EASA PART 66 - MODULE 11.02 - Airframe Structures

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MODULE 11.2 AIRFRAME STRUCTURES

CONTENTS 1

AIRFRAME STRUCTURES – GENERAL CONCEPTS................ 1-3 1.1 1.2

1.3

1.4

1.5

1.6

1.7

2

AIRWORTHINESS REQUIREMENTS FOR STRUCTURAL STRENGTH ..... 1-3 STRUCTURAL CLASSIFICATION ...................................................... 1-3 1.2.1 Primary structure ........................................................... 1-4 1.2.2 Secondary Structure ..................................................... 1-6 1.2.3 Tertiary Structure .......................................................... 1-6 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS ............ 1-6 1.3.1 Fail Safe........................................................................ 1-6 1.3.2 Safe Life........................................................................ 1-6 1.3.3 Damage Tolerance........................................................ 1-7 ZONAL AND STATION IDENTIFICATION SYSTEM................................ 1-9 1.4.1 Zonal System ................................................................ 1-9 1.4.2 Station Identification System ......................................... 1-10 LOADS FOUND W ITHIN THE STRUCTURE – STRESS AND STRAIN ...... 1-11 1.5.1 Compression ................................................................. 1-12 1.5.2 Tension ......................................................................... 1-12 1.5.3 Bending ......................................................................... 1-13 1.5.4 Torsion .......................................................................... 1-14 1.5.5 Shear ............................................................................ 1-14 1.5.6 Hoop Stress .................................................................. 1-15 1.5.7 Metal Fatigue ................................................................ 1-15 DRAINAGE AND VENTILATION PROVISIONS ..................................... 1-18 1.6.1 External Drains ............................................................. 1-18 1.6.2 Internal Drains............................................................... 1-20 1.6.3 Ventilation ..................................................................... 1-20 LIGHTNING STRIKE PROVISION ...................................................... 1-21

CONSTRUCTION METHODS ...................................................... 2-1 2.1

STRESSED SKIN FUSELAGE ........................................................... 2-1 2.1.1 Frames and Formers..................................................... 2-2 2.1.2 Bulkheads ..................................................................... 2-2 2.1.3 Longerons and Stringers ............................................... 2-3 2.1.4 Doublers and Reinforcement ......................................... 2-4 2.1.5 Struts and Ties .............................................................. 2-4 2.1.6 Beams and Floor Structures .......................................... 2-5 2.1.7 Methods of Skinning...................................................... 2-5 2.1.8 Anti-Corrosive Protection .............................................. 2-7 2.1.9 Construction Methods – Wing ....................................... 2-8 2.1.10 Construction Methods – Empennage ............................ 2-9 2.1.11 Construction Methods – Engine Attachments ................ 2-10 2.1.12 Structural Assembly Techniques ................................... 2-12 2.1.13 Solid Shank Rivets ........................................................ 2-12 2.1.14 Special and Blind Fasteners. ......................................... 2-14 2.1.15 Bolts and Nuts............................................................... 2-19 2.1.16 Adhesive Bonded Structures ......................................... 2-24 2.1.17 Methods of Surface Protection ...................................... 2-27 2.1.18 Exterior Finish Maintenance .......................................... 2-29

Issue 1 – Module 11.2 21 Dec 2001

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Issue 1 – Module 11.2 04 Sept 2001

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AIRFRAME STRUCTURES – GENERAL CONCEPTS

1.1 AIRWORTHINESS REQUIREMENTS FOR STRUCTURAL STRENGTH Airworthiness requirements are necessary with respect to aircraft structures, because established standards of strength, control, maintainability, etc. will ensure that all aircraft will be constructed to the safest possible standard. Requirements for aircraft above 5700kg MTWA (maximum total weight authorised) are listed in Joint Airworthiness Requirement 25 (JAR-25) and for aircraft below 5700kg MTWA, in JAR-23. These publications cover not only the basic requirements, like maximum and minimum 'g' loading, but a vast range of other requirements with respect to the structure such as: •

Control Loads

•

Door Operation

•

Effect of Tabs

•

Factor of Safety

•

Fatigue

•

High Lift Devices

•

Stability & Stalling

•

Ventilation

•

Weights

The list is all-embracing and provides a useful means of searching for specific structural details.

1.2 STRUCTURAL CLASSIFICATION

For the purpose of assessing damage and the type of repairs to be carried out, the structure of all aircraft is divided into three significant categories:•

Primary structure

•

Secondary structure

•

Tertiary structure

Diagrams are prepared by each manufacturer to denote how the various structural members fall into these three categories.


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AIRFRAME

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STRUCTURES

In the manuals of older aircraft the use of colour may be found to identify the three categories. Primary Structure is shown in Red, Secondary in Yellow and Tertiary in Green. Note: This system has been discontinued for many years, but with some aircraft having a life of 30 or more years and still being operated, it may still be possible to find the old system in use.

1.2.1 PRIMARY STRUCTURE

This structure includes all portions of aircraft, the failure of which in flight or on the ground, would be likely to cause: •

Catastrophic structural collapse

•

Inability to operate a service

•

Injury to occupants

•

Loss of control

•

Unintentional operation of a service

•

Power unit failure

•

Examples of some types of primary structure are as follows: •

Engine Mountings

•

Fuselage Frames

•

Main Floor members

•

Main Spars

Primary Structure – Engine mountings Figure 1 Page 1-4


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Primary Structure :Wing Spars Figure 2


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1.2.2 SECONDARY STRUCTURE

This structure includes all portions of the aircraft which would normally be regarded as primary structure, but which unavoidably have such a reserve of strength over design requirements that appreciable weakening may be permitted, without risk of failure. It also includes structure which, if damaged, would not impair the safety of the aircraft as described earlier. Examples of secondary structure include: •

Ribs and parts of skin in the wings.

•

Skin and stringers in the fuselage

1.2.3 TERTIARY STRUCTURE

This type of structure includes all portions of the structure in which the stresses are low, but which, for various reasons, cannot be omitted from the aircraft. Typical examples include fairings, fillets and brackets which support items in the fuselage and adjacent areas.

1.3 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS

1.3.1 FAIL SAFE

A fail safe structure is one which retains, after initiation of a fracture or crack, sufficient strength for the operation of the aircraft with an acceptable standard of safety, until such failure is detected on a normal scheduled inspection. This is achieved by part and full scale airframe testing and fatigue analysis by usually by the aircraft manufacturer and by subsequent in-service experience.

