Turbine aeroplane aerodynamics structure and systems -:airframe structure -general concept

STRUCTRUL CLASSIFICATION-
Aircraft structure is divided into three categories for
the purposes of assessing damage and the application of
repair protocol that are suitable for the structure under
consideration. Manufacturer manuals designate which
category a structure falls under and the technician
is required to repair and maintain that structure in
accordance with rules specified for the category under
which it falls. The three categories for structure are:

primary, secondary and tertiary.
Primary structure-Primary structure is any portion of the aircraft structure
that, if it fails, on the ground or in flight, would likely
cause any of the following:
• A loss of control of the aircraft.
• Catastrophic structural collapse.
• Injury to occupants.
• Power unit failure.
• Unintentional operation.
• Inability to operate a service.
Some examples of primary structure are wings spars,
engine mounts, fuselage frames, and main floor
structural members. Within the primary structure are
elements called principle structural elements (PSE's).
These elements are those which carry flight, ground
and pressurization loads.
Primary structure may also be represented as a structurally
significant item or SSL These elements are specified in a
supplemental structural inspection document. Due to
their structural importance, they may require special
inspection and have specific repair limitations.
Secondery structure-Secondary structure is all non primary structure portions
of the aircraft which have integral structural importance
and strength exceeding design requirements. These
structures weakening without risk of failure such as
those described for primary structure. Prominent
examples of secondary structure are wing ribs, fuselage
stringers and specified sections of the aircraft skin.
Tertiary structure-Tertiary structure is the remaining structure. Tertiary
structures are lightly stressed structures that are fitted to
the aircraft for various reasons. Fairings, fillets, various
support brackets, etc. are examples of tertiary structure.

DAMAGE OF TOLERANT-

Fail safe-Fail safe means the structure has been evaluated, usually
by the manufacturer, to assure that catastrophic failure is
not probable after fatigue failure or obvious partial failure
of a single, principal structural element. It is designed so
that the aircraft may continue to operate safely until the
defect is detected in a scheduled maintenance check.
Manufacturer testing and fatigue analysis is used when
developing fail safe structural elements. The elements
are considered damage tolerant.
Safe life-
Safe life structural elements are those which have a
very low risk of unacceptable degradation or failure for

a stated amount of time. The fatigue capability of thestructure is learned through testing. The stresses applied
while in service are designed to be significantly lower.
Also, the calculated time in service before failure is
greatly reduced so that failure of the structure before its
safe life is highly unlikely. The affects of corrosion, wear
and fatigue are considered when operating under the
safe life design principle.
Damage tolerance-Designing aircraft with fail safe principles can be
somewhat unreliable. Accidents have occurred that
prove this. Engineering improvements to a fail safe
structure typically come with the extra penalty of
adding weight. Thus, the damage tolerant concept of
engineering is favored.
By distributing loads over a larger area and designing
multiple load paths for carrying loads, a structure can be
damage tolerant. The structure retains its integrity and the
damage does not worsen in service between inspections
when the damage can be detected and repaired. Thus,
Damage tolerance means that the structure has been
evaluated to ensure that should serious fatigue, corrosion,
or accidental damage occur within the operational life
of the aeroplane, the remaining structure can withstand
reasonable loads without failure or excessive structural
deformation until the damage is detected.

STATION IDENTIFICATION AND ZONAL LOCATION SYSTEM-


Station numbering system-Even on small, light aircraft, a method of precisely
locating each structural component is required. Various
numbering systems are used to facilitate the location of
specific wing frames, fuselage bulkheads, or any other
structural members on an aircraft.
Most manufacturers use some system of station
marking. For example, the nose of the aircraft may
be designated "zero station," and all other stations are
located at measured distances in inches behind the zero
station. Thus, when a blueprint reads "fuselage frame
station 137," that particular frame station can be located
137 inches behind the nose of the aircraft.

To locate structures to the right or left of the center 

line of an aircraft, a similar method is employed. Many 

manufacturers consider the center line of the aircraft to bea zero station from which measurements can be taken to 

the right or left to locate an airframe member. This is often 

used on the horizontal stabilizer and wings. The applicable 

manufacturer's numbering system and abbreviated 

designations or symbols should always be reviewed before 

attempting to locate a structural member. They are not 

always the same. The following list includes location 

designations typical of those used by many manufacturers. 

• Fuselage stations (Fus. Sta. or FS) are numbered in 

inches from a reference or zero point known as the 

reference datum. The reference datum 

is an imaginary vertical plane at or near the nose of 

the aircraft from which all fore and aft distances are 

measured. The distance to a given point is measured 

in inches parallel to a center line extending through 

the aircraft from the nose through the center of the 

tail cone. Some manufacturers may call the fuselage 

station a body station, abbreviated BS.

 Buttock line or butt line (BL) is a vertical reference
plane down the center of the aircraft from which
measurements left or right can be made.
Water line (WL) is the measurement of height
in inches perpendicular from a horizontal plane
usually located at the ground, cabin floor, or some
other easily referenced location.
Aileron station (AS) is measured outboard from,
and parallel to, the inboard edge of the aileron,
perpendicular to the rear beam of the wing.
Flap station (KS) is measured perpendicular to the
rear beam of the wing and parallel to, and outboard
from, the inboard edge of the flap.
Nacelle station (NC or Nae. Sta.) is measured either
forward of or behind the front spar of the wing and
perpendicular to a designated water line .

