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Contents
1 MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURES AND SYSTEMS)
..........................................................................................
1-2
1.1 AEROPLANE AERODYNAMICS AND FLIGHT CONTROLS
..................... 1-2 1.1.1 Fixed Aerofoils
.............................................................. 1-2
1.1.2 Moveable Control Surfaces
........................................... 1-6 1.1.3 High Lift
Devices ...........................................................
1-13 1.1.4 Drag Inducing Devices
.................................................. 1-14 1.1.5
Airflow Control Devices Wing Fences ......................... 1-17
1.1.6 Boundary Layer Control
................................................ 1-18 1.1.7 Trim
Tabs
......................................................................
1-21 1.1.8 Mass Balance
...............................................................
1-24 1.1.9 Control Surface Bias
..................................................... 1-26 1.1.10
Aerodynamic Balance Horn Balance .......................... 1-26
1.1.11 Aerodynamic Balance Inset
Hinge.............................. 1-27
1.2 HIGH SPEED FLIGHT
.....................................................................
1-28 1.2.1 Speed of Sound
............................................................ 1-28
1.2.2 Subsonic Flight
............................................................. 1-29
1.2.3 Transonic Flight
............................................................ 1-30
1.2.4 Supersonic Flight
.......................................................... 1-32
1.2.5 Aerodynamic Heating
.................................................... 1-39 1.2.6
Area Rule
......................................................................
1-40 1.2.7 Factors Affecting Airflow in Engine Intakes of High
Speed Aircraft 1-41 1.2.8 Effects of Sweepback on Critical Mach
Number ............ 1-43
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1 MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURES AND SYSTEMS)
The principles of Aircraft Theory of Flight are covered in JAR
66 Module 8.
1.1 AEROPLANE AERODYNAMICS AND FLIGHT CONTROLS
An aircraft is equipped with fixed and moveable surfaces, or
aerofoils, which provide stability and control. Each item is
designed for a specific function during the operation of the
aircraft.
Typical Aircraft Flight Controls Figure 1
1.1.1 FIXED AEROFOILS
The fixed aerofoils are the wings or mainplanes, the horizontal
stabiliser or tailplane and vertical stabiliser or fin. The
function of the wings is to provide enough lift to support the
complete aircraft. The tail section of a conventional aircraft,
including the stabilisers, elevators and rudder, is occasionally
known as the empennage.
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1.1.1.1 Horizontal Stabiliser
The horizontal stabiliser is used to provide longitudinal pitch
stability and is usually attached to the aft portion of the
fuselage. It may be mounted either on top of the vertical
stabiliser, at some mid-point, or below it.
Conventional horizontal stabilisers are placed aft of the wing
and normally set at a slightly smaller or negative angle of
incidence with respect to the wing chord line.
This configuration gives a small downward force on the tail with
a value dependent on the size of the stabiliser and its distance
from the Centre of Gravity (CG).
Horizontal Stabiliser
Figure 2
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1.1.1.2 T-Tail Arrangement
The T-Tail Arrangement places the complete stabiliser/tailplane
and elevator assembly on top of the vertical stabiliser. This
ensures that pitch control is not affected by turbulent air from
the wing. It also makes the vertical stabiliser and rudder control
more effective, due to the so-called end plate effect.
However a T-Tail (and rear engine) configuration, would be
dangerous if the aircraft entered what is termed a deep stall. At a
very high angle of attack (i.e.: stalling angle), airflow could
make pitch control non-effective (and may cause the engines to
flame out). To prevent this, T-Tailed aircraft will have a stick
push system, in order to automatically recover them safely from
excessive angles of attack.
The T-Tail has another disadvantage in that the empennage
structure will be heavier than normal, due to the strengthening
required to combat greater bending loads. However since the pitch
moment arm is increased, the stabiliser and elevators can be made
smaller and therefore lighter than conventional designs.
Often, the complete stabiliser can be moved to provide
longitudinal trim, negating the use of trim tabs (later in Module
11.09).
TTail Arrangement
Figure 3
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1.1.1.3 Vertical Stabiliser
The vertical stabiliser for an aircraft is the aerofoils forward
of the rudder and is used to provide directional stability.