1.3.2 SAFE LIFE

Safe life structure and components are granted a period of time during which it is considered, that failure is extremely unlikely. When deciding its duration, the effects of wear, fatigue and corrosion must be considered. For example, if tests show that fatigue will cause a failure in 12,000 flying hours, then one sixth of this might be quoted as the safe life.(2000 hours then scrapped) If wear or corrosion prove to be the likely cause of failure before 12,000 hours, then one of these will be the deciding factor. The safe life time period may be expressed in flying hours, elapsed time, number of flights or number of applications of load, ie; pressurisation cycles.

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1.3.3 DAMAGE TOLERANCE

The fail safe method has proven to be somewhat unreliable following some accidents that proved that the concept was not 100% guaranteed. It was also a severe limitation that the addition of extra structural members to protect the integrity of the structure considerably increased the weight of the aircraft.. The damage tolerant concept, has eliminated much of the extra weight, by distributing the loads on a particular structure over a larger area. This requires an evaluation of the structure, to provide multiple load paths to carry the loading. The main advantage is that even with a crack present, the structure will retain its integrity and that during scheduled maintenance programmes, the crack will be found before it can become critical. For example, a wing attachment to the fuselage, which in the past would have been designed with one or two large pintle bolts, will now have a larger number of smaller bolts in the fitting. The single or dual bolt attachment had to be heavily reinforced to take the wing loading, adding more weight, whereas the multiple load path can be constructed in a lighter manner, whilst still maintaining its strength.

Single Pin Attachment Figure 3


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Multiple Pin Attachment Figure 4

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1.4 ZONAL AND STATION IDENTIFICATION SYSTEM

1.4.1 ZONAL SYSTEM

During many different maintenance operations including component changes, structural repairs and trouble shooting, it is necessary to indicate to the engineer where, within the structure, the correct location is to be found for the work to be carried out. When attempting to establish a specific location or identifying components, some manufacturers make use of two systems, a zonal system and a frame/station method. The zonal system divides the airframe into a number of zones, (usually less than 10), to give engineers and others a rough idea of where they need to look. The zonal system may also be used in component labelling and work card area identification. In the illustration below, an engineer might have for example a work card numbered 500376, indicating it was Job 376 located on the left wing (Zone 500).

Zonal Identification Figure 5 Issue 1 – Module 11.2 21 Dec 2001

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1.4.2 STATION IDENTIFICATION SYSTEM

Most manufacturers use a system of station marking where, for example, the aircraft nose is designated Station 0 and other station designations are located at measured distances aft of this point. Component and other locations within the wings, tailplane, fin and nacelles are established from separate dedicated station’s zero. Fuselage Locations A particular fuselage station (or frame) would be identified, for example, as Station 5050. This means that if the metric system of measurement is employed, the frame is located at 5.05 metres (5050mm) aft of station zero.

Frame Stations Figure 6 Lateral Locations To locate structures to the right or left of the aircraft, many manufacturers consider the fuselage centre line as a station zero. With such a system, the wing or tailplane ribs could be identified as being a particular number of millimetres (or inches) to the right or the left of the centre line. Vertical Locations These are usually measured above or below a ‘water line’, which is a predetermined reference line passing along the side of the fuselage, usually, somewhere between the floor level and the window line.

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1.5 LOADS FOUND WITHIN THE STRUCTURE – STRESS AND STRAIN Aircraft structural members are designed to carry a load or to resist stress and a single member may be subjected to a combination of stresses during flight. When an external force acts on a body, it is opposed by a force within the body. This force is called Stress. If the body is distorted by the stress, it is said to be subject to Strain. Stress and strain can be defined as follows: Stress is load or force per unit area acting on a body. Stress = Load or Force Cross Sectional Area Strain is the distortion per unit length of a body.

Strain = Distortion Original Length

There are five major stresses and all will be found somewhere within an aircraft structure. In the design stage, the stresses will have been assessed by the designer and the structure made strong enough to carry them adequately. Furthermore, a reserve of strength will also have been included for safety. The five types of stress are: 1. Compression 2. Tension 3. Bending (a combination of compression and tension) 4. Twisting/Torsion 5. Shear


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1.5.1 COMPRESSION

Compression is regarded as a primary stress and is the resistance to any external force which tends to push the body together. Compressive stresses applied to rivets for example, expand the shank as they are driven in, completely filling the hole and forming the head to hold sheet metal skins together.

Compression Figure 7 1.5.2 TENSION

Tension is the primary stress that tends to pull an object apart. A flexible steel cable used in flying control systems is an excellent example of a component designed to withstand tension loads only. It is easily bent, has little opposition to compression, torsion or shear loads, but has an exceptional strength/weight ratio when subjected to a purely tension load.

Tension Figure 8 Page 1-12


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1.5.3 BENDING

Bending, when applied to a beam, tends to try to pull one side apart while at the same time squeezing the other side together. When a person stands on a diving board, the top of the board is under tension while the bottom is under compression. Wing spars of cantilever wings are subject to bending stresses. In flight, the top of the spar is being compressed and the bottom is under tension while on the ground, the reverse occurs, the top is in tension and the bottom is under compression. If the wing is supported, the strut will be in tension in flight and in compression on the ground.

Bending Figure 9


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1.5.4 TORSION

A torsional stress is one that is put into a material when it is twisted. When we twist a structural member, a tensile stress acts diagonally across the member and a compressive stress acts at right angles to the tension. A good example is a crankshaft of an aircraft piston engine which is under a torsional load when the engine is driving the propeller.

Torsion Figure 10 1.5.5 SHEAR

A shear stress is one that resists the tendency to slice a body apart. For example a clevis bolt in a flying control system is designed to take shear loads only. It is normally a high strength steel bolt with a thin head and a fat shank. These bolts secure the flexible steel cables to the control surfaces and allow the cable to move with the control surface without bending. The airload on the control surface attempts to slice the bolt apart or shear it.

Rivet Joint in Shear Figure 11 Page 1-14


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1.5.6 HOOP STRESS

An aircraft which has its fuselage pressurised inside to allow the carriage of passengers at altitude, will have other stresses acting on the fuselage skin. The circumferential load about the fuselage is known as hoop stress and resisted by the fuselage frames and tension in the so called stressed skin. The longitudinal (axial) load along the fuselage is also resisted by tension in the skin and by the longerons and stringers.

