoid damaging the bearing and its seals.

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PART 66 CAT B1.1 MODULE 6 AIRCRAFT MATERIALS BEARINGS (EASA 6.7 L2)

Roller Bearings

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M RADIAL TAPERED ROLLER BEARINGS

NEEDLE ROLLER BEARINGS

Roller bearings are similar in construction to ball bearing except that polished steel rollers are used instead of balls. The rollers provide a greater contact area and a corresponding increase in rolling friction over that of a ball bearing. Roller bearings are available in many styles and sizes, but most aircraft engine applications use either a straight roller or tapered ROLLER bearing. Straight roller bearings are suitable when the bearing is subjected to radial loads only. Example: Most high- power aircraft engines use straight roller bearings as crankshaft main bearings.

Tapered roller bearings have cone-shaped inner and outer races that allow the bearing to withstand both radial and thrust loads. Examination of Bearings

Ball bearings and roller bearings should be closely examined for smoothness and freedom of movement. Visually inspect a bearing, feel the bearing parts carefully to detect any roughness, flat spots on balls or rollers, and dents or corrosion on the races. In addition, check for pitting, scoring and galling on the outside surfaces of races. Pitting on a thrust bearing race that cannot be removed by polishing with cloth or other mild abrasive usually requires part replacement.

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PART 66 CAT B1.1 MODULE 6 AIRCRAFT MATERIALS BEARINGS (EASA 6.7 L2)

Journal Bearings

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Check journal bearings for damage such as galling, burning, scoring, spalling, misalignment, or an out- of-round condition. Bearing inserts such as bushings and plain bearings are usually replaced, however, looking at them could help to detect wear on their mating surfaces or mounting bosses. Scratching and light scoring of aluminium bearing surfaces in the engine is usually acceptable if the damage is within the limits stated engine manufacturers overhaul manual. The presence of other defects could require rejection of the part even if falls within specific tolerance limits. To property handle bearings, lint- free cotton or synthetic rubber glover are used to keep the acids, oils and moisture on our hands from contaminating any bearing surface. Each bearing must be cleaned in a separate container filled with fresh cleaning solvent or white spirit. Shop cleaning rats and vapour degreasing should not be used because of possible contaminants left from cleaning other parts. Once clean, shop air should never be used to blow bearing dry since moisture in the air supply can corrode the bearing. It is better to use a lint-free cloth or let the bearing air dry. Once dry, immediately lubricate a bearing using the specified lubricant. When the cleaning procedure is completed, the individual bearings must be protected immediately to prevent the onset of corrosion. For temporary storage, the application of mineral oil to the bearing will be sufficient. For long term storage, a compound of lanolin and mineral oil should be used and the bearing wrapped in greaseproof paper.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE TRANSMISSIONS (DCAM 6.9 L2)

6.9 TRANSMISSIONS Gears A gear is a toothed wheel which when meshed with other gears transmits motion from one part of a mechanism to another. The design of the gears determines whether the speed, or the direction, of the motion will be maintained of changed.

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Gear is a machine part with gear teeth. When two gears run together, the one with the larger number of teeth is called the gear. Gear teeth could be manufactured with a wide variety of shapes and profiles.

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All stock gears are made in accordance with the diametral pitch system. The diametral pitch of a gear is the number of teeth in the gear for each inch diameter. Therefore, the diametral pitch determines the size of the gear tooth.

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Power transmission gears are usually made from chromium-molybdenum steel which provides good toughness and resistance to wear. Most gears are run lubricated either by regular maintenance lubrication or by being run semi submersed in oil. Various types of gears transmit power through gearboxes. The type selected for use in a specific application will depend on various factors: 1. 2. 3. 4. 5.

How much power to be transmitted? Is a change of RPM required? Is a change of torque required? Is a change of angle or direction of drive required? Is the gear system to be free from feedback (non-reversible)?

Gears are named according to the angle of intersection of the axis and the shape of teeth.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE TRANSMISSIONS (DCAM 6.9 L2)

Gear Types Spur Gears

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External spur gears have teeth, which point outward from the centre of the gear. Spur gears may have straight teeth, slanting teeth and herringbone teeth. Wheels with slanting teeth operate much more silently then wheels with straight teeth, because at all times several teeth are engaged. Slanting teeth create a force that acts in an axial direction. This force has to picked up by axial bearings. To avoid excessive axial force, the slanting angle of the teeth should not be larger than 20 º.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE TRANSMISSIONS (DCAM 6.9 L2)

Helical Gears

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These are a development of spur gears. Instead of the teeth being parallel to the axis of the gear they lie at an angle (a helix angle in fact) The main advantage of helical gears over straight cut gears is that more teeth area in contact at any one time. Meshing takes place along a diagonal line across the faces and flanks of the teeth. One pair of meshing teeth remain in contact until the following pair engage so the load on the teeth is distributed over a large area. This provides a smoother and quieter drive as well as enabling more power to be transmitted. The disadvantage of helical gears is that they give a heavy axial load to the shaft. This axial load can be eliminated by the use of double helical gearing but can also be absorbed by thrust bearings that support the gear shaft. A double helical gear has two sets of teeth, one with a right hand helix and the other with a left hand helix. Helical gears are stronger, quieter and smoother in operation. Helical gears are only suitable for the transfer of small forces, because the flanks of the teeth contact in only one spot.

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Bevel Gears

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Bevel gears that operate together are two conical surfaces which roll without slipping because of their tooth system.

In a bevel gear, a wheel cannot be exchanged for another with a different number of teeth because the angle in which the teeth are cut or slotted to the body of the wheel changes with the number of teeth. Bevel wheels are manufactured with straight, circular or helical teeth. Those with curved teeth operate with very little noise, but they create a larger axial force than wheels with straight teeth. They are commonly found on intermediate and tail rotor gearboxes on helicopters where a change in the direction of drive is required. They are also used in many gearbox accessory drives at the input stage of the turbine shaft and the accessory drive. Used to change the shaft axis direction and/or change the speed. External bevel gears have pitch angles less than 90 º. Internal bevel gears have pitch angles greater than 90 º.

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Hypoid Gears

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These are used where the centre lines of the two shafts neither intersect nor run parallel to each other. These are similar to bevel gears in application and form, but the basic surface on which they are cut are hyperboloids instead of cones. The teeth are helical and the axes of the shafts do not intersect.

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Worm Gear

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Used where a high reduction in speed and an increase in torque is required. Used on lifting equipment. These conned shafts at right angles which lie on different planes. The worm is essentially a screw which may have a single, double or triple start thread. These engage with teeth on the pinion gear. Older teeth on pinions were straight but now are usually wasted to give a greater contact area with the worm. Worms may be known as encircling worms. With parallel worms the teeth are straight sided on a section through the axis and have the same proportions as standard involutes track teeth. The worm is the drives and the pinion is the driven. Movement cannot be transmitted the other way.

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

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If, for example, the input gear has 25 teeth and the output gear has 75 teeth, then the output speed will be in the ratio of 25:75, or one third of the input speed. Conversely, if the input gear has 20 teeth and the output gear has 10 teeth, then the output speed will be in the ratio of 20:10, or twice that of the input speed.

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Gear trains may be used in a variety of ways, to change the direction of rotation or to increase or decrease the speed of the relevant output gear (and its shaft).

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The design of a gear train will be influenced by the amount of space available to accommodate the desired effect and by the power which is to be transmitted through the gears.

Spur and Pinion Reduction Gear Train

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

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Drive Gear (Pinion)

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Driven Gear (Spur)

Spur and Pinion Reduction Gear Train Fig. 98

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Accessory Unit Drives Aircraft engines also employ multiple gear trains (refer to Fig. 99), in their internal and external gearboxes. These provide the drives for accessories such as fuel, hydraulic and oil pumps, electrical generators, engine speed indicators and many other devices

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Here it can be seen that ‘idler’ gears are added to reverse the rotation and possibly to alter the final ratio of several drives and, while the majority of the gears are of spur and helical configuration, the drive from the engine shaft, to the gearbox, has bevel gears.

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Idler gears 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.

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Meshing Patterns

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

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Chains and Sprockets

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Chains and sprockets provide a strong flexible positive connection in control systems and are generally used where it is necessary to change direction or to connect to a push/pull rod system. These are used where high loads are encountered. E.g.: engine controls, flying control etc. The chain consists of: 1. Two inner and two outer plates 2. Rollers 3. Bearing pins and bushes

The chain has three principal dimensions: 1. Pitch – the distance between the centre of two rollers 2. Roller diameter 3. The width between the inner plates

These dimensions are important for the serviceability of the chain and for its correct fitment around sprocket wheels, pulleys etc. Chain assemblies are supplied from the manufacturer (approved supplier) as complete proof load tested units and no attempt should be made to dismantle riveted links or attachments. Only the bolted or screwed attachments can be disconnected. Any penned nuts and bolts and split pins must be used once only.

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The chain is supplied boxed, lightly oiled and coiled in oil-paper, it is identified by part number and name and should be accompanied by the appropriate stores release documentation. When fittings are connected to the end of the chain they must be fitted in a positive way using locked pins, locked nut and bolt assemblies. The standard for locking a nut, a bolt assembly is to open the bolt end for chains of 8mm pitch or under and use a split pinned lock nut for larger chains (the outer plate of the chain is normally tapped). Change of direction is achieved by the use of sprocket wheels. And the axis of the chain may be changed 90 º by the use of bi-planer block.