A problem encountered on single-engined propeller driven
aircraft is that the propeller causes the airflow to rotate as it
travels rearward. This strikes one side of the vertical stabiliser
more than the other, resulting in a yawing moment. These aircraft
may have the leading edge of the stabiliser offset slightly,
thereby causing the airflow to pass around it in such a manner to
counter the yaw.
Off-Set Vertical Stabiliser
Figure 4
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1.1.2 MOVEABLE CONTROL SURFACES
Moveable control surfaces are normally divided into Primary and
Secondary controls.
The primary control surfaces include the elevators, rudder,
ailerons and roll spoilers. The secondary control surfaces consist
of trim controls (tabs), high lift devices (flaps and slats), speed
brakes and lift dumpers (additional spoilers).
Note: Traditionally, spoilers have not been included as primary
controls, but those which operate in conjunction with the ailerons
during roll, are considered to be primary in the JAR 66 syllabus,
so this is how these notes will define them.
The primary control surfaces are used to make the aircraft
follow the correct flight path and to execute certain
manoeuvres.
The secondary controls are used to change the lift and drag
characteristics of the aircraft or to provide assistance to the
primary controls.
Moveable Control Surfaces
Figure 5
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1.1.2.1 Roll Control - Ailerons
These primary controls provide lateral (roll) control of the
aircraft, that is, movement about the longitudinal axis. They are
normally attached to hinges at the trailing edge of the wing, near
the wing tip. They move in opposite directions, so that the
up-going aileron reduces lift on that side, causing the wing to go
down, whilst the down-going surface increases the lift on the
opposite side, raising the wing.
Large aircraft often use two sets of aileron surfaces on each
wing, one in the conventional position near the wing tip and the
other set at mid-span or outboard of the flaps. The inboard set is
referred to as high speed ailerons. The outboard surfaces, or
sometimes both sets, work at low speeds to give maximum control
during take off and landing, for example when large movements may
be required.
At high cruising speed the outer ailerons are isolated and only
the inboard set operate. If the outer ailerons were permitted to
operate at high speed, the stress produced at the wing tips may
twist the wing and produce aileron reversal. This is particularly
likely with modern highly flexible thin wings, where the
possibility of structural damage may result if the outboard
surfaces were too powerful.
The ailerons are operated by a control wheel, a control column
or a side-stick. Movement of any of these inputs away from neutral
towards one side, will result in the aircraft rolling to that side.
Returning the control to neutral at this stage will leave the
aircraft in a banked condition and a similar but opposite movement
will be required to bring the aircraft level once more.
Aileron Controls Figure 6
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The ailerons are usually operated in conjunction with the rudder
and/or elevator during a turn and are rarely used on their own. A
co-ordinated turn is one that occurs without slip or skid. Too
little bank will cause the aircraft to skid outwards, too much bank
will cause the aircraft to slip downwards.
1.1.2.2 Roll Control - Spoilers
The use of spoilers as a primary control, will be to operate
asymmetrically in conjunction with aileron movement and are
normally referred to as Roll Spoilers.
Roll spoilers are mounted on the top of the wing just inboard of
the outboard set of ailerons.
Roll Spoiler Controls Figure 7
Movement of the aileron control wheel on the flight deck will
deploy each spoiler progressively upwards with the up-going
aileron, whilst on the side of the down-going aileron, the spoiler
will remain flush with the upper wing camber.
.This is achieved by the control system being routed via a
spoiler/aileron mixer unit. The up-going spoiler will effectively
spoil the lift on the down-going wing and augment the similar
effect of the up-going aileron.
Alternatively, on some aircraft the spoilers will replace the
ailerons completely to provide the sole means of roll control.
Note: Other spoiler functions are covered later under Secondary
Controls.
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1.1.2.3 Pitch Control - Elevators
The elevators are the control surfaces which govern the movement
of the aircraft in pitch about its lateral axis. They are normally
attached to the hinges on the rear spar of the horizontal
stabiliser.