In addition to the location stations listed above, other 

measurements are used, especially on large aircraft. 

Thus, there may be horizontal stabilizer stations (HSS), 

vertical stabilizer stations (VSS) or powerplant stations 

(PPS)In every case, the manufacturer's 

terminology and station location system should be 

consulted before locating a point on a particular aircraft.

Zonal identification-Another method is used to facilitate the location of 

aircraft components on air transport aircraft. This 

involves dividing the aircraft into zones. Large areas 

or major zones are further divided into sequentially 

numbered zones and sub zones. The digits of the zone 

number are reserved and indexed to indicate the location 

and type of system of which the component is a part.

Access and inspection penal-Knowing where a particular structure or component 

is located on an aircraft needs to be combined with 

gaining access to that area to perform the required 

inspections or maintenance. To facilitate this, access 

and inspection panels are located on most surfaces of 

the aircraft. Small panels that are hinged or removable 

allow inspection and servicing. Large panels and doors 

allow components to be removed and installed, as well 

as human entry for maintenance purposes. 

The underside of a wing, for example, sometimes 

contains dozens of small panels through which control 

cable components can be monitored and fittings 

greased. Various drains and jack points may also be 

on the underside of the wing. The upper surface of 

the wings typically have fewer access panels because a 

smooth surface promotes better laminar airflow, which 

causes lift.On large aircraft, walkways are sometimes designated 

on the wing upper surface to permit safe navigation 

by mechanics and inspectors to critical structures and 

components located along the wing's leading and trailing 

edges. Wheel wells and special component bays are 

places where numerous components and accessories are 

grouped together for easy maintenance access. 

Panels and doors on aircraft are numbered for positive 

identification. On large aircraft, panels are usually 

numbered sequentially containing zone and sub zone 

information in the panel number. Designation for a leftor right side location on the aircraft is often indicated in 

the panel number. This could be with an "L'' or "R," or 

panels on one side of the aircraft could be odd numbered 

and the other side even numbered. 

The manufacturer's maintenance manual explains 

the panel numbering system and often has numerous 

diagrams and tables showing the location of various 

components and under which panel they may be 

found. Each manufacturer is entitled to develop its 

own panel numbering system.

STRUCTRUL STRESS-

Aircraft structural members are designed to carry a 

load or to resist stress. In designing an aircraft, every 

square inch of wing and fuselage, every rib, spar, and 

even each metal fitting must be considered in relation 

to the physical characteristics of the material of whichit is made. Every part of the aircraft must be planned to 

carry the load to be imposed upon it. 

The determination of such loads is called stress analysis. 

Although planning the design is not the function ofthe aircraft technician, it is, nevertheless, important 

that the technician understand and appreciate the 

stresses involved in order to avoid changes in the 

original design through improper repairs.

The term "stress" is often used interchangeably with the 

word "strain." While related, they are not the same thing. 

External loads or forces cause stress. Stress is a material's 

internal resistance, or counterforce, that opposes 

deformation. The degree of deformation of a material is 

strain. When a material is subjected to a load or force, 

that material is deformed, regardless of how strong the 

material is or how light the load is. 

There are five major stresses to which all aircraft are 

subjected: (Figure 2-6) 

• Tension • Shear 

• Compression • Bending 

• Torsion 

Tension is the stress that resists a force that tends to pull 

something apart. (Figure 2-6A) The engine pulls the 

aircraft forward, but air resistance tries to hold it back. 

The result is tension, which stretches the aircraft. The 

tensile strength of a material is measured in pounds per 

square inch (psi) and is calculated by dividing the load 

(in pounds) required to pull the material apart by its 

cross sectional area (in square inches). 

Compression is the stress that resists a crushing force. 

(Figure 2-6B) The compressive strength of a material 

is also measured in psi. Compression is the stress that 

tends to shorten or squeeze aircraft parts. 

Torsion is the stress that produces twisting. (Figure 

2-6C) While moving the aircraft forward, the engine 

also tends to twist it to one side, but other aircraft 

components hold it on course. Thus, torsion is created. 

The torsion strength of a material is its resistance to 

twisting or torque. 

Shear is the stress that resists the force tending to cause 

one layer of a material to slide over an adjacent layer. 

(Figure 2-6D) Two riveted plates in tension subject the 

rivets to a shearing force. Usually, the shearing strength 

of a material is either equal to or less than its tensile or 

compressive strength. Aircraft parts, especially screws, 

bolts, and rivets, are often subject to a shearing force. 

Bending stress is a combination of compression and 

tension. The rod in Figure 2-6E has been shortened 

(compressed) on the inside of the bend and stretched on 

the outside of the bend. A single member of the structure 

may be subjected to a combination of stresses. In most 

cases, the structural members are designed to carry end 

loads rather than side loads. They are designed to be 

subjected to tension or compression rather than bending. 

Strength or resistance to the external loads imposed 

during operation may be the principal requirement in 

certain structures. However, there are numerous other 

characteristics in addition to designing to control the 

five major stresses that engineers must consider. For 

example, cowling, fairings, and similar parts may not 

be subject to significant loads requiring a high degree of 

strength. However, these parts must have streamlined 

shapes to meet aerodynamic requirements, such as 

reducing drag or directing airflow.