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Non-Reversible Chains

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These are the same as the standard chain except that they have extension pieces every other link and they are fitted to sprocket wheels where there is a guard close to the wheel. When fitted to the sprocket wheel, the extension pieces pass around the wheel either side of the wheel. If the chain is tried to be fitted to the wheel the wrong way round the extension pieces will be on the outside – circumference of the wheel and will not pass under the guard. Chains may have handed or non-interchangeable end fittings, this means that together with the chain extension pieces and guard it is impossible to fit the chain incorrectly into the system. Maintenance

Cleaned using paraffin but dried thoroughly afterwards to prevent corrosion and highly oiled. A control chain is checked for stiffness by running the (cleaned) chain over the finger so that each link rotates through 90 º as passes over the finger. The chain is then rotated along its length through 180 º and the process is repeated to rotate each link the other way when the chain is pulled over the finger. If there is a stiff link it will be immediately felt on the finger.

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To check for link wear in a control chain (ie. It will increase its overall length). Accurately measure the length of one link (pin centre to pin centre) using a vernier calliper and multiply the number of links in the length of the chain and compare it with the length of the new chain. This should not exceed that stated in the overhaul manual. Typically 2% maximum elongation is specified. If this figure is reached or exceeded the whole chain should be changed. A control chain is checked for twist by letting it hang, ensuring it is clean and each link articulates freely and sighting (looking) down the chain.

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Belts and Pulleys

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These are used to drive comparatively lightly loaded components such as generators – on some piston engine aircraft and timing mechanisms. The fabric reinforced rubber belt forms a continuous loop around two or more pulleys. Pulleys are called sheaves. On some systems the belt may go around more than one pulley with one being the driver and the others being driven. To maintain tension a spring loaded or adjustable idler pulley may be fitted (normally in the longest run) between the driver and driven pulleys. There are different types of belts and pulleys that may be found in service. For the actual design and maintenance, practices of a particular belt drive system, refer to the belt drive manufacturers manual and /or the AMM (Aircraft Maintenance Manual). Most belt drives are of the “V” type, though there are examples of flat belt drives in use and synchronous belts for applications where it is important that components operate synchronously- cam belts on piston engines for example.

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Flat Belts

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Flat belts are used with flat pulleys with flanges and/or with guides. The flanges or guides are to ensure the belt does not come off the pulley. The flat belt system is cheaper than other belt systems and used where very little load transmission is required. They are of thinner cross section and the specification dimension.

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V- Drive Belts

V-Drive belts are divided into 2 groups – heavy duty and light duty. The V design ensures it sit within the V shaped pulley with no tendency to come off and increases its grip as more tension (power) is applied. The belts are made of rubber or synthetic materials and are strengthened by fabric material, this provides strength in tension and reduces the belts ability to stretch. The rubber provides grip and a wearing surface. It also protects the fabric from moisture and contamination.

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The cross section is shown in the following figure. It is sometimes called Banded Construction. The main tension fabric yarn run longitudinal and the complete belt in enclosed by a fabric covering. Its loading is higher than the flat belt but the radius of the pulleys must not be too small. For smaller pulleys where a reasonably load is required a notched belt should be used. The Notched V belt with the tension fabric plies in the outer section where the tension loads are highest. The belt is designed to take similar loads to the Banded V Belt but will accommodate pulleys of smaller radii. Notched V belts are usually designated with an ‘X’, so a 3V notched belt, for example would be designated a ‘3VX’

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Sizes

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There are there measurements that are used to designate the size of a V belt: 1. Outside Circumference (OC) 2. Effective Length (EL) 3. Pitch Length (PL)

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Outside Circumference (OC) This is measured using a tape measure wrapped around the outside of the belt. It is not very accurate and does not provide a measurement of the belt when under tension (it will stretch slightly under load), which it would be under normal working conditions.

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Effective Length (EL)

This requires a special measuring rig consisting of two pulleys, one fixed and one loadable with an attached measuring scale. To measure the Effective Length (EL) of a belt it is placed around two pulleys with specified groove sizes. One pulley is fixed and the other is designed so it can be loaded to stretch the belt. There is a scale on the loaded pulley to indicate the length between the two pulley centres. Pitch Length (PL)

When the belt bends around a pulley the outside of the belt is in tension and the inside is in compression. Where the centre of the tension occurs is called the neutral axis or tensile chord line. The tensile chord is within the belt (towards to outer edge) and therefore cannot be measured. The pitch length (PL) is the length of the tensile chord around the complete belt.

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Synchronous Belts

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These are similar to flat belts in design except that they are toothed. The teeth are moulded as part of the inner surface and provide a positive drive with no slip. Synchronous belts are used with toothed pulleys and used with timing drives such as ignition systems and valve lifting mechanisms of some piston engines. The use of a synchronous belt system, it connects the tail plane from wheel in the flight deck of the A320 aircraft to sprocket drives under the floor for chain and cable connections back to the tail plane. The system is duplicated.

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Pulleys (Sheaves)

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Pulley (sheaves) are usually made of phenolic or micarta composition, plastic or aluminium alloy and supplied in various diameters and groove angles. Diameters specified include outside diameter and pitch diameter and include groove angles ranging from 32 º to 38 º. It is important that when replacing either a pulley or a belt of any system that it is checked for serviceability and also that is the correct part (check belt markings) Many pulleys/ belts, particularly of the V type construction look very similar and it is important that the Illustrated Part catalogue or Aircraft Maintenance Manual (IPC/AMM) is followed closely and documents such as clearly specify the correct part by name, part number, batch number and serial number.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

6.10 CONTROL CABLES A number of different systems are used to actuate flight and engine controls from the cockpit, flexible control cables are by far the most commonly used method. Multiple–strand control cables are simple, strong and reliable.

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Cable has several advantages over other types of linkage. It is strong and light in weight and its flexibility makes it easy to route.

In addition to primary flight, control cables is used on engine controls, emergency landing gear extension controls, trim tab systems and various other applications.

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Large aircraft have a rather complex automatic tensioning system to keep control cable tension relatively constant as the aircraft contracts and expands. Small aircraft must have their cable tension adjusted as a compromise so they are not too tight when the airplane is hot or too loose when it is cold.

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The cable used in British Aircraft control system is preformed. It complies with British Standards (BS) Specifications. British aircraft cables are graded by loadcarrying capacity (e.g. 10 cwt.) American aircraft cables are also preformed. It complies with American Specifications graded by extreme outside diameter (e.g. 3/32 inch) Cable should always be stored on suitably designed reels. The diameter of the reel barrel should be at least 40 times the cable diameter. Reels should be made of wood which will not corrode the cable and that interior surfaces should be lined with inert waterproof material. Non-flexible Cable

In areas where a linkage does not pass over any pulleys nonflexible cable can be used. It is available in either 1 x 7 or 1 x 19 configurations. The 1 x 7 cable is made up of one strand comprised of seven individual wires, whereas the 1 x 19 consists of one strand made up of 19 individual wires. Nonflexible cable is available in both galvanized carbon steel and stainless steel. Flexible Cable

Flexible steel cable is made up of seven strands of seven wires each is called 7x7 or flexible cable and is available in 1/6 and 3/32 inch sizes in both galvanised carbon steel and stainless steel. Both types are preformed which means that when the cable is manufactured each strand is formed into a spiral shape. This process keeps strands together when the cable is wound and also helps prevent the cable from spreading out when cut. Preforming gives cable greater flexibility and relieves bending stresses when the strands are woven into the cable.

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Extra-Flexible Cable

The most widely used, 7 x 19 is available in sizes from 1/8 inch up. It is extra flexible and is made of 133 individual wires wound in seven strands, each strand having 19 wires. There cables are preformed and are available in both galvanized and stainless steel. Galvanized cable is more resistant to fatigue than stainless steel, but in applications where corrosion is a factor, stainless steel is used.

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Swaged Terminals

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The cable fittings used most in large aircraft manufacture are MS-type swaged cable terminals. To install these terminals, cut the cable using mechanical methods, using cable cutters or heavy duty pliers and insert it into the end terminal. Use either a hand or power swaging tool to force the metal of the terminal down into the cable. To ensure that a terminal is properly swaged, a measurement is made of the swaged terminal with go/ no-go gauge. The swaging process must decrease the terminal’s diameter to the extent that the go end of go/no-go gauge over the swaged terminal, but the no-go end does not. An inspection aid to ensure the cable does not pull out of the terminal, a small mark of paint is placed over the terminal end onto the cable. A broken paint mark indicates the cable has slipped inside the terminal. Inspection should be carried out on completion of the swaging operations as follows:

1. Check that the correct combination of cable and fitting has been used. 2. Re- check the diameter of the swaged shank, using a go/ no-go gauge or a micrometer. If the diameter of the fitting is too small, it has been overswaged, the cable and the fitting must be rejected. Excessive work hardening of the fitting will cause it to crack and might also damage the cable. 3. Check, by means of the inspection hole or paint mark, that the cable is correctly engaged in the end fitting. 4. Check that the swaging operation has not disturbed the lay of the cable, where the cable enters the end fitting. 5. Ensure that the shank is smooth, parallel and in line with the head of fitting and the swaged length is correct. 6. Proof-load the completed cable assembly in accordance with the appropriate drawing. 7. Inspect the fittings for cracks using a lens of 10x magnification or carry out a crack detection test, using magnetic or dye processes as appropriate.