When the control column of the aircraft is pushed forward, the
elevators move down.. The resultant force of the airflow generated
lift', acting upwards, raises the tail and lowers the nose of the
aircraft. The reverse action takes place when the control is pulled
back.
1.1.2.4 Pitch Control Stabilators
A special type of pitch control surface that combines the
functions of the elevator and the horizontal stabiliser is the
stabilator, often referred to as a slab or all-flying tailplane .
The stabilator is a complete all-moving horizontal stabiliser which
can change its angle of attack when the control column is moved and
thereby alter the total amount of lift generated by the tail.
Elevator Controls Figure 8
Stabilator Controls Figure 9
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1.1.2.5 Pitch Control Variable Incidence Stabilisers
Incorporating a conventional elevator control system, the
variable incidence horizontal stabiliser is often used for pitch
trim. Normally a powerful electric motor is used to vary its angle
of attack when trim switches on the flight deck are operated.
Variable Incidence stabiliser
Figure 10
1.1.2.6 Canards
Some earliest powered aircraft, such as the Wright Flyer, had
horizontal surfaces located ahead of the wings. This configuration,
with the forward surface usually referred to as a canard or
foreplane, has been used on occasions, up to the present day.
Conventional aircraft have the tailplane located at the rear of
the fuselage which provides a small, stabilising down force. This
means that the wing has to produce slightly more lift to balance
this down force. As we have seen, in order for a wing to produce
lift it must also generate drag.
With the tailplane located at the front of the aircraft, the
stabilising force is directed upwards. This contributes to the
total lift of the aircraft, thereby reducing drag from the lift
producing wing.
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A fundamental feature of a canard design is that the angle of
attack of the foreplane, (in front of the CG of the aircraft) is
set at a greater angle than the main wing. This feature will ensure
that the foreplane reaches the stalling angle first, resulting in a
predictable dropping of the nose and a certain recovery.
Additionally, stall sensing systems (later), can be triggered
just before the foreplane reaches its critical angle of attack,
leaving the main wing safely below the stalling angle and still
producing adequate lift.
Canard Design Beech Starship Figure 11
1.1.2.7 Yaw Control - Rudder
The rudder is a vertical control surface that is hinged at the
rear of the fin and is designed to apply yawing moments. The rudder
rotates the aircraft about its vertical axis and is controlled by
rudder pedals that are operated by the pilots feet. Pushing on one
pedal, the right for example, causes the rudder to move to the
right also. This causes the rudder to generate a 'lifting' force
sideways to the left which turns the nose of the aircraft to the
right.
Because of the power of some rudder systems, particularly
assisted systems, they may have their range reduced at high speed
by means of a speed-sensitive range limiting system.(later).
The rudder is normally a single structural unit but on large
transport aircraft it may comprise two or more operational
segments, moved by different operating systems to provide a level
of redundancy.
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Rudder controls
Figure 12
1.1.2.8 Combined-Function Controls Elevons and Ruddervators
An example of combined-function controls is found on delta-wing
aircraft, where control surfaces for pitch and roll must be fitted
on the trailing edge of the wing.
Controls with a dual-function (elevators and ailerons) called
elevons, provide both pitch and roll, by moving symmetrically in
pitch or asymmetrically in roll via a mixer unit, when the control
column or control wheel are operated on the flight deck..
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Another example are ruddervators normally used on aircraft
fitted with a 'V' or Butterfly tail. These surfaces serve the
purposes of both rudder and elevator.
Ruddervator Controls Figure 13
1.1.3 HIGH LIFT DEVICES
Aerodynamic lift is determined by the shape and size of the main
lifting surfaces of the aircraft. In order to produce the
outstanding performance achieved by a large modern, swept wing,
passenger jet such as the Boeing 777, the wing is designed to give
optimum lift to support the aircraft whilst in cruise (typically
Mach 0.87).
This has meant, that to be able to control and land the aircraft
weighing around 200-tonne on runways of reasonable length, the
landing speed needs to be slower than the clean stalling speed of
the aircraft. In order to achieve this, more lift is required and
this is obtained from so-called high lift devices.
