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Issue 1 Module 11.2 21 Dec 2001 Page 1-1
JAR 66 CATEGORY B1
MODULE 11.2
AIRFRAME
STRUCTURES
engineering
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CONTENTS
1 AIRFRAME STRUCTURES GENERAL CONCEPTS ................ 1-3
1.1 AIRWORTHINESS REQUIREMENTS FOR STRUCTURAL STRENGTH .....
1-3
1.2 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
1.3 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
1.4 ZONAL AND STATION IDENTIFICATION SYSTEM
................................ 1-9 1.4.1 Zonal System
................................................................
1-9 1.4.2 Station Identification System
......................................... 1-10
1.5 LOADS FOUND WITHIN 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
1.6 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
1.7 LIGHTNING STRIKE PROVISION
...................................................... 1-21
2 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
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AIRFRAME
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Issue 1 Module 11.2 21 Dec 2001 Page 1-3
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MODULE 11.2
AIRFRAME
STRUCTURES
engineering
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1 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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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
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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
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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 stations
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
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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
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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.
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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
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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 -60C 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
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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' (LRUs) 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
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2 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 chickens 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
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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
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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 1950s 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
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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
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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 self-sealing. 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
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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 self-locking 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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INTENTIONALLY BLANK
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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
100-150C, 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
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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 left-hand 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
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INTENTIONALLY BLANK
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