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Proof-Loading

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All cables must be proof-loaded after swaging or splicing by subjecting the cable to specified load. The purpose of proof-loading is both to ensure that the end fittings are satisfactorily installed, and to pre-stretch the cable. British practice is to load the cable to 50% of its declared minimum breaking strength. American practice is to load the cable to 60% of its declared minimum breaking strength. A test rig suitable for proof-loading cables or other similar methods would be acceptable. Before Proof-Loading

A cable with swaged end fittings should be painted with a quick drying paint as its entry into the fitting and allowed to dry. Cracking of the dried paint during proof loading will indicate slipping of the cable resulting from an unsatisfactory joint. The test should consist of slowly applying the specified load maintaining this load for a minimum specified period. (Normally 30 seconds for swaged fitting, but up 3 minutes for splices) Releasing it and carefully examining the cable for sign of pulling out of the end fitting, or stretching of the splice. The end fitting should be checked for cracks using an electro-magnetic method. If the fitting is of stainless steel, a penetrant dye process. The length of the completed cable assembly should be measured after proof loading. Cables with different types of end fitting, or loops should be measured according to appropriate drawings or specifications. Check that the cable assembly is the correct length and ensure that any required identification marking, including evidence of proof loading has been carried out, and that any specified protective treatment has been applied.

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Note :

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The first swaged fitting in a production batch is usually sectioned after proof loading, so that the interior surface can be examined for cracks. If this check is satisfactory, the settings on the swaging machine should be noted, and used for completion of the batch.

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Nicopress oval sleeve / “TALURIT” Cable Splice

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Many light aircraft use Nicopress sleeves that are squeezed onto control cable to form terminal ends. A nicopress sleeve is made of copper and has two holes to accommodate a control cable. When a cable is wrapped around an AN 100 thimble and properly squeezed with the correct Nicopress squeezed, the terminal develops at least the strength of the cable.

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Turnbuckles Turnbuckles are a type of cable fastener that allows cable tension to be adjusted. A complete turnbuckle assembly consists of two ends, one with right-hand threads and the other having left-hand threads, with a brass barrel joining them. Minor cable adjustment is made by rotating the turnbuckle which effectively lengthens or shortens the cable’s length. To ensure that a turnbuckle develops full cable strength, there must be no more than three threads of either end sticking out of the barrel. After cable tension is adjusted, the turn burn buckle barrel is safe tied to the two cable ends so that it cannot turn.

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General Rules for Cable Inspection Cables shall be replaced under the following conditions:

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1. If there is any sign of fraying within three inches of either side of any fairlead or pulley when the cable is at its extreme travel in either direction. 2. 7 x 19 cables shall not have more than six, and 7 x 7 cables not more than three wires broken in any one inch outside the limits specified. 3. 7 x 19 cables shall not have more than 12 and 7 x 7 cables not more than six wires in any one inch are reduced in thickness by more than 50% of the original gauge thickness. 4. Cables must be replaced if they are kinked or corroded. Light surface corrosion that may be easily cleaned off does not necessitate cable replacement.

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Cable Tensiometer

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The cable tensiometer is an instrument which is used to check the tension of cables installed in aircraft control system. The operation of this instrument is similar to that of a torque wrench. The cable tension can be altered and rechecked with the meter until the desired tension is achieved. Cable tensiometers are not read directly off the face of the meter, a calibration chart being provided to convert meter reading to cable tensions in pounds. Three risers are supplied with each meter and are identified numerically for use with various size cables. Cable tensiometers are forwarded to the Instrument Overhaul Section (Standard Room) for calibration and checking at six monthly intervals. Before taking a reading, check the zero reading of the meter and also that the correct riser is being used for the size of cable to be tension – checked.

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No weight or side forces should be applied to the cable or meter when the check is being carried out. For accurate readings, the meter should be placed at least six inches away from turnbuckle, pulley or cable tension regulator. The controls and control surfaces of a cables system should be locked in the neutral position before any reading are taken. Cable Inspection

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When inspecting control cables pay particular attention to those sections of cable that pass through fairleads and around pulleys. To properly inspect each section which passes over a pulley or through a fairlead, remove the cable from the aircraft to the extent necessary to expose that particular section. Examine cable for broken wires by passing a cloth along the length of the cable. This cleans the cable as well as detects broken wires if the cloth snags on the cable.

When snags are found, closely examine the cable to determine the full extent of the damage. Replace flexible and nonflexible cables when the individual wires in each strand appear to blend together, or when the wires are worn 40 to 50 percent.

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

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M A Simple Elevator Control Run Fig. 107

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In a flight control system, pulleys are used to: a) change the direction of operation of the control cables. b) give support on long straight runs of the cable.

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

Aircraft pulleys are manufactured from various materials, such as aluminium alloy, Teflon, fluorocarbon resins, phenolic and other plastic materials.

 

The pulleys bearings are usually of the sealed type and require no lubrication. The pulley is bounded to the bearings in such a manner that the bearing cannot be removed.



A cable guide (or Retainer) is fitted to the pulley to ensure that the cable remains on the pulley.



When adjusting a control, it is important to ensure that the cable end 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 2°.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

Inspection and Maintenance of Pulley All pulleys in a flight control system should be examine or check for wear, cracks and alignment. If a pulley is worn or cracked to an appreciable extent, it should be replaced. The pulleys should turn freely when the control cables are moved. If a pulley is out alignment, it will cause wear to both the pulley and the cable.

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The mountings for such pulleys should be corrected and the cable carefully examine for wear. Typical pulley wear patterns are shown in Figure 112. Avoid contamination between pulleys and harmful substances such as hydraulic fluid, aircraft fuel, paint stripper, etc.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M Figure 112 : Typical Pulley Wear Patterns

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

The Bowden Cable Control System The system is used for lightly loaded controls (selector valve operation, parking brake operating cable etc) and relies on the cable working in tension only, with return being by a spring usually fitted at the component end. The flexible conduit is fixed at both ends which means that the cable system can be routed around bends (so long as they are not too sharp).

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Cables

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These are made of non-corrodible high tensile steel wire, unlike cables fitted to flying control systems. However, they are much smaller in diameter. Conduit

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The conduit consists of a close coiled wire designed to keep the cable system stiff and takes mainly compressive loads. This is covered with cotton braiding followed by a waterproof polymer coating. To give support at the ends and to prevent fraying, metal end-caps are fitted. On some installations rigid metal conduit is used on straight runs.

Fig. 25 BOWDEN CABLE

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

Cable Nipples These are made of brass. The conduit and cable is made up to the correct length (the cable end is tinned to prevent unraveling) and the metal end-caps are fitted over the cable and onto the conduit.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 26: NIPPLES

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The nipple is soldered onto the cable. The nipple recess is tinned. The cable is then passed through the nipple so that the end shows level with the top surface of the recessed end of the nipple. The strands of the cable are then unravelled as far as possible within the recess and the recess filled with molten solder. When the solder hardens the nipple is firmly attached to the cable. In some cases the cable may be swaged into the nipple using a special nipple and swaging machine.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

End Fittings These are usually levers and handles. They may be fitted with adjustable stops so that the range of movement can be set to those specified in the AMM. To fit the cable to an end fitting the AMM must be consulted, but in general terms the follow applies to systems that employ nipple type connections to both ends:

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1. Adjust both end fittings to give the greatest range of movement to each. 2. On those conduits that are adjustable for length, adjust them to their shortest length. (Some conduits have a turnbuckle type adjuster part way down their length which will adjust the length of the conduit but not the cable. The cable passes straight through the adjuster.) This means that there is more 'slackness' in the system in this condition than would otherwise be the case. It will allow easier fitting of the nipples. 3. Align the cable so that the nipple will pass into the fitting hole and the cable will pass through the cable slot (cable rotated to 90° to its normal position). 4. Move the control cable through 90° so that the control cable is now laying in its correct orientation with the metal end fitting of the conduit resting on the fixed part of the end fitting. 5. Carry out the same procedure at the other end of the system. This may require a higher level of motor skills because there is less slack in the cable system because the other end has taken up some of the free play between the cable and the conduit. 6. Adjust the conduit length adjuster to take up the slack in the conduit, which means increasing its length. Make sure the adjuster is in safety and correctly locked. 7. Ensure that both conduit metal end-caps are firmly in place at their respective ends - input end and component end. 8. Check for correct sense of movement, e.g. if it is a throttle system, pushing the throttle forward increases engine power. 9. Adjust the stops at the input end and the component end to give the correct range of movement (check the AMM). It is usual to adjust the stops at the input end so that they control the range of movement - but check the AMM.

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a) b) c) d) e)

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Check for free movement. Check the lay of the cable assembly. Ensure all adjusters are in safety and correctly locked. Carry out a full functional check. Record all the work done and sign.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

The Teleflex Control System

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Fig. 27: BOWDEN CABLE CONNECTION TO PARKING BRAKE LEVER

This uses a lightly loaded cable system moving inside a fixed rigid conduit that will transmit both a tensile (pull) load and a compressive (push) load. This means, for example, that a lever in the flight deck can be used to input a load in either direction to operate a remote device such as a hydraulic selector valve. There is no spring return as in the case of Bowden Controls for example.

The system uses wheel units where the helix winding of the cable engages with a toothed wheel and as the cable moves back and forth so the wheel is rotated. Rotation is limited by the amount of travel of the cable which is up to about 4in (102mm). Sliding end fittings (with a swivel joint) may be used in place of a wheel unit where a. linear movement is required.