These are divided generally into leading edge devices, namely
slots, slats and Krueger flaps and trailing edge devices including
plain, slotted and fowler flaps. They will increase lift and as a
result, reduce the stalling speed. Consequently the landing speed,
(about 1.3 times the stalling speed), will also be reduced, since
drag is also increased with large angles of trailing edge flap
deployment.
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Flaps and Slats Figure 14
Additionally, some aircraft incorporate ailerons, both of which
are designed to move downwards together whenever the trailing edge
flaps are extended to the landing position. These will act as
additional plain flaps and provide extra drag (and lift), but will
still provide roll control if required.
These surfaces are referred to as Droop Ailerons or
Flaperons.
Droop Aileron Figure 15
1.1.4 DRAG INDUCING DEVICES
There are several situations where the aircraft must slow down
fairly quickly. With slower, high drag, light aircraft, simply
closing the throttle allows the high drag of the airframe and the
idling propeller to slow the aircraft down, to gliding speed prior
to landing approach, for example.
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As previously stated, a modern airliner is an extremely smooth,
low drag design which, if only the throttles are retarded, will
continue in level flight for many miles before slowing down.
Furthermore, if the nose were lowered more than a degree or so, the
aircraft will begin to accelerate again.
In order to overcome the problems of low drag on large aircraft
with high momentum, the designers have introduced a variety of drag
inducing devices. These include spoilers, lift dumpers, speed
brakes and in unusual circumstances, lowering the landing gear and
operating in-flight thrust reversers.
1.1.4.1 Spoilers and Lift Dumpers.
Spoilers and Lift Dumpers are usually hinged panels located
about mid-chord position on the upper surface of the wing.
Hydraulically operated, they produce a large amount of turbulence
and drag when deployed, resulting in a reduction of lift.
Lift Dump Spoilers Figure 16
Spoilers, have a variety of uses, all of which involve spoiling
the lift of the wing. Some of the following facilities can be
combined, so that one set of panels can have more than one job.
Firstly, they can be the primary roll control of the aircraft as
described previously.
Secondly, the spoilers can be used in a symmetrical,
part-deployed position, allowing the aircraft to slow down quickly
in the cruise, or descend at a much steeper rate without
accelerating. On some aircraft, the deployment angle of the spoiler
panels can be varied by changing the position of the control lever
in the flight compartment.
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Lift dumpers are, as their name describes, are spoiler panels
incorporated solely to dump lift. They are normally deployed after
landing, destroying the lift of the wing and producing high drag,
to assist in stopping the aircraft efficiently and thereby allowing
the wheel brakes to be operated more effectively.
1.1.4.2 Speed Brakes
Whilst it is true that the in-flight use of spoilers may be
referred to as selecting the 'speed brakes', the term more
accurately describes devices which are solely for the production of
drag without any change of trim. The rear fuselage mounted
'clamshell-type doors which open up on the BAe 146 and Fokker
70/100 aircraft are true speed brakes (or air brakes) and have the
following major advantage over the use of spoilers for producing
drag.
When the wing mounted spoilers are deployed, vibration or rumble
is often felt in the passenger cabin, which some people may find
disturbing. The aft mounted speed brakes not only produce high drag
at any airspeed, but their selection is virtually vibration free.
Also, lift will be completely unaffected, thus permitting their
deployment on approach and making a go-around much safer. (This
will be covered later in powerplants).
Speed Brake Installation
Figure 17
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1.1.5 AIRFLOW CONTROL DEVICES WING FENCES
These devices are usually fitted to aircraft with swept wings.
Total airflow over a swept wing, splits into two components, one
moving across the wing chord parallel to the airflow and the other
flowing spanwise towards the wing tip.
The fences are fitted about mid-span, on the leading edge of the
wing and extending rearwards. They are designed to control the
spanwise flow of the boundary layer air over the top of the wing.
Also they will straighten the airflow over the ailerons, improving
their effectiveness and straighten the air nearer the wing tip,
resulting in less 'spillage' of air from beneath the wing to the
top, thereby producing less drag. (See Winglets later).