Hoop stress Figure 12 1.5.7 METAL FATIGUE

The phenomenon of metal fatigue has long been known, but has become of greater concern in recent years with aircraft which remain in service long after their original expected fatigue life has expired. It is relatively easy to design a structure to withstand a steady load, but aircraft are subjected to widely varying loads in flight and many components experience load reversals, an example being the wings, where the aerodynamic forces during flight manoeuvres cause tension and compression loads to alternate continually. Unfortunately, any metal part subjected to a wide variation or reversal of even a relatively small load is gradually and progressively weakened. The subject was vividly highlighted in 1954, with another type of load reversal, that of pressurisation cycles of the passenger cabin. which resulted in a number of disastrous accidents with the De-Havilland Comet airliner. Small fatigue cracks in the fuselage skin accumulated around the corners of the square shaped windows and hatches and led to a fatal explosive decompression of the cabin. Following the incidents the most extensive research to this hitherto unwarranted menace was undertaken, and led to fatigue loading being included into future design considerations. Metal fatigue refers to the loss of strength, or resistance to load, experienced by a component or structure as the number of load cycles or load reversals increases. Issue 1 – Module 11.2 21 Dec 2001

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Load reversals refer to a material being continually loaded and unloaded and as long as the elastic limit is not exceeded, the material should be unaffected and return to its original state. In reality, however, the load application may result in minute, seemingly inconsequential cracks, which, as the cycles continue, get larger and join up with other, newer cracks. Eventually, after many cycles, the cumulative effect will be such that the strength of the metal will be compromised and could result in catastrophic failure. The fatigue strength of a metal can be found by experimentation on full scale fatigue rigs, which can be subjected to a programme of load reversals, 24 hours a day, 365 days a year, to accumulate information and a fatigue life, years ahead of the oldest aircraft of the particular type in the fleet. How the in-service aircraft subsequently consumes this fatigue index, depends on its operating theatre. For example, the number of times the pressurisation cycles are applied to aircraft on long or short haul flights, steep or conventional take off and landing etc., are taken into account to calculate fatigue life consumed. Stress amplitude can be plotted against endurance for one particular value of mean stress, the so-called ‘S/N Curve. Using a chart such as this, it can be determined at what point, in cycles, the metal has reached its minimum acceptable strength. This will be the ultimate fatigue life and is normally allotted a fatigue index of 100.

Fatigue Graph Figure 13 Page 1-16


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Even when the fatigue index of 100 is eventually reached on each individual aircraft, the designers can extend it beyond 100, by examining, as previously mentioned, how the fatigue was consumed and recommending specific structural inspection and possibly strengthening or replacement of fittings and components. Fatigue is a natural phenomenon and cannot be prevented. The ability to correctly predict its effects and take the necessary action is the problem faced by the aircraft design and maintenance personnel. Different metals have different fatigue characteristics and the way parts are designed, also affects their fatigue life. Fastener holes, sharp changes in thickness and small seemingly insignificant cracks for example, can directly affect the fatigue life of a part. Fatigue cracking can also accelerate the onset of corrosion, by exposing unprotected metal to the elements. The crack growth and the consequential increase in corrosion, can cause serious structural problems over a relatively short period. With the ageing of the airliner fleet, a number of extra inspections, including non-destructive testing and structural sampling techniques have been introduced. The maintenance technician must carefully monitor the aircraft structure, paying particular attention to the integrity of surface finish and general corrosion.


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1.6 DRAINAGE AND VENTILATION PROVISIONS

Drainage The aircraft structure requires many different types of drain holes and paths to prevent water and other fluids such as fuel, hydraulic oil etc., from collecting within the structure. These could become both a corrosion and fire hazard. The forms of drainage can be divided into two areas. 1. External drains 2. Internal drains

1.6.1 EXTERNAL DRAINS

These ports are located on exterior surfaces of the fuselage, wing and empennage to ensure fluids are dumped overboard. In small unpressurised aircraft and unpressurised areas of larger airliners, these drains may be permanently open. However, in pressurised aircraft, the cabin air would leak uncontrollably through the drains and so it is necessary to use drain valves to prevent loss of cabin pressure. There are a number basic types of drain valve used for this purpose. Two similar types rely upon pressurised air in the cabin to keep the valve closed. One valve has a rubber flapper seal and the other a spring loaded valve seal. Normally located on the keel of the fuselage, both are open when the aircraft is unpressurised on the ground, allowing the fluids to drain overboard. During flight, the increased air pressure in the cabin closes the valves, thus preventing any pressurisation losses. These valves are shown below, where it can also be seen that a levelling compound has been used in areas which might become fluid traps. This compound is usually a rubberised sealant which fills the cavity, bringing the level up to the lip of the drain hole.

Fuselage Drains Figure 14

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Another similar type of drain valve also uses the cabin air pressure to close off the drain path, this time by moving the plunger down to seal the drain. This valve will also be open when cabin pressure is removed.

Fuselage Drains Figure 15 Fluids from some places, such as galleys and wash basins, require more than simple drain holes. The temperature at cruising altitude can fall to -60°C and water draining overboard could freeze and cause blockage problems. The method used in these cases are drain masts, which are like small aerofoils projecting from the bottom of the aircraft skin, on the centre line, through which the water is discharged. The drain masts are heated to prevent icing and also discharge the liquids well away from the aircraft's skin.

Boeing 747 Drain Masts Figure 16 Issue 1 – Module 11.2 21 Dec 2001

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1.6.2 INTERNAL DRAINS

To enable the external drains to function as designed, means must be provided within the various locations of the airframe and powerplant installation, to ensure that all fluids are directed towards the site of the external drain points. This is achieved by using internal drain paths and drain holes. The internal structure is provided with tubes, channels, dams and drain holes, to direct the flow of fluid towards the external drain points. All structural members are designed so that they do not trap fluids by ensuring, for example, that all lightening holes and ribs face downwards, allowing fluids to run off them.