The conduit must be supported at regular intervals and may have quick release break units fitted for ease of dismantling.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 20 GENERAL LAYOUT OF A TELEFLEX CONTROL SYSTEM

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

Cable

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 21: TYPES OF CABLE

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These may be of various designs but shown in figure 21 is a number 2 and a number 300 type cable. (See manufactures literature for further types). They have helix windings of opposite hand, are not interchangeable, each having there own fittings. The cable will take a tensile and compressive load with the core cable taking the tensile load and the compression windings taking the compressive load (the type 2 suitable for higher compressive loads). The helix winding is designed to be threaded into an end fitting. Conduit

These are made of aluminium alloy, steel, or tungum (a copper alloy). The conduit should be supported every 3ft (0.9m) but clamp supports should not be fitted where the conduit curves. Clamp Blocks

These are fitted on straight sections to support the conduit.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M For Training Purposes Only

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

Connectors These are used to connect one section of conduit to another. There are several types:   

Nipple type - similar to flare-end hydraulic pipe-line connections but without the olive. Clamp type - this clamps the two conduits together as a butt joint. Quick break type - these allow for the disconnection of the system for component removal etc, and the re-assembly of the joint without having to setup the system again. The cable joining fittings consist of machined rods with interlocking slotted ends attached to the end of each cable.

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Wheel Units

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These consist of a housing in which a "threaded" wheel engages with the helix winding of the cable. They allow for conversion of linear movement to rotary movement and vice-versa. There are several types including the:

   

Single entry type Straight lead type Junction box type 90° and 180° types

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

The cable enters/leaves the unit via a conduit connector and in the case of the single entry unit the cable must have a minimum engagement (at its extreme end of travel) as laid down by the equipment manufacturer/AMM.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Sliding End Fittings

These are used where the linear movement of the cable is not converted to rotary movement. A sliding end fitting is attached after a swivel joint and the assembly is used to move levers etc.

End Fittings

These are fitted to the end of the push/pull rod which is connected to the lever arm of a sliding end fitting or to an arm fitted to the rotating shaft of a control unit. Some push/pull rods will have an end fitting at both ends. They are adjustable for length and have ball-end or ball and socket connections.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

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Fig. 23: END FITTINGS

When adjustment is required it is important that the correct range of movement is achieved and that the fitting is in safety (checked by not being able to pass a piece of wire the same diameter as the hole through the inspection hole). The unit should be locked after final adjustment either using the lock-nut, or a tab washer, or locking wire (as per the AMM of course).

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Figure 24 shows how the cable is screwed into the screwed-end fitting which is also screwed into the outer sleeve locking the slider tube, cable and complete end fitting together. When the cable is caused to move, it will move slider tube and end-fitting together. Note - the slider tube is passed through the outer sleeve and over the conduit first with the belled end resting inside the taper of the outer sleeve.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 24: CONNECTION OF CABLE TO END FITTING

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE CONTROL CABLES (DCAM 6.10 L2)

Split Collet Type End Fitting These are fitted direct to the cable for the operation of sliding end fittings .

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M For Training Purposes Only

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

6.11 ELECTRICAL CABLES In the early days the cables used in aircraft were manufactured to a similar standard to those used in the automobile industry. It was soon learnt that these cables didn't stand up to the severe climatic and environmental conditions encountered during aircraft operation and therefore had to be designed specifically for aircraft use. A variety of cable types have been developed, the choice of cable for a particular function will be governed by its purpose and location.

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Requirements

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These are laid down in BCAR's section D, K and G (old system), now JAR 25 (for large aeroplanes), JAR 27 and 29 (helicopters) etc.

Reliability is of prime consideration for aircraft cables since the performance and safety of an aircraft and its occupants is usually dependant on electrically operated systems. Care, therefore, must be exercised during the manufacture of cable looms and circuits and these must be thoroughly tested on completion.

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Listed below are a number of qualities which an aircraft cable should possess. Minimum Weight and Dimensions

A large aircraft may require many miles of electrical wiring and even small reductions in the size and weight of a cable will result in a considerable weight saving, therefore allowing an increased payload. Resistance to Fluids

The likelihood of an aircraft cable encountering a variety of aircraft fluids is high. It is therefore important that aircraft cables are able to withstand the effects of: water, engine oils, hydraulic oils, fuels, solvents, etc. Non-Inflammability

Wiring is necessary in high fire risk areas such as engine nacelles, and APU bays. Such wiring should not cause the fire to spread and for this reason the protective covering should be of self extinguishing material. There has been doubt about Capton wiring in this respect - although it is still in use. During flight many cables will experience a large temperature range and must remain flexible within this range with the insulation remaining in tact. Resistance to Abrasion

An aircraft cable must possess a number of 'physical' qualities and in particular must have high resistance to abrasion (induced by aircraft vibration). Cables should also be physically strong and easily workable.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Electrical Requirements The conducting element must have a low resistance with a low volts drop per unit length and the insulation must have a sufficiently high insulation resistance.

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Current Rating

The normal current rating of a cable can be defined as: "The amount of current it will carry without sustaining a temperature rise sufficient to cause the value of the insulation resistance to deteriorate to an unacceptable level or without exceeding a specified voltage drop per unit length". Earlier cables either had the current rating stamped on the outer sheath or coloured identification related to the current rating.

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However, because a cable's current carrying capacity is influenced by a number of factors other than electrical load current, it is nowadays the practice of cable manufacturers to use a classification based on the American Wire Gauge (AWG).

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Modern aircraft cables have a wire gauge number stamped on the outside. The electrical systems designer will take into account the factors listed below before choosing a cable for a particular job:      

The electrical loading of the cable. The amount of heat generated by neighboring cables. The number of cables in the loom. The ambient temperature of the surrounding air (its location on the aircraft - near an engine for example). Whether the cable is enclosed or in free air. The thermal conductivity of the cable.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Deterioration Aircraft cables are designed to provide the best possible combination of resistance to deterioration caused by extremes of temperature, mechanical damage and contamination by fluids, and in general, are suitable for installation without additional mechanical protection.

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Working conditions and environment, however, may necessitate the provision of extra protection in those places where the cables are exposed to the possibilities of local damage or conditions which could cause deterioration.

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Receipt and Storage of Cables

Prior to delivery, cable ends are sealed to prevent ingress of moisture. The cables are supplied on drums suitably labelled and protected to prevent damage during transit and storage.

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Smaller sizes of cable may sometimes be supplied in wrapped coils. Visual examination of cables on receipt, by nature of the packing, is often restricted to the outer turns. Such an examination is of little value in checking for faults in the cable, therefore, if the condition of the packing, as received, gives rise to doubt regarding the soundness of the cable, it should be returned to the manufacturer. Note. Check the cable part number/batch number and confirm its identification against its documentation/stores release certificate (JAA form 1). Cables should be stored in a clean, well-ventilated store. They should not be stored near chemicals, solvents or oils and, if necessary, protection should be provided against accidental damage. Loose coils, whether wrapped or not, must not be stored so that a heavy weight is imposed on them, since this may cause unacceptable distortion of the insulation or damage to the protective coverings. The ends of cables in store should be sealed against the ingress of moisture by the use of waterproof tape or sealing compound. Handling of Cables

It is important that cables should be handled carefully at all stages of storage and installation.

When taking long lengths of cable from a drum or reel, the cable should not be allowed to come in contact with rough or dirty surfaces. Preferably the drum or reel should be mounted so that it can rotate freely, but heavy drums may need some means of control over rotation. Care should be taken to remove the twist out of each turn of cables drawn from loose coils, otherwise severe kinking, with consequent damage to the cable, may occur.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Made-up Cabling Cable looms and cable runs made-up on the bench should be inspected before installation in the aircraft. Check for the following: a) Ensure that all cables, fittings, etc, are of the correct type, have been obtained from an approved source, have been satisfactorily tested before making up and have not deteriorated in storage or been damaged in handling. b) Ensure that all connectors and cable looms conform to the relevant AMM, Wiring Diagram Manual, or Modification Drawing, with respect to terminations, length, angle of outlets, and orientation of contact assemblies, identification, and protection of connections. c) Ensure that all crimped joints and soldered joints have been made in accordance with the relevant AMM, Wiring Diagram Manual or Modification Drawing. They should also be clean and sound, and insulating materials should not be damaged in any way. d) Ensure that cable loom binding and strapping is secure. e) Ensure that continuity, resistance and insulation tests are carried out. f) Ensure that all cables should be identified using the correct aircraft wiring code in accordance with the wiring diagram.