Hoop stress-Hoop stress is the stress on the airframe structural 

components caused by pressurization. All transport 

category aircraft are pressurized. A circumferential 

load is experience in hoop stress. The structural 

fuselage framework resists this load with the aid of 

the stressed skin. 

Note that axial loads in the fuselage are also partial 

resisted by the stressed skin construction as well 

as the longitudinal structural members such as 

longhorns and stringers.

 Metal fatigue-Metal fatigue is experienced by a component or 

structural member when a load is repeatedly applied 

and released or applied and reversed. This cycling 

weakens the material over time even though the load 

applied may be well below that which causes damage 

in a single application. 

All materials have an elastic limit. If applied loads do 

not exceed this limit, the material should be unaffected 

by the load and returns to its original state when 

the load is removed. However, an aircraft in flight 

constantly experiences varying loads. Over time, 

these small load changes cause fatigue in the form of 

minute cracks in the metal structure. Each tiny, seemlyinconsequential crack exposes new material to the 

elements. This may weaken the material through corrosion. 

Additionally, when a multitude of tiny fissures 

combine, larger significant cracks may develop and 

weaken the metal to the point of failure. 

Aircraft structure is tested at the manufacturer to 

determine a limit not to be exceeded for an aircraft in 

service. Often, fatigue testing is accomplished on full 

scale fatigue rigs which subject the elements to cycles 

of loading and unloading or reversal well beyond that 

which will be experienced in service by the aircraft. 

A fatigue index is applied and the aircraft is monitor 

throughout its service life. If its fatigue life limit is 

consumed, an aircraft may be reevaluated to perceive its 

actual condition. If the loading cycles and environmental 

exposure of the structure was not as harsh as calculated, 

it is possible to extend the service life of the aircraft. An 

increased in inspection frequency and/or strengthening 

modification(s) may be required to do so. 

Fatigue characteristics vary with the type of metal 

and how it is worked. The thickness of the material 

and type and number of fastener hole can alter the 

fatigue life. Aging aircraft are monitored and treated 

by technicians to protect against corrosion which 

accelerates metal fatigue

DRAINAGE AND VENTILATION PROVISIONS 

DRAINAGE 

The collection of water and other fluids in the many 

cavities found on an aircraft can lead to corrosion and 

could present a fire hazard. Drainage and ventilation 

are used to address this issue. There are two types of 

drains, internal and external. 

External drains have openings to the exterior of the 

aircraft. They are found on the wings, empennage and 

fuselage as well as engine nacelles. An external drain 

dumps the fluid overboard,. In unpressurized aircraft 

the drains may remain open at all times. Drain valves 

are used in pressurized sections of aircraft so that they 

may remain sealed during pressurization. Typically 

located along the aircraft keel, some external drains use 

the pressurizing air to hold the valve closed. A rubber 

flapper type valve, a plunger type valve or a normally 

open spring loaded valve are closed by pressurization 

air. When depressurized, such as when the aircraft is 

on the ground, the drain valves open. 

Leveling compound is sometimes used to build up a low 

area near a drain valve to ensure that no fluid is trapped and 

it flows out the drain orifice. This is typically a waterproof 

rubber like sealant without structural characteristics. 

Some fluids accrue during flight and need to be drained. 

Galley and lavatory drain masts must be heated to 

prevent ice formation and blockage caused by cold 

temperatures at high altitude. A drain mast is nothing 

more than an airfoil shaped projection designed to guide 

the fluid overboard away from the skin of the aircraft. 

Most have electric resistance heating elements or use hot 

air from the pneumatic system to combat icing. 

Internal drain paths are required to direct fluid to 

the external drain sites. Tubes, channels, dams and 

internal drain holes are all common. The design of 

structural members often includes considerations that 

prevent fluids from being trapped. 

VENTILATION 

Any cavity in the aircraft structure that may experience 

the presence of a flammable vapor or water must 

be ventilated to permit the vapor to evaporate. If 

necessary, vent pipes are used provide an escape route 

for the vapor. Some highly susceptible areas, such as 

an engine nacelle, may even contain ram air inlets and 

exit points to enable a full flow of fresh air through the 

cavity. The technician should ensure that all openings 

designed for ventilation are unobstructed.

SYSTEM INSTALLATION PROVISIONS 

In addition to designing functioning support systems 

for operation of the aircraft, design engineers must 

also make the system components fit into the aircraft.Depending on the system and components, provisions 

for access and servicing must also be addressed. Items 

that receive regular maintenance such as filters, fluidlevel checks, bearing lubrication, etc. must be located so 

that technicians can easily access them. Line replaceable 

units (LRU's) must be able to be quickly uninstalled and 

installed. Aircraft maintenance is a significant expense 

for the operator. Anything that can be done to locate 

system components for easy access for maintenance saves 

time and lowers the cost of operating the aircraft. 

Modern airliner designers often group the components 

of a various systems in a single bay for easy access. 

Air conditioning, for example, may have its several 

key components mounted next to each other in an air 

conditioning bay. The hydraulic reservoir, pumps and 

filters may all be located in a different bay or in the 

wheel well area. Avionics and electronics are frequently 

mounted in an avionics bay. Not only are the "black 

boxes" easily accessible but environmental conditions 

can be better controlled than if the units were spread 

throughout the aircraft.