Wing Fences Figure 18
1.1.5.1 Airflow Control Devices Saw Tooth Leading Edges
This form of airflow control is more common on military aircraft
than modern commercial airliners. The saw tooth or notch is simply
a small increase in wing chord on the outer portion of the wing.
The step where the change occurs, tends to form an invisible 'wall'
of high velocity air, which flows over the wing and straightens the
spanwise flow. It functions in much the same way as the wing fence
but removes the extra drag and weight penalty.
Leading Edge Notch Figure 19
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1.1.5.2 Airflow Control - Winglets
These can be seen on a variety of the later generation airliners
and business jets. The outboard part of the wing are upswept to an
extreme dihedral angle. These winglets work best at higher speeds
and, by clever aerodynamic design, will give better airflow control
and reduce the drag produced by the wing. It does this by using the
up-flow from below the wing to produce a forward thrust from the
winglet, rather like a yacht sail. The winglets add weight to the
aircraft as well as increasing parasitic drag, but the large
reduction in induced drag at the wingtip, results in a significant
fuel saving.
Winglets Figure 20
1.1.6 BOUNDARY LAYER CONTROL
The boundary layer is that layer of air adjacent to the aerofoil
surface (the boundary between metal and air). If measured, the air
velocity in the layer will vary from zero directly on the surface,
to the relevant velocity of the free stream at the outer extremity
of the boundary layer.
Normally, at the leading edge of the wing the boundary layer
will be laminar, (in smooth thin sheets close to the surface), but
as the air moves over the wing towards the trailing edge, the
boundary layer becomes thicker and turbulent. The region where the
flow changes from laminar to turbulent is called the transition
point. .As airspeed increases, the transition point tends to move
forward, so the designer tries to prevent this thus maintaining
laminar flow, over the top of the wing for as far back as possible.
Methods of boundary layer control are as follows:
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1.1.6.1 Boundary Layer Control - Vortex Generators
One way of stimulating the boundary layer and stopping the
airflow becoming increasingly sluggish towards the trailing edge is
the use of vortex generators.
Vortex generators are small plates or wedges projecting up from
the surface of an aerofoil about 25mm.(about 3 times the typical
boundary layer thickness), into the free stream air. Their purpose
is to shed small but lively vortices from their tip, which act as
scavengers to direct and mix the high energy free stream air into
the sluggish boundary layer air and invigorate it. This action
pushes the transition point backwards towards the trailing edge
.
In this way,the small amount of drag created by the vortices is
far more than compensated by the considerable boundary layer drag
which they save. They also weaken the shock wave at high speed and
reduce shock drag also. (later).
Vortex Generators Figure 21
1.1.6.2 Boundary Layer Control - Stall Wedges
We have seen previously that washout on a wing permits the root
of the wing to stall first, allowing the pilot to retain roll
control during the stall. Even with a degree of washout, the
aircraft will drop a wing on occasions due to adverse boundary
layer air causing the outer part of the wing to stall first. This
can be overcome with the use of stall wedges, or stall strips, as
they are sometimes known.
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Stall Wedges are small, wedge-shaped strips mounted on the
leading edge of the wings at about one third span. The are designed
to disrupt the boundary layer airflow, at large angles of attack
approaching the stall, thus ensuring the airflow breaks
away,(stalls), at the root end of the wing first.
Additionally they produce a similar effect to a wing fence at
smaller angles of attack resulting in a smoother airflow over the
ailerons, thus retaining optimum roll control.
Stall Wedges Figure 22
1.1.6.3 Boundary Layer Control - Leading edge Devices
Other devices to prevent laminar separation at the low speed end
of the range and thus control boundary layer air are leading edge
droop flaps and Kreuger flaps. They can be a droop snoot or
permanent droop type, or can be adjusted during flight.
Krueger (left) and Drooped (right) Leading Edge Flaps Figure
23
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1.1.7 TRIM TABS
During a flight an aircraft will develop a tendency to deviate
from a straight and level hands-off attitude. This may be due to
changes in fuel state, speed, load position or flap/landing gear
selection and could be countered by applying a continuous
correcting force to the primary controls. This would be fatiguing
for the crew and difficult to maintain for long periods, so trim
tabs are used for this purpose instead.