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Identification marking may be carried out by printing on sleeves and attaching sleeves at the end of each cable run, or the cable may be printed on at regular intervals along its length. If direct cable marking uses a heat marking system, then the cable must be inspected to check that the insulation has not been damaged and an insulation check must be carried out. Many looming shops have special machines that will automatically mark the cable along its length at regular intervals with the identification - at the same time carrying out insulation tests etc. Installation of Cabling in Aircraft

Guidance on the factors requiring special attention during the installation is given in the following paragraphs – but check the AMM first. Contamination: To prevent moisture from running along the cables and seeping into the associated equipment, the cables should be so routed as to run downwards away from the equipment. Where this is not possible, the cable should incorporate a descending loop immediately before the connection to the equipment. Where conduits, tubes or ducts are used, they should be installed in such a way that any moisture accumulating in them will be able to drain safely away. Cables which are routed through such fittings should be capable of withstanding any such moisture as may be encountered. Interference: Interfering magnetic fields may be set up by electrical equipment, electrical currents in the cabling, or the aircraft structure, and also by magnetic materials. Cables are required, therefore, to be installed so as to reduce electrical interference to a minimum and to avoid interaction between the different electrical services.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

NOTE. Requirements for the avoidance of compass and radio interference are given in Chapter J4-1 of British Civil Airworthiness Requirements. (Now JAR 23 - light aircraft, JAR 25 - large aeroplanes, JAR 27 & 29 -helicopters) Protection of Cabling: The cables are required to be protected from abrasion, mechanical strain and excessive heat and against the deleterious effects of fuel, oil and other aircraft fluids, water in either liquid or vapour form and the weather. Cables should be spaced from the skin of the aircraft so as to reduce the effect of the high skin temperatures likely to be reached in the tropics. The cables should not be run near the hot parts of an engine or other hot components unless a cooled-air space or heat barrier is provided.

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Where cables are routed through metal fittings or bulkheads etc, the edges of the holes through which they pass must be radiised and smoothed and fitted with an insulated bush or sleeve. Cables which are drawn through holes or tubes must be an easy fit requiring only a moderate, steady pull, care being taken to keep the cables parallel to one another and to avoid the formation of kinks (which may cause fracture). Conduits, ducts and trunking used for carrying cables should have smooth internal surfaces.

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Cables being fitted through pressure bungs should be fitted into the correct size holes for the size of cable, to ensure efficient sealing. Only the recommended cable threading tool should be used for this purpose to avoid damaging the bung. Support of Cabling: The cabling must be adequately supported throughout its length, and a sufficient number of cable clamps must be provided for each run of cable to ensure that the unsupported lengths will not vibrate unduly. Bends in cable groups should riot be less than eight times the outside diameter of the cable group. However, at terminal blocks, where the cable is suitably supported at each end of the bend, a minimum radius of three times the outside diameter of the cable, or cable bundle, is normally acceptable. Cables must be fitted and clamped so that no tension will be applied in any circumstances of flight, adjustment or maintenance, and so that loops or slackness will not occur in any position where the cables might be caught and strained by normal movement of persons or controls in the aircraft, or during normal flying, maintenance or adjustment. Where it is necessary for cables to flex in normal use, eg connections to retractable landing gear, the amount and disposition of slack must be strictly controlled so that the cable is not stressed in the extended position, and that the slack will not be fouled, chafed, kinked or caught on any projection during movement in either direction. Cables should normally be supported independently of, and with maximum practicable separation from, all fluid and gas carrying pipelines. To prevent contamination or saturation of the cables in the event of leakage, cables should be routed above rather than below liquid carrying pipelines. Cables should not be attached to, or allowed to rub against, pipelines containing flammable fluids or gases.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Cable Types The pages at the back of this section give information on various types of cables to be found on aircraft. You would not be required to remember the details but you should understand the information that is given.

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All Cables: Cables and equipment should meet the requirements laid down in BCARs and JARs to provide electric shock protection to personnel as well as heat protection - if equipment gets hot during normal operation.

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Airframe Cables: Used for runs throughout the airframe.

Interconnecting Cables: These are used for the interconnection of equipment within racks. Therefore, their insulation is thinner than normal airframe cabling. They are lighter and more flexible.

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Equipment Wire: Sometimes known as “wire” it is used within equipment and is therefore flexible and suitable for soldering. It is not designed as interconnecting wiring though some aircraft manufacturers do use it for this in protected parts of the airframe. Fire Resistant Cables: This type of cable is required to retain a defined level of resistance in certain fire or overheat conditions. The cable is classed as Fire Resistant if able to withstand 1100°C for 5 minutes, and Fire Proof if able to with withstand the same temperature for 15 minutes (JAR 25 & JAR 1 - if close to the outside of a firewall should not suffer damage if firewall heated to 1100°C for 15 minutes). Fireproof Cables: These cables are required to operate for 15 minutes in a designated zone defined in BCARs and JAR 1 and are used in designated fire zones. Cable Maintenance

The requirements, laid down by the CAA for the installation of electrical cables, are laid down in BCARs section J and JARs 23, 25, 27, 29. Only the cables as specified in the AMM, or approved equivalents, should be used. This will ensure that the cables will be capable of taking the voltages (during operation and testing) and the maximum current in the most adverse conditions, without damage to the cables.

Cable Identification

Cables are identified by the manufacture of the cable and further identified by the aircraft manufacturer - to comply with the wiring diagrams.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Cable Manufacturers Identification Each manufacturer will stamp its identification code on the cable at regular intervals along its length. It may include: a) b) c)

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For example: Minyvin GBx XX X 22 (1) (2) (3) (4) (5) a. b. c. d. e.

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The cable size. The manufacturers name. The manufacturer’s code and cable name.

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Manufacturers name of the cable Country of origin Manufacturers cable code Year of manufacture Cable size

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 1 AIRCRAFT WIRING DIAGRAM - EXAMPLE

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Installation Identification Besides the identification of the cable by the manufacturer there is a requirement to identify the cable in the aircraft installation. During aircraft manufacture a cable is installed (suitably routed, supported and connected -crimped etc). Prior to assembly the cable is marked with a code that identifies it and relates it to the aircraft wiring diagram.

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The code - made up of a series of letters and numbers - may be printed on sleeves which are placed on the cable ends prior to being made up - or more likely - printed on the cable length itself.

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The printing may be carried out by a small heated hand operated machine. It is ribbon fed and prior to cable marking is set up with the correct numbers and letters (cable code). These are found by reference to the appropriate aircraft wiring diagram.

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The cable may be marked by being put through an automatic identification and testing machine - once set up this will pull the cable through and print the code on at the required intervals. Any cable faults found the machine will stop and give an aural warning. It will stop automatically at the end of the cable run. It is important that the cable is coded at both ends and at any point where it passes through bulkheads, seals, etc. Most cables are coded at regular intervals (say 2 ft - 0.6m). Always visually check the cable insulation for damage after identing as the ident may have penetrated the insulation and exposed the conducting core. (Fires have been caused by this, so it is important to check carefully end reject the cable if found). This is why the automatic identing machines carry out an insulation test at the same time as the identing procedure. The code will identify such things as: (a) Cable size (b) Circuit (c) Circuit function (d) Cable number

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

The code may be devised by the aircraft manufacturer or may be based on the ATA 100 specification system. An example of this is shown below. I EF G B 22 (1) (2) (3) (4) (5)

NMSV (6)

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1. Unit number, used where components have identical circuits. 2. Circuit function letter and circuit designation letter which indicate circuit function and the associated system. 3. Cable number, allocated to differentiate between cables which do not have a common terminal in the same circuit. Generally, contacts of switches, relays, etc, are not classified as common terminals. Beginning with the number one, a different number is given to each cable. 4. Cable segment letter, which identifies the segment of cable between two terminals or connections, and differentiates between segments of the circuit when the same cable number is used throughout. Segments are lettered in alphabetical sequence, excluding the letter I and O. A different letter is used for each of the cable segments having a common terminal or connection. 5. Cable size. 6. Suffix data, used to indicate the type of cable and to identify its connection function. For example code NMS V indicates nyvin metsheath (a BICC cable) ungrounded cable in a single-phase system.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

The recommendation is that the cable is coded at regular intervals along its length and it is most important that it corresponds to the appropriate aircraft wiring diagram. When replacing cables it is important to: a) b) c) d) e)

Fit the correct replacement cable. Correctly route and support the cable. Ensure its correct identification along its length. Employ the correct terminations. After replacement carry out appropriate electrical tests followed by a functional test.

For certain electrical systems, cables are required to perform a more specialized function than that of the cables already referred to. Some examples of what are generally termed 'special purpose cables' are described below.

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Ignition Cables (Ignition Harnesses) These are used for the transmission of high tension voltages (high voltages) in both piston and turbine engine ignition systems. They are usually of the singlecore stranded type with a high level of insulation, and screened by metal braided sheathing to prevent interference.

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The number of cables required for a system correspond to the number of spark plugs or igniter plugs as appropriate, and they are generally made up into a complete ignition cable harness. Depending on the type of engine installation, the cables may be enclosed in a metal conduit, which also forms part of the harness, or they may be routed without conduit.

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Cables are connected to the relevant system components by special end fittings comprising either small springs or contact caps secured to the cable conductor, insulation, and a threaded coupling assembly.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Thermocouple Cables

These cables are used for connection of cylinder head temperature indictors and turbine engine exhaust gas temperature (egt) indicators to their respective thermocouple sensing elements. The conducting materials are normally the same as those in the thermocouple sensing element, for example, iron and constantan or copper and constantan for cylinder head thermocouples, and chromel (an alloy of chromium and nickel) and alumel (an alloy of aluminium and nickel) for egt thermocouples. In the case of cylinder head temperature indicating systems, only one thermocouple sensing element is used and the cables between it and a firewall connector are normally asbestos covered. For egt measurement a number of thermocouples are required to be radially disposed around the jet pipe in the gas stream. It is usual practice to arrange the cables in the form of a harness tailored to suit a specific engine installation. The insulating material of the harness cables is either silicone rubber or PTFE impregnated fibreglass. The cables terminate at an engine or firewall junction box from which cables extend to the flight deck indicator. The insulating material of extension cables is normally of the polyvinyl type, since they are subject to lower ambient temperatures than the engine harness. In some applications extension cables are encased in silicone paste within a metal-braided flexible conduit.