CONSTRUCTION METHODS 

FUSELAGE 

The fuselage is the main structure or body of the fixed 

wing aircraft. It provides space for cargo, controls, 

accessories, passengers, and other equipment. In single 

engine aircraft, the fuselage houses the powerplant. In 

multi engine aircraft, the engines may be either in the 

fuselage, attached to the fuselage, or suspended from 

the wing structure. There are two general types of 

fuselage construction: truss and monocoque. 

TRUSS TYPE 

A truss is a rigid framework made up of members, 

such as beams, struts, and bars to resist deformation 

by applied loads. The truss framed fuselage is generally 

covered with fabric. The truss type fuselage frame is 

usually constructed of steel tubing welded together in 

such a manner that all members of the truss can carry 

both tension and compression loads. (Figure 2-9) 

In some aircraft, principally the light, single engine models, 

truss fuselage frames may be constructea.of aluminum 

alloy and may be riveted or bolted into one piece, with cross 

bracing achieved by using solid rods or tubes. 

STRESSED SKIN FUSELAGE 

MONOCOQUE TYPE 

The monocoque (single shell) fuselage relies largely 

on the strength of the skin or covering to carry the 

primary loads. The design is called stressed skin and 

may be divided into two classes: 

1. Monocoque 

2. Semimonocoque

Different portions of the same fuselage may belong to 

either of the two classes, but most modern aircraft are 

considered to be of semimonocoque type construction. 

The true monocoque construction uses formers, 

frame assemblies, and bulkheads to give shape to the 

fuselage. (Figure 2-10) The heaviest of these structural 

members, bulkheads, are partition type walls that 

typically span the entire fuselage diameter often with 

an opening for access through the partition. They are 

located at intervals to carry concentrated loads and 

at points where fittings are used to attach other units 

such as wings, powerplants, and stabilizers. Since no 

other bracing members are present, the skin must 

carry the primary stresses and keep the fuselage rigid. 

Thus, the biggest problem involved in monocoque 

construction is maintaining enough strength while 

keeping the weight within allowable limits.

SEMIMONOCOQUE TYPE 

To overcome the strength/weight problem of monocoque 

construction, a modification called semimonocoque 

construction was developed. It also consists of frame 

assemblies, bulkheads, and formers as used in the 

monocoque design but, additionally, the skin is 

reinforced by longitudinal members called longerons. 

Longerons usually extend across several frame members 

and help the skin support primary bending loads. They 

are typically made of aluminum alloy either of a single 

piece or a built up construction. 

Stringers are also used in the semimonocoque fuselage. 

These longitudinal members are typically more numerous 

and lighter in weight than the longerons. They come in 

a variety of shapes and are usually made from single 

piece aluminum alloy extrusions or formed aluminum. 

Stringers have some rigidity but are chiefly used for 

giving shape and for attachment of the skin. Stringers 

and longerons together prevent tension and compression 

from bending the fuselage. (Figure 2-11) 

Other bracing between the longerons and stringers can 

also be used. Often referred to as web members, these 

additional support pieces may be installed vertically 

or diagonally. It must be noted that manufacturers use 

different nomenclature to describe structural members. 

For example, there is often little difference between 

some rings, frames, and formers. 

One manufacturer may call the same type of brace a ring 

or a frame. Manufacturer instructions and specifications 

for a specific aircraft are the best guides. 

The semimonocoque fuselage is constructed primarily 

of alloys of aluminum and magnesium, although steel 

and titanium are sometimes found in areas of high 

temperatures. Individually, no one of the aforementioned 

components is strong enough to carry the loads imposed 

during flight and landing. But, when combined, those 

components form a strong, rigid framework. This is 

accomplished with gussets, rivets, nuts and bolts, screws, 

and even friction stir welding. A gusset is a type of 

connection bracket that adds strength. (Figure 2-12) 

To summarize, in semimonocoque fuselages, the strong, 

heavy longerons hold the bulkheads and formers, 

and these, in turn, hold the stringers, braces, web 

members, etc. All are designed to be attached together 

Module 11 A - Turbine Aeroplane Structures and Systems 

and to the skin to achieve the full strength benefits of 

semimonocoque design. It is important to recognize that 

the metal skin or covering carries part of the load. The 

fuselage skin thickness can vary with the load carried 

and the stresses sustained at a particular location. 

The advantages of the semimonocoque fuselage are 

many. The bulkheads, frames, stringers, and longerons 

facilitate the design and construction of a streamlined 

fuselage that is both rigid and strong. Spreading loads 

among these structures and the stressed skin means no 

single piece is failure critical. 

This means that a semimonocoque fuselage, because ofits 

stressed skin construction, may withstand considerable 

damage and still be strong enough to hold together. 

BEAMS FLOOR STRUCTURES 

In addition to the structural members already mentioned, 

additional beams, floor structural members and 

various other reinforcement members are also used to

construct an aircraft. A beam may be installed laterally 

or longitudinally. Beams typically support the floor of 

the flight deck and the passenger compartment. They 

are situated to provide secure attachment of the floor 

panels and also the seats tracks into which the passenger 

seats are secured. The floor itself is typically made up 

of numerous honeycomb constructed panels that are 

screwed to the floor support structure. Flight deck floor 

panels may be constructed from sheet metal. 