1.6.3 VENTILATION

It is essential that the internal cavities within the structure are properly vented to prevent the build up of flammable vapour from the drain lines and to allow any other moisture residue to properly evaporate. Consequently sumps, tanks and cavities will all be provided with vent pipes and in some cases, such as engine cowlings, ram air inlets and outlets are utilised to ensure all zones where fluids are contained are adequately ventilated. System Installation Provisions The installation of various systems within the airframe, require adaptations from the perfect ‘drawing-board’ design. When systems like the air conditioning and pressurisation, hydraulic, pneumatic, electrical, avionics and others are designed, there must be facilities incorporated in the plans, to provide a location for all the system components, their associated lines and cables. It must also be borne in mind that many components have to be either serviced ‘in-situ’, or will be a line replaceable unit (LRU), both of which requires easy access for the maintenance engineers. To this end, on modern aircraft, there are normally compartments allocated to each of the major systems where the majority of components will be installed. Thus, it can be possible to find dedicated Avionics bays, Hydraulic bays, Air conditioning bays, etc., all of which allow access for the easier replacement of 'black boxes' (LRU’s) and mechanical components like control units, valves, filters etc,. Older aircraft will still have components scattered throughout the airframe, with difficult access in some places through small panels, all of which will obviously make maintenance on these systems much more difficult.

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1.7 LIGHTNING STRIKE PROVISION When aircraft are flying in cloud or in close proximity to storms, there is always the risk of the aircraft being struck by lightning. Whilst this is a rare occurrence, there are many protection devices installed in the aircraft to ensure that a strike does as little damage as possible when it does happen. A lightning strike on an aircraft can have a peak current of up to 100,000 amperes, so precautions must be taken to ensure that the least damage is done to the aircraft, its systems and components as the charge passes through. Most important is the electrical bonding of all the major components of the airframe. Bonding is achieved by electrically connecting all the components of an aircraft structure together. These precautions will ensure all components are at the same electrical potential by providing a return path through the airframe, since modern aircraft utilise an earth return system. This means that current from the lightning strike cannot build up on one part of the structure and create a voltage high enough to allow it to jump to another part, that might be electrically separated, such as flying control surfaces. Note: Electrical bonding also protects equipment from the build up of static electricity, which is produced as the aircraft collects ions from the atmosphere as it passes through. Bonding cables are referred to as secondary conductors. As well as electrical bonding, dedicated lightning protection systems are employed to cater for the high current and these are usually known as primary conductors. They can be found, connecting system earth returns, as mentioned earlier, connecting power-plants to the airframe and ensuring that all major structural items, (which are often manufactured in different factories in different countries), are properly connected together after final assembly. Occupants of the aircraft are also protected from electrical shock in this way by the surrounding aircraft structure with what is referred to as a ‘Faraday Cage’.

Electrical Bonding Figure 17 Issue 1 – Module 11.2 21 Dec 2001

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CONSTRUCTION METHODS

2.1 STRESSED SKIN FUSELAGE As previously described, a variety of loads act on the airframe during flight. If a proportion of these loads can be carried by the skin covering, the underlying framework can be made lighter without loss of overall strength. In early aircraft, all loads were taken by the framework and the covering of fabric, doped to pull it taught or of thin sheets of wood achieved streamlining, but contributed little or nothing to the strength of the airframe. As aircraft design evolved, the fabric and wood was replaced with aluminium alloy sheet. Because of its extra strength, a large part of the load can be borne by this skin, reducing the weight of underlying structure. This is called Stressed Skin construction and this method also provides a very smooth surface, because the skin is stiff enough not to be distorted by the airflow. With the advent of pressurised cabins the usefulness of a strong skin is evident when considering pressurisation loads. A method of construction where the skin carries all the loads without supporting structure is called pure monocoque construction. A good example of a pure monocoque construction is a chicken’s egg, since it has no internal support, only the egg shell carries the load. In practice, this construction is difficult to achieve, as the skin would have to be so thick, that the extra weight penalty incurred, would severely impair the ability to fly. However, the principle is sometimes used in the construction of composite material external fuel tanks, mainly for military aircraft and even here some internal strengthening is necessary.

Monocoque Construction Figure 18 Issue 1 – Module 11.2 21 Dec 2001

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STRUCTURES

In a stressed skin fuselage construction, about half the loads are carried by the skin and half by the supporting structure. This type of construction is called semi monocoque and its advantage is that the space within the structure is unobstructed and is used for passengers and freight.

Semi-Monocoque Construction Figure 19 2.1.1 FRAMES AND FORMERS

Frames and formers provide the basic fuselage shape, with the frames, being of more robust construction, providing strong points for attachment of other fittings such as the wings and tailplane.

2.1.2 BULKHEADS

Where extra support is required within a fuselage for mounting of components such as wings and landing gear, bulkheads are to transfer the loads to the fuselage structure without producing stress raising points. Bulkheads can be either a complete or a partial circular frame, which usually reinforces a fuselage frame. Other examples are solid pressurisation bulkheads which are normally found at the front of the fuselage ahead of the flight deck and at the rear of the pressure cabin, or an engine firewall on the nacelles..

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2.1.3 LONGERONS AND STRINGERS

Longerons are used in fuselage construction, where either an aperture such as a door or window requires greater support, or where a number of structural high load points such as floors, landing gear attachments, etc. need to be interconnected. They are usually of much heavier construction than stringers and can be solid extrusions or fabricated multiple part construction. Stringers provide longitudinal shape and support to the fuselage skin. They are also the spanwise members of the mainplanes, vertical and horizontal stabilisers and flying control surfaces. Often stringers are attached to frames with fillets or gussets.

Longerons and Stringers Figure 20


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2.1.4 DOUBLERS AND REINFORCEMENT

Where the skin requires extra strengthening, at the junction of plates or around small apertures, a second layer of skin is attached over the original to reinforce it. This extra plate is known as a doubler or a doubler plate. Where loads are concentrated within the structure, it can be strengthened at these places by either making the material thicker, or by the addition of a number of layers of similar material. The actual amount of reinforcement being dictated by the amount of stress carried in each area.

Doubler Plate Figure 21 2.1.5 STRUTS AND TIES

Any structural item that is designed solely to take a compressive load is called a strut. Whereas an item that only takes a tensile load is called a tie. They can be found throughout a modern aircraft structure, although an ideal example would be a high performance biplane. In this type of aircraft often used for aerobatics, the struts which separate the pairs of wings, in compression and the interconnecting flying wires, in tension, take all the loads produced by the wing.