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Co-axial Cables (Figures 2 and 3) Co-axial (co-ax) cables contain two or more separate conducting elements -one inner and one outer. The innermost conductor may be solid or stranded copper wire, and may be plain, tinned, silver-plated or even gold-plated in some applications, depending on the degree of conductivity required.

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The outer conductor is made in the form of a circle usually of fine wire braid surrounding the inner core. The two conductors are separated by an insulation usually of polyethylene or Teflon.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 2 CROSS SECTION OF CO-AX CABLE

Outer coverings or jackets serve to weatherproof the cables and protect them from fluids, mechanical and electrical damage. The materials used for the coverings are manufactured to suit operations under varying environmental conditions. Co-axial cables are used for the transmission of low power signals, with the signal line (the inner conductor) protected from unwanted signals (noise) by the outer wire braid. The outer braid provides a shielded against electrostatic and magnetic fields. Any electrostatic field does not extend passed the outer braid and the fields due to current flow in the inner and outer conductor's cancel each other. Also, since co-axial cables do not radiate and fields, then likewise they will not pick up any energy, or be influenced by other strong fields. Co-axial cables are used on radio equipment, for the connection of antennae to receivers/transmitters, and capacitance type fuel quantity indicating systems for the interconnection of tank units to amplifiers. The construction of a typical co-axial cable and also the sequence adopted for attaching the end fitting are shown. The outer insulation covering is cut back to expose the braided outer conductor which is then fanned out and folded back over the adapter. At the same time, the inner insulation is cut back to expose the inner conductor.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 3 CO-AX CABLE & END FITTING

The next step is to screw the sub-assembly to the adapter thereby clamping the outer conductor firmly between the two components. In some cases the outer conductor may also be soldered to the sub-assembly through solder holes. Soldering a contact on to the inner conductor and screwing the coupling ring on to the sub-assembly completes the assembly. Cable Types

The following pages give technical data on a selection of cables. You would not be required to remember the details but you should read and understand the information. You should note the performance rating of the cables, the properties and the identification. You should note the current ratings and how they are affected by being “bunched” and the reasons why.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

MINYVIN

Lightweight flexible airframe wiring cable. Operating temperature -30°C to +105°C. Single- or multi-core screened and sheathed versions are available.

PACKAGING AND IDENTIFICATION: Packaging Cables are supplied on reels labelled in accordance with specification or order requirements and suitably packed for transport to destination. Cable Identification Cables are printed with the cable code, country of origin -GBX (UK); manufacturer • BB (BICC); code letter for year of manufacture, number indicating conductor size, and G221 (BS reference) Colour White Conductor Tinned copper, size range 22 to 12. Silver plated copper alloy size 24 only. SPECIFICATIONS AND APPROVALS These cables are produced to specification BS 2G221.

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SUMMARY OF PERFORMANCE: Voltage rating: 300* volts rms at 1600 Hz Maximum service temperature +105°C Minimum service -75°C for fixed installation temperature -30°C for flexing * Size 24 is 250 volts rated PROPERTIES: Mechanical Resistant to tape abrasion Chemical Resistant to fuels, hydraulic fluids, petroleum and ester based oils, de- icing fluids, fire extinguishing liquids and cleaning solvents, fungus and mildew. Physical Resistant to flame. Cold bend at 30°C. Readily printed.

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Current ratings (MINYVIN) The ratings given in Table 4 are based on a conductor temperature rise of 40°C in an ambient of 65°C. The maximum permissible conductor temperature is 105°C If the ambient temperature (t°C) is continuously in excess of 65°C, the rating must be multiplied by a factor K where

K=

 105  t     40 

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These current ratings are in line with those given for Pren cables in British Civil Airworthiness Requirements, Section J.

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Cable No.

Rating Uninyvin Uninyvinal condition 22

–

20

–

18

–

16

14

12

10

A B C A B C A B C A B C A B C A B C A B C

Rating in amperes (maximum) Single 3 7 12 cable cables cables cables 11 7 5 4 12 8 7 6 15 12 9 9 14 9 7 5 16 12 9 8 22 19 15 15 18 13 10 6 23 17 13 12 30 26 19 18 21 15 11 7 25 19 14 13 33 28 26 25 31 24 17 12 36 28 24 21 50 47 43 42 43 30 22 15 50 38 32 30 72 67 62 60 61 47 36 25 71 56 48 45 110 107 104 101

8

6

6

4

4

2

2

0

1

00

0

000

00

0000

000

–

0000

–

A B C A B C A B C A B C A B C A B C A B C A B C A B C

–

–

8

Table 4: Maximum ratings for cables bunched in free air A = Continuous rating B = 5-minute rating C = 1-minute rating

For Training Purposes Only

65 89 165 87 122 236 120 185 378 155 265 530 165 300 600 185 350 690 210/240* 410 810 235/265* 460 955 270/305* 555 1240

49 82 159 65 115 230 92 175 360 120 250 620 130 290 590 165 340 680 190† 405† 800† 210† 455† 940† 245† 550| 1225†

36 80 153 – – – – – – – – – – – – – – – – – – – – – – – –

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M –

87 105 173 115 152 250 160 225 390 200 305 545 220 330 620 240 370 705 270 420 820 300 470 965 350 570 1255

*The higher rating relates to 2 cables only

†5 cables only

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AIRCRAFT THERMOCOUPLE EXTENSION CABLE

These cables are used for the transmission of thermocouple currents within an operating temperature range -65°C to +260°C. SUMMARY OF PERFORMANCE: Maximum continuous service temperature: +260°C Minimum service temperature (flexing): -55°C PROPERTIES: Mechanical Resistant to tape and scrape abrasion and cut through Low notch sensitivity Low surface creepage Withstands climatic test BS 3 G 100 Service life: 10,000 hours at 260°C Chemical Resistant to fuels, hydraulic fluids, petroleum and ester-based oils, de-icing fluids, fire extinguishing liquids and cleaning solvents, fungus and moulds. Physical Resistant to flame Non-blocking Retains flexibility after ageing Non-cracking when flexed at -50°C Readily printed

PACKAGING AND IDENTIFICATION: Packaging Cables are supplied on reels labelled in accordance with specification or order requirements and suitably packed for transport to destination. Colours Insulation: The nickel chromium core (+ve) is coloured white The nickel aluminium core (-ve) is coloured green Sheath: Size 20 sheath colour is green Size 22 sheath colour Is green with white stripes Conductors Positive, nickel chromium; negative, nickel aluminium; size range 20 and 22

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KPA 150

PACKAGING AND IDENTIFICATION: Packaging Cables are supplied on reels labelled In accordance with specification or order requirements and suitably packed for transport to destination. Cable Identification Cables are printed with the cable code name, country of origin – G. manufacturer – BB BICC, year of manufacture M for 1974 etc. and number Indicating conductor size. white Single core: two core red and blue three core red, blue, yellow Multicore: four core red, blue, yellow, green outer sheath white

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SUMMARY OF PERFORMANCE (KPA 150): Voltage rating: 600 volts rms at 2000 Hz Maximum continuous service temperature 150°C Minimum service temperature (flexing) -65°C PROPERTIES: Mechanical Resistant to tape and scrape abrasion and cut through. Low notch sensitivity. Low surface creepage Withstands climatic tests BS 2 G 100. Service life: 50,000 hours at 150°C Chemical Resistant to all fuels, hydraulic fluids, petroleum and ester-based oils, de-icing fluids, fire extinguishing liquids and cleaning solvents, fungus and moulds. Physical Resistant to flame No smoke emission at twice the operating temperature (300°C) Non-blocking Retains flexibility after ageing Non-cracking when flexed at -65°C Readily printed

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TERSEL PACKAGING AND IDENTIFICATION

Flexible airframe cable. Maintains essential circuits after a fire with an ultimate life of five minutes at 1100°C. Operating temperature -55° to +190°C. Single and multicore, screened and sheathed versions also available.

Packaging Cables are supplied on reels labelled in accordance with specification or order requirements and suitably pecked for transport to destination. Cable Identification Cables are printed with the cable code, country of origin GBX (UK); manufacturer - BB (BICC) code letter for year of manufacture, and number indicating conductor size. Colour Orange Conductor Nickel plated copper, size range 22 to 0000. SPECIFICATIONS AND APPROVALS: These cables are produced to specification BS G189 which is interchangeable with MIL–W–8777.

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SUMMARY OF PERFORMANCE: Voltage rating: 600 volts rms at 1600 Hz. Maximum service temperature +190°C. Minimum service temperature - 75°C for fixed Installation. PROPERTIES: Mechanical Resistant to tape abrasion Chemical Resistant to aviation fuels, hydraulic fluids, petroleum and ester based oils. Physical Resistant to flames, ultimate life of 5 minutes at 1100°C. Flexible throughout temperature range (-55° to +190°C).

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FEPSIL

Fluid resistant airframe wiring cable. Maintains essential circuits after a fire with an ultimate life of five minutes at 1100°C. Operating temperature -75°C to +190°C. Single or multicore, screened and sheathed versions also available.