STRUTS AND TIES 

Struts and ties are also used in aircraft structure. A strut 

is a bar or rod shaped reinforcement designed to resist 

compression loads. A tie is a rod or beam designed to 

take a tensile load. Both are used as needed to reinforce 

the aircraft structure throughout the fuselage to carry 

the loads experienced. 

METHODS OF SKINNING 

Attached to the outside of the aircraft structure is the 

aircraft skin, be it stressed or not. Simple, light aircraft 

generally have skin made from sheet aluminum which 

is formed to fit, wrapped and riveted to the structural 

members. Larger, more complex and heavier aircraft 

used heavier material to form the aircraft skin. This 

is to transfer and carry the greater loads experience 

during high performance flight. Some simple sheet 

metal skin may be found. However, various skin 

thickness are used to meet the design loads which vary 

by location around the aircraft. 

Since in many areas the skin thickness varies, 

machining the skin, including integrally formed 

stringers and risers, from a solid billet of material 

has become a standard practice. By milling the skin 

out of a single piece of material, the skin thickness 

may be varied precisely to meet design requirements. 

Maximum strength is achieved with minimum weight 

and no excess. (Figure2-13) 

Another process used in skinning a large aircraft is 

chemical etching. Etching of thicker skin material to 

form thinner material with supporting raised patterns 

of material are produced without any stress. Skin with a 

"waffle plate" pattern is produced this way. 

DOUBLERS 

A simpler way to reinforce an area of skin on the aircraft 

which receives greater loads than can easily be carried 

by a single sheet of material is to create a doubler for 

that area. A doubler is simply a second, reinforcing 

layer of skin material used to strengthening the load 

caring capacity of the skin. It has the advantage of being 

inexpensive and is able to be shaped for a specific area 

identified as needing reinforcement. Doublers are also 

used in sheet metal repair work. 

WING, EMPENNAGE AND ENGINE ATTACHMENT 

The wings, empennage and engines must be attached 

to the fuselage. The type of attachment varies with the 

aircraft design. Typically, special pins or bolts are used. 

Wings and empennage structure is often constructed 

with load carrying main members called spars. Attach 

lugs securely fitted to these spars mate with lugs that 

are fitted to strengthened sections of the fuselage and 

mounting pins or bolts are passed through both lugs 

and secured. Figure 2-14 shows the internal fuselage 

structure of what is considered the center section of the 

horizontal stabilizer on a Boeing 737. Its lugs are mated 

with the lugs on the horizontal stabilizer front spar (each 

side, top and bottom) and attached with bolts. 

Various wing and empennage attach methods exist 

including a single piece structure that passes through 

the fuselage making it basically non removable. 

Configurations where numerous smaller bolts and 

permanent fasteners are used to attach wings and 

empennage airfoils are also common. Strength and 

spreading the load throughout the fuselage attach 

structure is achieved with any of these methods.

Engine attachments vary widely on aircraft depending on 

where the engines are located and the size and design of 

the aircraft and engine. A typical arrangement found on 

transport aircraft is to extend support structure forward 

and down from the wing spars. The structure is called 

a pylon. Figure 2-15 is a rough cutaway drawing of a 

turbofan engine pylon. It is built to be very strong to 

support the engine. Attached to the pylon structure are 

engine mounts to which the engine is bolted or visa versa. 

The engine mounts on most turbofan engines, for 

example, perform the basic functions of supporting 

Genter Section 

Clevis Lugs 

the engine and transmitting the loads imposed by 

the engine to the pylon and aircraft structure. Most 

turbine engine mounts are made of stainless steel, and 

are typically located as illustrated in Figure 2-16. Some 

engine mounting systems use two mounts to support 

the forward end of the engine and a single mount at the 

rear end. The mounting arrangement depends on the 

position of the engine and the pylon structure. Figure 

2-17 illustrates the pylon and the side engine mount 

configuration for a rear engine aircraft.

STRUCTURE ASSEMBLY TECHNIQUES 

The structures of the majority of today's aircraft are 

primarily aluminum. However, advances in the used 

of composite materials such as glass and carbon fiber 

is steadily increasing. A myriad of fasteners are used 

to join together aluminum structural elements. 

Most common are rivets, bolts and nuts and a 

wide variety of special application fasteners. A full 

discussion of aircraft materials and hardware is found 

in Module 06 - Materials and Hardware of this series.As early "rag and tube" aircraft construction was 

replaced by aluminum construction, assembly using 

rivets dominated assembly techniques. Light and heavy 

aircraft today still use the rivet as a primary fastener on 

structural and non structural elements. But as aircraft 

design evolved, larger and heavier aircraft were produced. 

Structural members increased in size and complexity. 

Rivets were not always suitable to assemble the new 

structure. Stronger fasteners, some designed specifically 

for use in aircraft assembly, were introduced.

Bolts are used in many locations on aluminum aircraft 

when fastening large structural members and when 

attaching both fixed and moveable components. 

Special bolts such as Hi-loks, Jo-bolts and lock-bolts 

are common as are clevis bolts where hi shear loads 

are present. Close tolerance bolts are used where a 

tight drive fit is required. 

Special fasteners called blind fasteners are used in areas 

where access to only one side of an assembly is possible. 