Struts and Wires Figure 22 Page 2-4


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2.1.6 BEAMS AND FLOOR STRUCTURES

Beams are often used laterally and longitudinally along the fuselage to support the flight deck and passenger cabin floors. Additionally they provide strong point attachments for the crew and passenger seats and as such, constitute primary structure. Modern cabin flooring is usually made up from a number of removable composite honeycomb core panels, examples of which are shown below, whereas the flight deck is often made from metal panels supported on beams.

Floor Structures Figure 23 2.1.7 METHODS OF SKINNING

Skins for light aircraft are usually simple, thin sheets of aluminium alloy, wrapped around and riveted to the internal structure. Larger aircraft, developed since the 1950’s have their skins manufactured from heavier material with the additional use of even thicker sections in certain places where more strength is required. As the aircraft designs became more complex, the excess weight of thicker skins in places where they are not necessarily required, became too big a penalty. To overcome this problem, the skins were rolled individually to produce a variety of differing thickness across each sheet, to cater for variations in stress.


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The latest methods are to machine or mill each skin panel individually from a solid billet, to include all stringers and risers and to provide a varying thickness all over the sheet. In this way, the skin panel is exactly the right thickness at each location, with no excess material and hence no extra weight. This method results in what is termed milled skin or machined skin. Milled wing skins give maximum strength and rigidity with minimum weight. Panels containing areas of different thickness can also be produced from a chemical etching process where areas which have been treated, will be removed to about half their thickness by the chemical etch. The nature of the etching process ensures that no ‘stress raisers’ are introduced into the material. So called ‘waffle plates’ can be produced in this way and are shown in Fig 24.

Skinning Methods Figure 24

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2.1.8 ANTI-CORROSIVE PROTECTION

Materials used in aircraft construction are selected primarily for their strength and tenacity. Unfortunately, many may readily suffer serious damage from corrosion unless effectively protected and the rate of corrosion attack can be extremely rapid in certain environments. One of the main considerations in the design of aircraft structure therefore, are measures for the control and prevention of corrosion. During manufacture and assembly, a range of surface treatments are applied. Materials are heat treated to refine grain structure, sacrificial coatings in the form of plating and cladding are employed, to retard the onset of corrosion. Epoxy primers, special paint finishes, wet-assembly techniques and the use of barrier sealants to prevent the ingress of dirt and moisture between component parts, all help to reduce the risk of corrosion. Additionally, drain holes, drainage paths and attention to good corrosion resistant design techniques for each component part, ensure that aircraft newly off the production line are protected as much as possible, before entering airline service. Aircraft are required to operate in widely varying, often highly corrosive environments throughout the world and despite the high standard of protective treatments applied during manufacture, corrosion will still occur. Corrosive attack may extend over an entire metal surface, may penetrate locally to form deep pits or may follow the grain boundaries within the metal. The weakening effect of corrosive attack may be aggravated by stresses in the metal and result in premature failure of the component. These stresses may be due to externally applied loads or may be internal stresses locked into the metal structure during manufacturing processes, despite the care taken to keep the risk to a minimum. Whatever the cause and type of corrosive attack, unless preventative maintenance is carried out, damage may become so severe, it could present a serious hazard to the airworthiness of the aircraft. Rectification of advanced corrosion damage is time consuming and much of the corrosion during service can be prevented or contained by simple corrosion prevention measures Corrosion seldom occurs on a clean dry aircraft especially if the protective coatings are completely in tact. Since aircraft have to operate outside throughout their lives, they are difficult to keep dry, but keeping the protective coatings free from scratches, dents and scores, ensuring drains which might allow water to accumulate are kept clear and keeping the aircraft clean and free of dirt are all within the scope of a good maintenance engineer. In addition, the engineer should clear up spills from the galleys and toilets and remove deposits from engine exhausts as these are also very corrosive if left on the skin for too long.


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2.1.9 CONSTRUCTION METHODS – WING

The basic requirement for wing construction, particularly with cantilever types is for a spanwise member of great strength, usually in the form of a spar. Conventionally, there are three general designs, monospar, two-spar or multispar. Most modern commercial airliners, have a wing comprising top and bottom skins complete with spanwise stringers, front and rear spars and a set of wing ribs running chordwise across the wing between the spars. This forms a box-like shape which is very robust and the addition of nose ribs and trailing edge fittings produce the characteristic aerofoil shape. Wing structures carry some of the heaviest loads found in aircraft structure. Fittings and joints must be carefully proportioned so they can pick up loads in a gradual and progressive manner and redistribute them to other parts of the structure in a similar manner. Special attention must be paid to minimising stress concentrations, by avoiding too rapid a change in cross section and to provide ample material to handle any concentration in stress or shock loading that cannot be avoided, such as landing loads.

Typical Wing Construction Figure 25

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2.1.10 CONSTRUCTION METHODS – EMPENNAGE

The vertical and horizontal stabilisers, elevators and rudder are constructed in a manner similar to the wings but on a smaller scale. The main structural members are the spars, with the stringers, ribs and stressed skin completing the basic design.

Typical Stabilizer Construction Figure 26


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2.1.11 CONSTRUCTION METHODS – ENGINE ATTACHMENTS

Engine mountings consist of the structure that transmits the thrust provided by either the propeller or turbojet, to the airframe. The mounts can be constructed from welded alloy steel tubing, formed sheet metal, forged alloy fittings or a combination of all three. Some typical examples are shown in Figures 27 to 29. All engine mounts are required to absorb not only the forward thrust during normal flight, but the reduced force of reverse thrust and the vibrations produced by the particular engine/propeller combination..