PACKAGING AND IDENTIFICATION: Packaging Cables are supplied on reels labelled in accordance with specification or order requirements and suitably packed for transport to destination. Cable Identification Cables are printed with the cable code, country of origin -GBX (UK); manufacturer - BB (BICC);code letter for year of manufacture, and number indicating conductor size. Colour Green Conductor Nickel plated copper, size range 22 to 0000 SPECIFICATIONS AND APPROVALS: These cables are produced to specification BS G206 which is interchangeable with MIL-W-8777.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

SUMMARY OF PERFORMANCE: Voltage rating: 600 volts rms at 1600 Hz Maximum service temperature: +190°C Minimum service temperature: -75°C for fixed installation PROPERTIES: Mechanical Resistant to tape abrasion Chemical Resistant to fuels, hydraulic fluids, petroleum and ester based oils, deicing fluids, fire extinguishing liquids and cleaning solvents, fungus and mildew. Physical Resistant to flames, ultimate life of 5 minutes at 1100°C Flexible throughout temperature range (-55 to +190°C)

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EFGLAS

Flexible, abrasion resistant airframe wiring cable, operating temperature from -75°C to +260°C. Screened and sheathed versions are also available. SUMMARY OF PERFORMANCE: Voltage rating: 600 volts rms at 1600 Hz Maximum service temperature: +260°C Minimum service temperature: -70°C PROPERTIES: Mechanical Resistant to tape and scrape abrasion, dynamic cut through Chemical Resistant to fuels, hydraulic fluids, petroleum and ester-based oils, deicing fluids, fire extinguishing liquids and cleaning solvents, fungus and mildew. Physical Resistant to flames and smoke emission, blocking and low temperature cracking. Flexible throughout temperature range (-75°C to +260°C). Resists retraction.

PACKAGING AND IDENTIFICATION: Packaging Cables are supplied on reels labelled in accordance with specification or order requirements and suitably packed for transport to destination. Cable Identification Cables are printed with the cable code, country of origin GBX (UK); manufacturer - BB (BICC); code letter for year of manufacture, and number indicating conductor size. Colour White Conductor Nickel plated copper, size range 22 to 0000 SPECIFICATIONS AND APPROVALS: These cables are produced to specification BS G222 which is interchangeable with AIR 4524 (GROUP 250-280)

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NYVIN

General purpose flexible airframe wiring cable. Operating temperature -30°C to +105°C. Screened and sheathed versions are also available. SUMMARY OF PERFORMANCE: Voltage rating: 600 volts rms at 1600 Hz Maximum service temperature: +105°C -75°C for fixed installation Minimum service temperature: -30°C for flexing PROPERTIES: Mechanical Resistant to tape abrasion Chemical Resistant to all fuels, hydraulic fluids, petroleum and ester based oils, deicing fluids, fire extinguishing liquids and cleaning solvents, fungus and mildew. Physical Resistant to flames. Cold bend at -30°C Readily printed

PACKAGING AND IDENTIFICATION: Packaging Cables are supplied on reels labelled in accordance with specification or order requirements and suitably packed for transport to destination. Cable Identification Cables are printed with the cable code, country of origin -GBX (UK); manufacturer - BB (BICC); code letter for year of manufacture, and number indicating conductor size. Colour White Conductor Tinned copper, size range 22 to 0000. SPECIFICATIONS AND APPROVALS: These cables are produced to specification BS G177

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

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Crimping A crimped connection is one in which a cable conductor is secured by compression to a termination so that the metals of both are held together in close contact. A typical crimp termination has two principal sections, crimping barrel and tongue, together with, in some types, a pre-insulated copper sleeve which mates with the crimping barrel at tone end and is formed, during the crimping process, so as to grip the cable insulation at the other in order to give a measure of support.

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The barrel is designed to fit closely around the cable conductor so that after pressure has been applied a large number of points of contact are made. The pressure is applied with a hand or hydraulically operated tool fitted with a die or dies, shaped to give a particular cross-sectional form to the completed joint.

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The precise form of the crimp is determined by such factors as the size and construction of the conductor, the materials, and the dimensions of the termination. It is, therefore, most important that only the correct type of die and crimping tool should be used, and that the necessary calibration checks have been made to the tool.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M

There is a vast range of terminations available, many of which are colour-coded, and suitable for use only with specific types of aircraft cable. It is, therefore, vital that the appropriate manufacturer's instructions regarding the use of cables and terminations are followed.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 1 TYPICAL CRIMP TERMINATIONS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Only aluminium or bimetal (AlCu) terminations should be used to terminate aluminium cables and the cable should be stripped immediately prior to making the joint. The barrel of some aluminium terminations may contain a quantity of inhibiting compound, others not so filled require that inhibiting compound be applied before crimping takes place.

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Some specifications also require additional sealing after crimping. The compound will also minimise later oxidation of the completed connection by excluding moisture and air.

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Tools

f T o g y n r i r a t e e e i n r i p g o n r E P S A M

These include: * *

Crimping pliers Cable strippers

Both come in a variety of shapes and sizes and the descriptions that follow are typical of some that are available. The AMP Crimping Tool

The special tool used for crimping AMP terminals has several important design features to ensure a consistent quality of completed crimp joints. These include: (1) (2) (3) (4)

Crimp ratchet. Locator. Insulation adjusting pins. Colour and dot coding.

Crimp ratchet. This device ensures the bottoming of the die jaws before the jaws can be opened again. This ensures that the tool cannot be released until a complete crimp is made. Locator. The locator holds the terminal in the correct position in the die jaws and allows the conductor strands to protrude 0.8 mm from the terminal barrel when the wire is fully inserted.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Insulation Adjusting Pins. To allow for small variations in wire size and to ensure optimum mechanical strength of the joint the insulation die head has three degrees of adjustment e.g.: 1. 2. 3.

Tight. Medium. Loose.

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Colour & Dot Coding. The "dot" coding system is needed to identify the terminals which have been crimped in the correct AMP hand tool. If a red terminal is crimped in a red handled tool, a single dot impression will be left on the insulation at the barrel end.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 2 AMP CRIMPING TOOL

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Cable Strippers These are used to cut the insulation away from the conducting part of the cable.

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These may have separate locations in the one tool to be used with different size cables. They may be adjustable to fixed positions to cater for different size cables or they may be adjusted *by trial and error' to obtain the correct amount of cut (into the insulation).

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Which ever cable stripping method is used it is most important that a spare piece of cable is used to practice on first. The 'spare' piece of cable should be the same size and type as the actual cable being worked on, and the strip should be accomplished so that all the insulation is removed but no conducting strands are cut or weakened.

f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Some cable strippers will cut the insulation and pull it off the end in one action, others will cut the insulation only - the fingers being used to remove the insulation. Crimping Procedures

Before carrying out crimping of a termination, the following should be verified: (a) (b) (c) (d) (e)

Correct size and type of wire for the job. Correct size and type of terminal with suitable size crimp barrel to accommodate wires and if necessary, the insulation. Correct crimping tool and associated dies, selected to be compatible with type of terminal and wire size. Correct tool being used. Note that the ratchet and pawl hand type tools will only release on completion of crimping cycle. Correct cable strippers.

Preparation of Wire 1. 2. 3. 4. 5.

Using wiring diagrams, AMM etc choose correct wire, inspect, put on aircraft wiring code, and cut to length - allowing for some error when stripping etc. Using approved stripping tool, remove specific length of insulation. Inspect stripped end for severed or damaged conductor strands. If any found cut off cable to beginning of insulation and re-strip. Insert all conductor strands into barrel. Ensure that no insulating materials enter.

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Conductor strands must be laying together to allow for 100% insertion. If the lay of the strands is disturbed they should be re-imposed with a light twist. Excessive twisting should be avoided as this increases the conductor diameter. Preparation of Tool 1. 2. 3. 4. 5. 6.

Crimping AMP Terminals – Example

7. 8.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M

Always ensure that both insulation adjustment pins are in the same position.

1. 2. 3. 4. 5. 6.

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Insert insulation adjustment pins into the No. 3 position. Locate terminal in crimping jaws. Insert wire into barrel with the insulated part entering the grip portion of the terminal. Close handles slowly and fully until crimp ratchet releases. Open handles, remove terminal and check insulation support as follows: Bend the wire back and forth once, terminal sleeve should retain grip on wire insulation. If wire pulls out set insulation adjustment pins in next tighter position. (No. 2) and re-crimp.

Select the appropriate terminal for the size of wire being terminated and to suit the stud size of the terminal fitting. Select a tool by reference to the colour of the terminal check wire size range stamped on tool face. Inspect the tool for serviceability and adjust the insulation crimping adjustment pins. Insert the terminal into the jaws so that the barrel rests against the locator. Squeeze handles until terminal is lightly gripped by the jaws. Insert prepared wire end into terminal barrel ensuring that all conductor strands enter. When fully inserted the conductor should extend beyond the barrel by approximately 0.8 mm. Hold wire in position and crimp by squeezing handles until ratchet releases. Remove completed crimped joint and inspect for dot code impression.

On completion of crimp, check: 1. 2. 3. 4. 5.

Correctness of form and location of crimp. Adequate insertion of conductor strands in barrel. If insulation support is provided, check correctness of form and location of insular crimp. Check any codification by crimp dies is correct in detail and position. Check joint for freedom from fracture, rough or sharp edges and 'flash'.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Crimping Butt Splices (1) (2) (3) (4) (5) (6) (7) (8) (9)

Select the required Butt Splice and a tool of the same colour coding. Adjust the insulation crimping adjustment pins as detailed above. Insert Butt Splice into crimping jaws until properly located. Squeeze handles until Butt Splice is lightly gripped. Insert prepared wire into terminal barrel. When inserted the conductors should be visible in the inspection window. Hold wire in position and complete crimping operation. Inspect for correct formation of completed crimp. Insert other end of Butt Splice into jaws until properly located. Complete crimping operation by repeating Items 4, 5, 6 and 7.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 3 A TYPICAL CRIMPED JOINT

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 4 'DOT' INDICATOR

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In-Line Crimping

Each barrel of a connector must carry only one cable unless specifically permitted by the CAA.