Module 11 A - Turbine Aeroplane Structures and Systems 

A variety of blind fasteners are used including several 

classified as rivets. Structural sections and components 

of the aeroplane that are made from composite material 

may be assembled and attached in a variety of ways. 

Sleeves and fitting incorporated during construction of a 

panel, for example, facilitate the use of bolts. 

Other fasteners may be specified depending on the 

design and location of the structure. The panel itself is 

constructed using methods described in Module 06

Materials and Hardware of this series. It is of the utmost 

importance to follow manufacturer's instructions when 

assembling composite structures. Many components 

are bonded or require special fasteners with specific 

torque considerations. Note also that some metal 

structural members are bonded. Epoxy sheet bonding 

using autoclave curing is sometimes used to bond metal 

components resulting in extremely high strength joinery. 

Large aircraft maintenance manuals contain specific 

instruction for the bonding of all materials and sections 

of the aircraft. ATA section 51 gives a descriptive 

overview of the aircraft structure and general rules 

followed in construction of airframe components 

. 

and sections. The manufacturer's structural repair 

manual (SRM) details numerous repair procedures 

and techniques for all aircraft structure repair. A large 

aircraft fuselage is manufactured in sections that are 

then mated and fastened together. 

The structural sections of a Boeing 737 are shown in 

Figure 2-18. Sections 41, 43, and 48 comprised the 

pressurized portion of the fuselage. Section 48 is not 

pressurized but does supply the support structure for 

the vertical and horizontal stabilizer. It also contains 

a bay for installation of the auxiliary power unit. A 

rear pressure bulkhead separates body section 46 

from body section 48.

ANTI-CORROSION PROTECTION 

Preventing the corrosion of aircraft structures is 

a consideration when materials are selected for its 

construction. Suitable anti-corrosion measures are then 

taken before and during construction. These range from 

heat treatment of the material to a variety of surface 

treatments to design and assembly techniques all 

designed to prevent corrosion. 

Heat treatment of a metal can refine its grain structure so 

that it has the properties required for a specific function 

while reducing its susceptibility to corrosion. Surface 

treatments can protect metals from contaminants and 

moisture which cause corrosion. Plating and cladding of 

materials are common methods of corrosion protection. 

When these are designed to degrade rather than having 

the material they cover degrade, they are known as 

sacrificial coatings. Common surface treatments such as 

paints and primers are used as well as metal specific thin 

surface treatments such as anodizing and chromating. 

Numerous similar surface treatments have been 

developed for specific metals in specific applications all 

of which endeavor to keep the causes of corrosion at bay. 

The design of an aircraft part or assembly can be very 

instrumental in preventing corrosion. Something as 

simple as a well designed drain path or a drain hole 

placed in a strategic location can prevent corrosion of 

material in a vulnerable area. Wet assembly techniques 

and the use of sealants also provide a barrier to 

corrosion causing agents. 

Module 11A-Turbine Aeroplane Structures and Systems 

Manufacturers use all techniques at their disposal to 

produce a corrosion resistant aircraft. However, varied 

aircraft operating environments and maintenance 

practices combine with service loads sustained during 

operation make corrosion inevitable. Processing of 

susceptibility data obtained from field operations is used 

with a wide variety of inspection and testing techniques 

to find and correct corrosion before it reaches a critical 

phase. Anti-corrosion treatments and repairs are 

detailed throughout the manufacturer's maintenance 

manuals, especially in ATA chapter 51, Structures. 

Rarely does corrosion occur on a clean, dry aircraft 

properly treated by the manufacturer during 

construction. While in service, it is impossible to avoid 

exposure of the aircraft to the elements. The agents of 

corrosion, namely dirt and moisture, are encountered. A 

program of keeping aircraft clean and diligence to keep 

the condition of surface treatments in good condition 

are main combatants for operators when preventing 

corrosion. Technicians must assist by wiping up spills 

and removing deposits that contribute to the corrosive 

environment. Scratches, dents, and scoring should be 

avoided while performing maintenance. Drain holes 

must not be plugged so they can function as designed.

METHODS OF SURFACE 

PROTECTION 

The manufacturer's maintenance manual details the 

surface protection compounds that must be applied by the 

technician for all of the various areas of the aircraft. Again, 

ATA Chapter 51 in the maintenance manual and the SRM 

should be consulted. Different areas on the aircraft may 

be prone to different contaminants and the recommended 

treatments are designed accordingly. Do not assume that a 

product is suitable for treatment of an area of the aircraft 

structure without consulting the manufacturer's data. 

ANODIZING 

Manufacturers use a variety of methods of surface 

protection on structural metals and hardware. One 

of the most common for aluminum based alloys is 

anodizing. Anodizing is an electrolytic treatment that 

coats the metal with a hard, waterproof and airtight, 

oxide film. Anodizing usually contains a dye. Various 

colors are used. This permits easy identification that a 

part has be anodized. The oxide film acts as an isolator. 

When attaching a bonding lead, the film must carefully 

be removed to ensure electrical conductivity. 

Anodizing provides an excellent base for many finishes 

as well as for bonding adhesives. Acrylic lacquers, and 

polyurethane paints adhere well to anodized parts and 

provide good resistance to chemical attack and wear. 

CHROMATING 

An alternative to anodizing used for surface protection 

on magnesium and zinc alloy parts is chromate. 