Fabricated Piston Engine Mounting Figure 27

Tubular Turbopropeller Mounting Figure 28 Page 2-10


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Machined Turbojet Side Mounting Figure 29


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2.1.12 STRUCTURAL ASSEMBLY TECHNIQUES

The integrity of an aircraft joint depends on the way the parts are attached together. The most common method of attachment is by the use of rivets or more sophisticated types of rivets, known as fasteners. However, where high strength is required, nuts and bolts are used whilst other structural assembly is achieved by the use of adhesive bonding techniques. Although aluminium alloy is the most common material for aircraft construction, more and more structural components and in some cases, complete aircraft, are being manufactured from composite materials like glass or carbon fibre. Riveting is generally divided into two types: (1) solid shank rivets and (2) special fasteners. The special fastener category being sub-divided further into special and blind fasteners. 2.1.13 SOLID SHANK RIVETS

The vast majority of aircraft structure is held together with solid rivets. As will be explained later, many of the more modern designs use special fasteners and some bonded construction, but the majority are still solid rivets. Head Shapes In the past there have been a large number of rivet head shapes used in aircraft, but in recent years these have been reduced and standardised to four main types: The Universal Head, sometimes known as AN70 or MS20470, is most popular and may be used to replace any protruding-head rivet. It is streamlined on top but thick enough to provide strength without protruding too much into the airflow. A Round Head rivet, AN430, is used on internal structure where the thicker head is more suitable for automatic riveting equipment. In internal locations where a flat head rivet can be driven more easily than either a round or universal head rivet, the AN442 Flat Head rivet may be used. Where a smooth skin is important, flush rivets such as AN426 or MS20426, with a 100° countersink head are used. Additionally, rivets with a different countersink angle, such as 90° and 120° degrees can be found.

Rivet Head Types Figure 30 Page 2-12


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Types of Alloy used for Solid Shank Rivets The identification marks on rivet heads serve two important functions. Firstly, the marks are used to identify the rivet alloy required for a special installation area and, secondly, the head markings are necessary when trying to identify which kind of rivets are being removed from an aircraft during disassembly or repair. The alloy identifying marks are made on rivet heads at the time they are being stamped out during manufacture. Generally, solid rivets are manufactured in five different materials:

Solid Rivet Identification Figure 31 For non-structural applications, rivets made from pure aluminium, sometimes known as 'A' rivets, may be used. A very popular rivet is the 'AD' rivet, which has copper and magnesium added to the aluminium base metal. This rivet is heat treated during manufacture to make it strong, whilst still being soft enough to be formed easily. When much more strength than the 'AD' rivets is required, there are two stronger rivets available. These are 'D' and 'DD' rivets but they must be heat treated to make them softer before they can be formed. The 'D' types are of 2017 alloy and the 'DD' types are manufactured from 2024 alloy. Both of these rivet types, after heat treatment, must be formed within a specific period of time (one hour for 'D' and ten minutes for 'DD' types) or they may be put into a refrigerator to maintain the softening effect. Once refrigerated they will remain useable for about 10 days.


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When riveting magnesium alloy sheets, there must be no copper in the rivet alloy, or dissimilar metal corrosion will set in. Therefore, a 'B' rivet, manufactured from 5056 alloy is used. This contains a large amount of magnesium with a little manganese and chromium but no copper. Dimensions Aircraft rivet dimensions are categorised by the diameter of the shank, ‘D’, and the length, ‘L’, measured from the end of the shank to the portion of the head that will be flush with the surface of the metal. This means that a countersink rivet is measured from the top of its head, whilst the remainder are measured from under the head.

Rivet Dimensioning Figure 32 Identification The complete identification of a rivet includes its head style, its material, its diameter and its length. The identification code shows the diameter as a number of 1/32ths of an inch and the length as a number of 1/16ths of an inch. For example, An MS20470AD4-4 has a universal head (MS20470), is made from alloy 2117 (AD), is 1/8" diameter (4 x 1/32”) and 1/4" long (4 x 1/16”).

2.1.14 SPECIAL AND BLIND FASTENERS.

When solid shank rivets become impractical to use, then special fasteners are used. These, you will remember, are of two types; special and blind fasteners. The term ‘Special Fasteners’ refers first to their job requirement and second to the tooling needed for the installation. In certain locations, aircraft require strength that cannot be produced by a solid shank rivet, so a special high strength fastener is used. For example, if high shear strength is required, then special High Shear rivets are used. These are usually installed with special tools and will be discussed later in this chapter.

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Blind Fasteners There are several different types of blind fasteners which can be hollow or selfsealing. They include the following types, all of which can be installed from one side of the work. •

Chobert

•

Avdel

•

Tucker/Pop

•

Cherry

Note: It is most important that the correct tools are always used with the types of rivets mentioned above. Chobert Rivets These are available with a snap (round) head or a countersink head and are closed by forcibly pulling a mandrel through the bore of the rivet. This closes the 'tail' and expands the rivet tightly into the hole. To seal Chobert rivets, a separate sealing pin is driven into the hollow bore of the rivet.

Chobert Rivet Figure 33


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Tucker or 'Pop' Rivets Tucker/'Pop' rivets are manufactured with either domed or countersunk heads and are supplied on individual mandrels. The rivets can be either ‘break head’ or ‘break stem’ and when closed, can be sealed or open depending upon their application. Break head rivets are rarely used due to the 'foreign object' risk from the broken off heads lying within the internal aircraft structure. Break stem rivets are be divided into two groups, short and long break mandrels. Long break types leaves the stem in place, greatly increasing the shear strength of the rivet.

Tucker ‘Pop’ Rivet Figure 34 Cherry Rivets These rivets, of American manufacture, are similar to Avdel rivets, except that the stem is positively locked in the rivet bore. During final forming, a locking collar is forced into a groove in the stem, preventing further movement. After the closing operation, the remainder of the stem is milled flush with the skin. There are many different types of Cherry rivets, two of the most popular being the Cherry Lock and the Cherry Max. The Cherry Lock, however, requires a range of closing tools for different sized rivets, whilst the Cherry Max series can all be closed with a single tool. Cherry Lock rivets are manufactured from 2017 or 5056 alloys, Monel metal or Stainless Steel, whereas Cherry Max are made from 5056 alloy, Monel or Inconel 750. They are all available with either universal or countersink heads and due to their positive locking method, can be installed in place of solid shank rivets.

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Cherry Lock Rivet Figure 35

Cherry Max Rivet Figure 36 Issue 1 – Module 11.2 21 Dec 2001

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Avdel Rivets These are similar to Chobert rivets, but each is fitted with its own stem (mandrel). The stem is pulled through the rivet body to close the rivet and at a predetermined load, breaks off proud of the manufactured head. This leaves part of the stem inside the body which seals the rivet. The excess stem is then removed by nipping it off and carefully milling it until flush with the surface of the aircraft skin. The shear strength of an Avdel rivet is greater than a Chobert rivet of equivalent material and size and similar to a solid rivet.