Trim tabs move the primary control surface aerodynamically in
the opposite direction to the movement of the tab. To correct an
aircraft nose down out of trim condition, the elevator tab is moved
down, resulting in the elevator moving up, the tail of the aircraft
moving down, so that the nose comes up, correcting the fault.
1.1.7.1 Fixed Trim Tabs
A fixed trim tab may be a simple section of sheet metal attached
to the trailing edge of a control surface. It is adjusted on the
ground by simply bending it up or down, to a position resulting in
zero control forces during cruise. Alternatively, the tab is
connected to the primary control by a ground-adjustable connecting
rod. Finding the correct position for both types is by trial and
error.
Fixed Trim Tab Figure 24
1.1.7.2 Controllable Trim Tabs
A controllable trim tab is adjusted from the flight deck, with
its position being transmitted back to a flight deck indicator
showing trim units, left and right of neutral.
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Flight deck controls are trim-wheel, lever, switch, etc., with
the actuation of the tab by mechanical, electrical or hydraulic
means. Trim facilities are normally provided on all three axes.
Controllable Trim Tab
Figure 25
Note: Aircraft with hydraulic fully powered controls do not have
trim tabs. Since fully powered controls are termed irreversible,
trim tabs if fitted, would be aerodynamically ineffective. With
these systems, trimming is achieved by moving the primary control
surface to a new neutral datum.(later).
1.1.7.3 Servo Tabs
Sometimes referred to as the flight tabs, servo tabs are
positioned on the trailing edge of the primary control surface and
connected directly to the flight deck control inputs. They act as a
form of power booster, since pilot effort is only required to
deflect the relatively small area of the servo tab into the air
stream.
Movement of the flight deck control input moves the tab up or
down and the aerodynamic force created on the tab, moves the
primary control, until the aerodynamic load on the control surface
balances that on the tab. Moving the tab down will cause the
primary control to move up and vice-versa.
Servo Tab Figure 26
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1.1.7.4 Balance Tabs
Balance tabs assist the pilot in moving the primary control
surface. The flight deck controls are connected to the primary
control surface whereas the balance tab, hinged to the trailing
edge of the primary surface, is connected to the fixed aerofoil.
For example, the elevator balance tab, will be connected by an
adjustable rod to the horizontal stabiliser and is so arranged,
that it tends to maintain the tab at the same relative angle to the
stabiliser when the pilot moves the elevator.
Aerodynamically, therefore, the tab is moving in the opposite
direction to the control surface and assists its movement.
Adjusting the length of the connecting rod will alter the
displacement of the effective range of the tab about the mid-point
datum.
Some types of balance tab have more than one point of attachment
and it is possible with these so called geared balance tabs, to
alter the range of tab deflection.
The function of a balance tab can also be combined with that of
a trim tab, by adjusting the length of the balance tab connecting
rod from the flight deck. This is usually achieved by installing a
form of linear actuator in the rod and is termed a trim/balance tab
(Geared balance and trim/balance tabs will be covered later in the
notes).
Balance Tab Figure 27
1.1.7.5 Anti-Balance Tabs
Anti-balance tabs operate in a similar way aerodynamically as
balance tabs but with a reverse effect. The difference is in the
way it is connected to the fixed aerofoil. It is routed so that the
tab moves, relative to and in the same direction as, the primary
control surface. The effect is to add a loading to the pilot
effort, making it slightly heavier and thus providing feel, to
prevent the possibility of over-stressing the airframe
structure.
Anti-Balance Tab Figure 28
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1.1.7.6 Spring Tabs
At high speed, control surfaces operated directly from the
flight deck, become increasingly difficult to deflect from neutral,
due to the force of the aerodynamic loads caused by the airstream
around them.
The spring tab is progressive in its operation and provides
increasing aerodynamic assistance in moving the control surface,
with an increase in aircraft forward speed. The flight deck
controls are connected to the spring tab in a similar manner to the
servo tab previously described, except the linkage is routed via a
torque rod assembly (or spring box) attached to the primary control
surface.