When chromated, parts are generally immersed in a 

potassium bichromate solution. The chromate coating 

protects the surface from corrosive elements and has a 

yellowish appearance on magnesium alloys. Products 

are available to obtain a chromate coating on a part in 

the field. Alocrom 1200 is one such product. 

CLADDING 

Cladding a material with another, non corrosive material 

is a popular means of material surface protection. This 

is done as the raw material is formed into the product 

material. Sheet aluminum, for example, may be clad to 

protect the corrosive copper or zinc aluminum alloy from 

which many aluminums products are made. Alclad is a 

process of cladding aluminum in which a pure aluminum 

skin is rolled onto the face of an alloy aluminum sheet. 

Pure aluminum forms a stable aluminum oxide surface 

2.18 AIRCRAFT - TECH N ICAL 

Boo k Comp any 

when exposed to air that protects the pure aluminum 

itself and the material that has been clad. 

PAINTING 

Many aircraft structural elements and parts are painted 

to protect them from corrosion. The paint acts as a 

barrier so that the agents of corrosion cannot reach the 

material being protected. To be effective, paint must be 

applied to a clean dry surface. It must be compatible 

with the material composition so that a good bond 

is formed and it adheres when it is applied. Material 

surface treatments such as paint primer and alodine are 

used before painting because they bond strongly to the 

base material as well as to the paint. 

SURFACE CLEANING 

Nearly all surface treatments to aircraft metals 

begin with a thorough cleaning of the material. 

This may include stripping of old paint before new 

paint or primer is applied. Strippers are specifically 

recommended by the manufacturer that do not react 

with the base metal of the structure. Therefore, only 

use strippers that are recommended. 

A cleaned surface is often treated with alodine before 

a primer or painted coating is applied. Clad aluminum 

parts use a different formula of alodine than non clad 

alloys. Be sure to use the correct formula. 

Personal safety procedures should be followed when 

cleaning, stripping and applying any surface treatment. 

Solvents, strippers, cleaners, etchants and conversion 

coatings can all be hazardous to the health of the 

technician. Avoid breathing vapors from products 

of this type and avoid prolonged skin contact. Use 

protective gloves, goggles, respirators and other 

protective gear. Know the location of the nearest 

eyewash fountain when working with these substances. 

Flush eyes with water if one splashes into the eyes and 

get medical attention immediately. Generally, specified 

paint strippers are used on metal surfaces only. Protect 

all surrounding areas from accidental contact with the 

stripper. Polyethelene film and suitable adhesive tape is 

used for masking. 

In particular, Teflon lines, self lubricated bearings, 

electrical terminal plugs, nylon coated wires and 

nylon bushings should be protected from contact

with chemicals used in strippers. Plastics, laminates, 

composites, fiberglass and bonded structures usually 

have paint removed by abrasive cleaning. Do not use 

stripper on composite structures. Use only the methods 

described by the manufacturer. 

EXTERIOR AIRCRAFT CLEANING 

Aircraft are cleaned before major inspections. Typically 

a high pressure water or steam is sprayed in conjunction 

with cleaning agents to clean the exterior of the aircraft. 

While a clean aircraft aids in corrosion prevention, the 

cleaning process may put water and agent where it is not 

desirable and, thus, it may even cause corrosion. Areas 

into which the cleaning spray should not enter must be 

Module 11 A - Turbine Aeroplane Structures and Systems 

covered or sealed from its entrance. Pitot tubes and static 

ports are such areas as well as tires and brake assemblies. 

The manufacturer's maintenance manual gives detailed 

instructions on cleaning procedures. Areas to be 

protected and the proper cleaning agents to use must 

be noted. A cleaning agent that is suitable for one 

area of the aircraft may not be for another. Follow all 

manufacturer instructions when cleaning. 

Aircraft are generally washed outside in an area with 

adequate and environmentally responsible drainage. 

Washing with cleaning agents should not be performed in 

high temperatures where the agent may dry before being

rinsed off. In certain locations, this may relegate washing 

to inside of a hangar. Use the ratio of agent to water that 

is recommended. Use of the wrong agent may cause the 

agent to attack materials. Hydrogen embrittlement occurs 

when certain agents soak into an aircraft metal. Minute 

cracks form and stress corrosion develops. 

ALIGNMENT AND SYMMETRY 

The position or angle of the main structural components 

is related to a longitudinal datum line parallel to the 

aircraft center line and a lateral datum line parallel to a 

line joining the wing tips. Before checking the position 

or angle of the main components, the aircraft must be 

jacked and leveled. 

Small aircraft usually have fixed pegs or blocks attached 

to the fuselage parallel to or coincident with the datum 

lines. A spirit level and a straight edge are rested across 

the pegs or blocks to check the level of the aircraft. 

This method of checking aircraft level also applies to 

many of the larger types of aircraft. However, the grid 

method is sometimes used on large aircraft. The grid 

Special Dihedral Board with 

Spirit Level Incorporated 

Engine and wheel well areas may require a special 

washing technique or cleaning agents due to dirt, 

oil, grease and exhaust debris buildup. Again, follow 

manufacturer's instructions. Be aware that some 

cleaning procedures are followed by greasing various 

locations that may have had grease washed out during 

the cleaning process. 

plate is a permanent fixture installed on the aircraft 

floor or supporting structure. (Figure 2-19) 

When the aircraft is to be leveled, a plumb bob is 

suspended from a predetermined position in the ceiling 

of the aircraft over the grid plate. The adjustments to the 

jacks necessary to level the aircraft are indicated on the 

grid scale. The aircraft is level when the plumb bob is 

suspended over the center point of the grid. 