Avdel Rivet Figure 37 Special Fasteners These can include Hi-Shear, Avdelock, Jo-Bolts, and Rivnuts. The first three are all formed by means of a collar which is swaged into the grooves in fastener shank or expanded over the shank to form a blind head. Rivnuts are formed using a similar method to cherry locks, but with a threaded mandrel screwed into the Rivnut. The advantage of Rivnuts, (see Fig 38), is that after closing, a fixed nut is left behind which may be used for the attachment of de-icing boots, floor coverings and other non-structural parts.

Rivnuts After Installation Figure 38

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2.1.15 BOLTS AND NUTS

Bolts A bolt is designed to hold two or more parts together. It may be loaded in shear, in tension, or both. Bolts are designed to be used with nuts and have a portion of the shank that is not threaded, called the grip, whereas Machine screws and Cap screws have the entire length of the shank threaded. The dimensions required to identify a bolt are expressed in terms of the diameter of the shank and the length from the bottom of the head to the end of the bolt. The grip length should be the same as the thickness of the material being held together. This measurement can be found by reference to the applicable charts. Bolt heads are made in a variety of shapes, with hexagonal being the most common.

Bolt Terminology Figure 39 General Purpose Bolts All-purpose structural bolts used for both tension and shear loading is made under 'AN' standards from 3 to 20, the bolt diameter is specified by the AN number in 1/16"; for example: AN3 = 3/16" diameter AN11 = 11/16" diameter The range is from AN3 to AN20 which have hexagon heads, are made from alloy steel and have UNF (fine) threads. The length of the bolt is expressed as a dash number. Bolts increase in length by 1/8" and the dash number(s) will show the length.


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For example: AN3-7 = 7/8" long AN3-15 = 1 5/8" long Other markings will identify whether the bolt has a drilled shank, a drilled head for locking and indicate what material the bolt is made from. Clevis Bolts These bolts (AN21 to 36) are designed for pure shear load applications such as control cables. The slotted, domed head results in this bolt often being mistaken for a machine screw. A clevis bolt has only a short portion of the shank threaded with a small notch between the threads and the plain portion of the shank, which allows the bolt to rotate more freely in its hole. Because the length of this bolt is more critical than normal bolts, its length is given in 1/16" increments.

Clevis Bolt Identification Figure 40

Nuts All nuts used on aircraft must have some sort of locking device to prevent them from loosening and falling off. Many nuts are held in place on a bolt, by passing a split pin through a hole in the bolt shank and through slots, or castellations, in the nut. Others have some form of locking insert that grips the bolt's thread, whilst others rely on the tension of a spring-type lock-washer to hold the nut tight enough against the threads to prevent them from vibrating loose. Sometimes, nuts that are plain with no locking devices are used and prevented from coming undone, once they have been tightened, by the use of locking wire attached to an adjacent nut or to the aircraft structure.

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There are two basic types of nuts, self-locking and non self-locking. As the name implies, a self-locking nut locks onto a bolt with no external help, whilst a non selflocking nut relies on either a split pin, lock-nut, locking washer or locking wire, to stop it from undoing.

Standard Nuts Figure 41 Another type of nut in general use is the Anchor nut. These are permanently mounted on nut plates that enable inspection panels and access doors to be easily removed and installed, without access being required on the reverse side of the work. To make fitment of the panel easier when there is a large number of screws, the nuts are often mounted 'floating' on their mounts, which allows for small differences in the position of the attaching screws. Although rarely used on large commercial airliners, Tinnerman nuts are manufactured from sheet steel and are used mainly on light aircraft, for the fitting of instruments into the flight deck panels, the attachment of inspection panels, etc. Some light aircraft engine cowlings have U-type tinnerman nuts fitted over the inner edge of the cowling frame. When the retaining screws are tightened, spring action holds them tightly and safely in place.


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Examples of self-locking nuts, anchor nuts and U-type tinnerman nuts are shown in figures 42 and 43 below.

Self Locking and Anchor Nuts Figure 42

U-Type Tinnerman Nut Figure 43

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2.1.16 ADHESIVE BONDED STRUCTURES

Adhesive Bonding is the technique of joining materials using special adhesives. In the past a common type of adhesive widely used in metal to metal joints was the ‘Redux’ epoxy resin system. ‘Redux’ is the trade name for a range of adhesives produced by the Ciba-Geigy company and the epoxy bonding procedure in general, refers to a hot-melt, hot-cure adhesive, which is available in partly cured strips or sheets. Note: This type of epoxy resin is also used to provide the reinforcement for fibre composite construction and has already been covered as a separate topic in Module 6. In metal to metal bonding, the sheets of partly cured adhesive, which at this stage resemble strips of chewing gum, are cut to exact size. With the backing paper peeled away, they are carefully placed between each of the components being joined together and the joint securely clamped. The complete assembly, which for example might consist of a wing skin with all its stringers and ribs in place, is then loaded into an autoclave (pressure cooker) to complete the curing process. The adhesive melts and flows evenly into the narrow gaps between the component parts and cures to produce a very strong bond. In the autoclave the temperature limits are strictly controlled, (typically not above 100°-150°C, depending on type of adhesive used), and subjected to a constant clamping force (usually by a vacuum process), resulting in perfect bonded joints which are as strong as, or stronger than, equivalent riveted joints. For composite repairs, figure 45, a portable Autoclave process is employed. There are a number of aircraft, in which the majority of the primary metal structure is joined together entirely with adhesive bonding, with very few rivets being used. The Fokker 50/70/100 and BAe 146/RJ are good examples of aircraft employing this technique extensively. In fact British Aerospace claims that by using adhesive bonding techniques on the BAe 146/RJ airframe, over 10,000 rivets are not required. This means the weight of the rivets, the work that would be expended in closing them and the risk of subsequent in-service cracks (see Figure 44) emanating from rivet holes, are all saved on each airframe. A further important advantage of using adhesively bonded structures, is improved sealing of integral fuel tanks, eliminating the leakage problems that are typical of riveted assemblies.