If in-line crimps are allowed they must be fitted either horizontal or positioned so that an ingress of fluid is impossible. Protective sleeves, additional to the crimp insulation, will not be provided to prevent an ingress of fluid. Care must be exercised to ensure that in-line crimps are only used in positions where the operating temperatures do not exceed the specified limits.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Specific approval must be obtained from the air worthiness authority before incorporating in-line crimps in the following: (1) (2) (3) (4) (5) (6) (7) (8)

Screened cable Coaxial cable. Multicore cable. Cables in excess of size 10. Thermocouple cables. HV cables ie above 250 V rms (eg igniter ht leads, aerial feeders). Cables used in fire-resistant circuits (fire detector and extinguisher circuits within the protective zone). Types of cable, totally enclosed in conduits or ducts, which cannot readily be visually inspected.

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Fig. 5 VOLTAGE DROP TEST ON CRIMPED TERMINATIONS

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Restrictions 1) The use of in-line crimps is currently restricted to cable size 10 (35 amp), or smaller; low temperature (105°C) connectors must not be crimped on size 12 or larger Efglas cable.

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2) Repair schemes are restricted to the following: a) The minimum distance between joints in any one cable must be two feet. b) Not more than two joints are to be made in any ten feet of cable. c) Multiplicity of joints in cables must be avoided, if possible, and in no case must the number exceed the following: (i) Runs up to 20 feet - 3 joints. (ii) Runs up to 200 feet - 5 joints. (iii) Runs over 200 feet - 8 joints.

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3) On installation, wherever possible, observe the following:

a) All joints must be accessible for visual inspection. b) Joints must be positioned so that they do not touch one another or touch duct cable-retaining straps and other fixtures which may set up 'tracking' paths. c) Joints must, if possible, be positioned on the outside of the looms unless special fixing attachments are preferable. All fixing attachments, such as corrugated wrapping strip, must be approved. d) If it is impracticable to accommodate a stagger of joints along a cable run, positive separation, eg using insulation or cable clips, must be carried out.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Erma Hand-Operated Hydraulic Crimping Machine For large size cables various hydraulic crimping machines are available. Described here is the Erma crimping machine. This machine is supplied as a kit containing eight sets of dies for cable size from AWG 6 to AWG 0000, and an alien key used for fitting the dies to the machine. The crimp formed is a regular hexagon shape and has two code letters impressed on it by the dies during crimping. These code letters are HG, HH - HN (for cable sizes AWG 6, 4 - 0000) and are the same as those marked on the cable lugs by the manufacturer.

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Preparation of the Machine

The machine operating handles should be screwed into position and the code letters stamped on the dies checked for size. If the dies are to be changed carry out the following:

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a) Select the two matched dies bearing the correct code letters for the size of cable in use. Check that the lugs to be used have the same code letters marked on the terminal palm. b) Remove the upper die adapter by sliding it from the dovetailed head of the tool. This leaves the slotted head of the tool open to allow the lower die to be fitted to the ram. Insert the spigot on the upper die into the hole in the die adapter until it is held in position by a spring-loaded steel ball. c) Close the hydraulic valve by turning the knob clockwise. Pump the handle a few times to move the ram forwards and disclose the hexagon socket screws which hold the lower die. Slacken these screws using the alien key provided with the kit. Fit the lower die into the ram so that the screws fit into the recesses on either side of the die. Tighten the screws to hold the die, ensuring that they are below the surface of the ram body. Open the hydraulic valve to retract the ram. d) Slide the upper die adapter, complete with die, into the dovetailed grooves until it is located centrally by a spring-loaded steel ball.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 6 ERMA CRIMPING MACHINE

Operation

Check that the two-letter code on the cable lugs and on both dies is correct for the size of the cable to be terminated. (a) Close the hydraulic valve. Place the lug centrally between the dies and pump the handles until the lug is lightly gripped. (b) Strip the cable insulation so that when it is inserted in the lug the insulation lies flush against the end of the barrel and the conductor projects slightly from the other end. (c) Insert the conductor into the barrel of the lug and pump the machine until the dies are fully closed. A safety valve will operate with an audible click and pressure on the pump handle is greatly reduced.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2) (d) Open the hydraulic valve to allow the ram to retract. The crimped termination can then be removed from the machine. Plugs And Sockets Plugs and sockets are provided to ensure a secure connection for one or more circuits. They are designed to prevent entry of moisture and to provide a positive connection for a multi pin system. They are small and have light mass but may be difficult to assemble and are expensive.

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To prevent damage, debris and moisture entry, protective caps are provided and should be fitted at all times other than when the connectors are being worked on and in their assembled condition. During work protection may then be in the form of a linen or plastic bag, totally enclosing the connector and secured to the cables. This temporary protection should only be removed just prior to connection being made in the aircraft.

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Miniature Connectors

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Extreme care should be taken when handling and connecting miniature and sub-miniature connectors. Both plugs and sockets should be checked for any signs of dirt, bent pins or physical damage to the shells before attempting to connect. If connectors will not mate, check the reason, and rectify or renew. On no account should force be used to effect mating. If a bent pin is found, on no account should it be straightened as it will almost certainly fracture. The pin should be removed and a new one crimped into position. Lubrication

Some ranges of plugs and sockets require the engaging threads to be lubricated with a suitable lubricant to ensure that they can readily be disconnected.

Removal of Wired Contacts

There are two basic types of contact (pin) retention used in plug and socket connectors in aircraft, one which the contacts being released for removal from the rear and one where release is from the front using the insertion/extraction tools. Therefore, it is essential that the correct procedures and tools are used for a particular type of plug or socket. 1. Front Release - The contact is removed by pushing from the front of the connector and removing from the rear. 2. Rear Release - The extraction tool enters the connector from the rear of the connector and the contact is also removed from the rear. Multiway connectors, terminal junctions, inline single wire connectors, switches, motors, indicators, instruments and other electrical components; all may now be terminated by a rear release system which requires the use of a few tools and the minimum of operator training.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Contacts crimped with a standard crimping tool are inserted and removed using a single fail-safe plastic tool for each size of contact. The Hellermann Deutsch 460/450 Series Connectors, terminal junction modules and custom-made component termination modules form the central part of the integrated termination system.

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All terminations are inserted and removed by a single expendable plastic tool which is fail-safe in that mishandling will result in damage to the tool rather than to the connector or termination modules.

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Fig. 7 REMOVAL AND INSERTION TOOL

Fig. 8 PIN RETENTION

The tines of the clip snap in behind the shoulder of the contact. The removal tool displaces the tines of the clip sufficiently to allow the contact to be withdrawn rearwards.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

Contact Insertion: 1. Remove the backshell or other accessory from the rear of the connector and move onto the cable loom. 2. Ensure that the correct hole in the connector has been selected to insert the pin - check the hole numbering system on the front face of the plug/socket and the wiring diagram. 3. Snap the coloured end of the appropriate insertion/removal tool on to the wire. When inserting the wire into the tool, use the thumb and not the thumb nail as this could damage the insulation (see figure 9). Position the tool on the contact shoulder, except in the case of size 22 contacts, in which case the tool should be positioned on the back of the crimp bucket. See figure 9. 4. Holding the connector with the rear insert facing you, slowly push the contact straight into the connector. A positive stop will be felt when the contact is locked in by the retention clip. 5. Inspect the contact/pin for correct alignment, straightness, security and the height compared with the other pins. If damaged/not correctly fitted, renew and refit. 6. Carry out continuity test. Assemble plug and socket and test system.

1. 2. 3. 4.

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Contact Removal:

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The removal procedure is similar but the reverse of the insertion procedure. Holding the connector with the rear insert facing you, snap the white end of the appropriate insertion/removal tool over the wire to be removed. Slowly slide the tool along the wire into the connector, until a positive stop is felt. The retention clip will now be unlocked. Press the wire against the serrations of the central section of the tool and withdraw both wire and tool together.

As you can see, to release the contact, you must put the extraction tool over the rear of the contact and down between the contact and clip to release the clip from behind the front shoulder.

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 9 INSERTION OF PIN INTO PLUG

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DCAM PART 66 CAT B1.1 MODULE 6 MATERIALS AND HARDWARE ELECTRICAL CABLES (DCAM 6.11 L2)

This method has had wide usage. Some of the connectors you are likely to use with this feature are Amphenol 246 and 48 series, Bendix PT-SE, Cannon FRF, KPSE, Flight FH, FC Hellermann Deutsch SLPT, DS, Cinch C0909, Pyle National RPL/FPK, ZZ and the AMP/AM series of rack and panel connectors.

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f T o g y n r i r a t e e e i n r i p g o n r E P S A M Fig. 10 FRONT RELEASE

In the case of the rear release, the extraction tool, which is usually plastic, enters from the rear of the connector between the contact and the clip to release the contact. The contact is then pulled out through the rear whilst still in the tool.

Fig. 11 REAR RELEASE

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