Certain precautions must be observed in all instances 

when jacking an aircraft. Normally, rigging and 

alignment checks should be performed in an 

enclosed hangar.

If this cannot be accomplished, the aircraft should be 

positioned with the nose into the wind. 

The weight and loading of the aircraft should be exactly 

as described in the manufacturer's manual. In all cases, 

the aircraft should not be jacked until it is determined 

that the maximum jacking weight (if applicable) 

specified by the manufacturer is not exceeded. 

With a few exceptions, the dihedral and incidence angles 

of conventional modern aircraft cannot be adjusted. 

Some manufacturers permit adjusting the wing angle 

of incidence to correct for a wing heavy condition. The 

dihedral and incidence angles should be checked after 

hard landings or after experiencing abnormal flight 

loads to ensure that the components are not distorted 

and that the angles are within the specified limits. 

There are several methods for checking structural 

alignment and rigging angles. Special rigging boards 

that incorporate, or on which can be placed, a special 

instrument (spirit level or inclinometer) for determining 

the angle are used on some aircraft. On a number of 

aircraft, the alignment is checked using a transit and 

plumb bobs or a theodolite and sighting rods. The 

particular equipment to use is usually specified in the 

manufacturer's maintenance manual. 

When checking alignment, a suitable sequence should 

be developed and followed to be certain that the checks 

are made at all the positions specified. The alignment 

checks specified usually include: 

• Wing dihedral angle 

• Wing incidence angle 

• Verticality of the fin 

• Engine alignment 

• A symmetry check 

• Horizontal stabilizer incidence 

• Horizontal stabilizer dihedral 

CHECKING DIHEDRAL 

The dihedral angle should be checked in the specified 

positions using the special boards provided by the aircraft 

manufacturer. If no such boards are available, a straight 

edge and a inclinometer can be used. The methods for 

checking dihedral are shown in Figure 2-20. 

It is important that the dihedral be checked at the 

positions specified by the manufacturer. Certain 

portions of the wings or horizontal stabilizer may 

sometimes be horizontal or, on rare occasions, 

anhedral angles may be present. 

CHECKING INCIDENCE 

Incidence is usually checked in at least two specified 

positions on the surface of the wing to ensure that the 

wing is free from twist. A variety ofincidence boards are 

used to check the incidence angle. Some have stops at 

the forward edge, which must be placed in contact with 

the leading edge of the wing. Others are equipped with 

location pegs which fit into some specified part of the 

structure. The purpose in either case is to ensure that the 

board is fitted in exactly the position intended. In most 

instances, the boards are kept clear of the wing contour 

by short extensions attached to the board. A typical 

incidence board is shown in Figure 2-21. 

When used, the board is placed at the specified 

locations on the surface being checked. If the incidence angle is correct, a inclinometer on top of the board 

reads zero, or within a specified tolerance of zero. 

Modifications to the areas where incidence boards are 

located can affect the reading. For example, if leading 

edge de-icer boots have been installed, the position of a 

board having a leading edge stop is affected. 

CHECKING FIN VERTICALITY 

After the rigging of the horizontal stabilizer has been 

checked, the verticality of the vertical stabilizer relative 

to the lateral datum can be checked. The measurements 

are taken from a given point on either side of the top of 

the fin to a given point on the left and right horizontal 

stabilizers. (Figure 2-22) The measurements should be 

similar within prescribed limits. 

When it is necessary to check the alignment of the 

rudder hinges, remove the rudder and pass a plumb bob 

line through the rudder hinge attachment holes. The 

line should pass centrally through all the holes. It should 

be noted that some aircraft have the leading edge of 

the vertical fin offset to the longitudinal center line to 

counteract engine torque. 

CHECKING ENGINE ALIGNMENT 

Engines are usually mounted with the thrust line parallel 

to the horizontal longitudinal plane of symmetry. 

However, this is not always true when the engines are 

mounted on the wings. Checking to ensure that the 

position of the engines, including any degree of offset 

is correct, depends largely on the type of mounting. 

Generally, the check entails a measurement from 

the center line of the mounting to the longitudinal 

center line of the fuselage at the point specified in the 

applicable manual. (Figure 2-23) 

2.22 AIRCRAFT - TCCHNICAL 

Doak C omp;111y 

SYMMETRY CHECK 

The principle of a typical symmetry check is illustrated 

in Figure 2-23. The precise figures, tolerances, and 

checkpoints for a particular aircraft are found in 

the applicable service or maintenance manual. On 

small aircraft, the measurements between points are 

usually taken using a steel tape. When measuring long 

distances, it is suggested that a spring scale be used 

with the tape to obtain equal tension. A five pound pull 

is usually sufficient. 

On large aircraft, the positions at which the dimensions 

are to be taken are usually chalked on the floor. This is 

done by suspending a plumb bob from the checkpoints 

and marking the floor immediately under the point 

of each plumb bob. The measurements are then taken 

between the centers of each marking.