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Comparison between Machined and Bonded Structure Failure Rates Figure 44

Autoclave Curing Process During Composite Repair Issue 1 – Module 11.2 21 Dec 2001

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MODULE 11.2 AIRFRAME STRUCTURES Figure 45

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2.1.17 METHODS OF SURFACE PROTECTION

As mentioned in an earlier chapter, there are many different types of surface protection added to the basic structural materials and hardware. Anodising A method of protecting aluminium based alloys from corrosion, especially when cladding is impractical, is by a process called Anodising. This is an electrolytic treatment which coats the host metal with a film of oxide. This film is hard, waterproof, air-tight and to aid in identification of some parts, will permanently accept a coloured dye. The film also acts as an insulator, so when bonding leads are to be attached to an anodised part, the surface treatment must be carefully removed before the bonding lead is attached. Finally, anodising a part also provides an excellent base for the addition of an organic finish and bonding adhesives. There are a number of different organic finishes applied to aircraft to protect the surfaces: Synthetic Enamel.- An older finish which cures by the process of oxidation It has a good surface finish, but is poor when it comes to its resistance to chemicals or wear. Acrylic Lacquer.- A popular finish in the mass production market, easy to apply and has a fairly good resistance to chemical attack and weather. Polyurethane.- One of the most durable finishes which has high resistance to wear, fading and chemicals. It also has a 'wet look'. Chromating Chromate coatings are used to protect Magnesium-based alloys, as well as zinc and its alloys. Components are immersed in a bath containing potassium bichromate and results in a yellowish coating on magnesium alloys. The coating can be restored locally with Alocrom 1200 treatment. Cladding There are two metals most commonly alloyed with aluminium, to produce high strength skin and component parts for aircraft manufacture. These are, Copper and Zinc. These alloys suffer extensively from the effects of corrosion, so a cladding technique is used as a form of corrosion protection. ‘Alclad’ as it is termed is a soft, highly corrosion-resistant, pure aluminium skin, rolled onto the face of each base alloy sheet, effectively sandwiching the alloy.


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Surface Cleaning Most aircraft will be cleaned before starting on large inspections, but it is common sense to keep an aircraft clean all of the time. Dirt can cover up cracked or damaged components as well as trap moisture and solvents which can lead to corrosion. Note: Materials mentioned in this chapter are only used as an example, each aircraft type will have a list of suitable and prohibited materials in its maintenance manuals (AMM). Exterior Cleaning Exterior cleaning is an important facet of corrosion control, but there are a number of points which must first be protected from cleaning materials and high pressure water sprays. The pitot tubes and static vents must be properly blanked off to prevent water ingress and the wheels, tyres and brake assemblies need to be covered to keep them free of aggressive cleaning agents. Only cleaning agents and chemicals recommended by the manufacturer are to used. for the job in hand or the risk of serious contamination may result. One of the unseen effects of using non-approved cleaning agents is hydrogen embrittlement. This is caused by hydrogen from the agent being absorbed into the metal, causing minute cracks and will lead to stress corrosion failure. Aircraft should ideally be washed on a proper platform with suitable drains. It is better if the outside air temperature is not too high, so the cleaning agent does not evaporate. Typically, a mix of water and an emulsion-type cleaner, to a ratio of between 3:1 and 5:1 is applied, allowed to soak for a few minutes and then rinsed off with a high pressure stream of water. Engine cowlings and wheel well areas usually have grease, oil or brake dust deposits that require special treatment. These require stronger mixtures ratios and scrubbing with a soft bristle brush to loosen the dirt before rinsing off with a high pressure water jet. It must be borne in mind however, that oil and grease could be accidentally removed from places where they are meant to be, for example in wheel bearings etc. These will often require re-lubrication after washing has been completed. Exhaust residue from both piston and jet engines is very corrosive and must be removed on a regular basis. These deposits usually require a special proprietary solvent to mix with the water. Sometimes a simple emulsified mix of kerosene and water may be approved. Dry-cleaning solvent or naptha is sometimes used for oil and grease removal. Some naptha compounds are harmless to rubber or acrylic items, whilst others will attack these same materials, so only approved specifications are to be used.

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2.1.18 EXTERIOR FINISH MAINTENANCE

All materials used on the exterior of an aircraft must be approved by the manufacturer of that aircraft to ensure no abrasives or solvents are applied where they can do damage. Non-Metallic Cleaning Non-metallic components sometimes require different cleaning techniques from metal parts. For example, the slightest amount of dust on plastic or acrylic panels will scratch and severely reduce the optical quality if rubbed with a dry cloth. This can also build up a static charge and attract more dust so the correct procedure in this situation is to wash down, rinse with water without rubbing with a cloth. Oil and hydraulic fluid also attack rubber components such as tyres, so any spillages must be cleaned up immediately. Neoprene rubber leading-edge de-icer boots and composite structures are other examples of parts that need special cleaning procedures, all of which will be detailed in the AMM. Engine Cleaning Apart from external cleaning carried out on the engine cowlings, with the associated protection of electrical components; gas turbine engines are regularly washed internally to remove the deposits of dust, sand and salt, that tend to accumulate on internal parts of the engine. This coating if not removed, can have a serious effect on the engine's performance. Indeed, the output of the engine could fall below the manufacturers minimum figures, resulting in an unscheduled and expensive engine change Alignment and Symmetry Aircraft can have abnormal occurrences during their life, when for example, a very heavy landing could occur, some accidental external damage or the need to replace a major component, etc. All of these instances will require special checks to be carried out to guarantee that the aircraft is perfectly symmetrical and aligned before its next flight. The checks consist of measuring very accurately from a number of datum points on the airframe, such as from wing tips, the nose, the horizontal stabiliser and the top of the vertical stabiliser. The checks vary, depending on the aircraft manufacturers requirements, but all ensure that measurements taken on the lefthand side of the aircraft are within a minimum tolerance of the measurements from the right-hand side. These checks are usually taken with the aircraft on jacks and in the rigging position, ie: a nominally level ‘in flight’ attitude.


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On light aircraft, these measurements are usually taken using a surveyors tape measure. (It is a check of comparison, not of outright measurement). As the aircraft get larger, optical theodolite style methods are used. These can be a microscopic level with the use of sighting rods or even a laser ranging alignment device. Deeper checks that are carried out after any of the above mentioned situations, as well as on a routine basis, include checks on the wing, tail and control surfaces to ensure that they are set at the correct angles. These checks are usually known as 'rigging checks' and are carried out using purpose built levelling boards and an accurate measuring device known as a Clinometer.

Rigging Checks - Older Aircraft Figure 46

Symmetry Checks – Modern Aircraft Figure 47 Page 2-30


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EASA PART 66 - MODULE 11.02 - Airframe Structures

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