When the aircraft is stationary or flying at low airspeed the
airloads are non-existent or very small. If the flight deck
controls are deflected from neutral, the rigidity of the torque
tube (or spring force) causes the primary control to be deflected
together with the spring tab. The tab will remain in the same
relative position with the primary control and consequently
provides no additional aerodynamic assistance.
As the aircraft flies faster, the increased force produced by
the airflow, opposes the movement of the primary control surface
from its neutral position. Deflection of the
flight deck controls in this case causes the torque tube to
twist (or the spring to compress), resulting in a deflection of the
spring tab.
The tab deflection provides an added aerodynamic load which
assists the flight deck effort. The faster the aircraft flies, the
greater the airflow force and therefore the greater the spring tab
deflection, resulting in a progressively increasing assistance in
moving the primary control.
Spring Tab Figure 29
1.1.8 MASS BALANCE
All aircraft structures are distorted when loads are applied. If
the structure is elastic, as all good structures are, it will tend
to spring back when the load is removed, or its point of
application is changed.
Since a control surface is hinged near its leading edge, the
centre of gravity (C of G) will be behind the hinge and as a
consequence, there will be more weight aft of the hinge line than
in front of it .
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In the case of an aileron for example, should the air load
distort the wing upwards, it is likely that the aileron will lag
behind and distort downwards. This effectively produces an extra
upward aerodynamic force which pushes the wing up even further.
Due to its elasticity, the wing will spring back and the aileron
will lag again but this time upwards, aerodynamically forcing the
wing down further than it would normally go due to elastic recoil
alone. Now the cycle is repeated and a high speed oscillation will
result. This unwanted phenomenon is referred to as flutter.
Flutter can be prevented if the C of G of the control surface is
moved in line with, or slightly in front of, the hinge line. The
normal way of achieving this is to add a number of high density
weights, either within the leading edge of the surface itself or
externally, ahead of the hinge line. The addition of these weights,
normally made from lead or depleted uranium, is closely controlled
and calculated to ensure that the exact balance is obtained.
This procedure of adding weights is referred to as mass
balancing of the controls.
External Mass Weights Figure 30
Integral Mass Weights
Figure 31
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1.1.9 CONTROL SURFACE BIAS
When a control surface is set so it is not in the true neutral
position it is referred to as having a bias. There are many reasons
for not having the controls in a true central position, including
compensating for design features. As an example, a single propeller
aircraft may have a tendency to roll in the opposite direction to
the engines torque, to counteract this moment the ailerons could be
offset with one slightly up and the other down. Once the aircraft
is flying level with the bias set the trim gauge in the cabin would
then be set to read zero.
1.1.10 AERODYNAMIC BALANCE HORN BALANCE
In order to overcome the high stick forces on larger aircraft at
higher speeds, the surfaces themselves are used to lighten the
forces.
This is referred to as Aerodynamic Balancing and the three
principal ways of achieving it are: horn balance, inset hinge and
pressure balancing.
This method, a small part of the primary control surface ahead
of the hinge will project into the airflow when the control is
deflected from neutral. The airflow on this side assists the
movement of the control in the desired direction and will attempt
to move the control further away from the neutral position.
Air loads on the control side, aft of the hinge, try to push the
surface back towards neutral. (This is the force that would
normally make the controls heavy).
If the proportion of balance area forward of the hinge and
control area aft of the hinge is correct, the pilot will feel that
his control loads are more manageable, making the aircraft easier
to fly.
Horn Balance
Figure 32
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1.1.11 AERODYNAMIC BALANCE INSET HINGE
This method is similar to and has the same effect as the horn
balance. Instead of having a forward projection at one or both ends
of the control surface, the hinges are set back so that the area
forward of the hinge line, which projects into the air flow when
the control surface is moved from neutral, is spread evenly along
its whole length.
Inset Hinge Balance
Figure 33
1.1.11.1 Aerodynamic Balance Balance Panels
A device fitted to a few aircraft is the aerodynamic balance
panel. Often used in the aileron system, the panel is fitted
between the leading edge of the aileron, ahead of the hinge and the
rear face of the wing. When the aileron is deflected upwards
(downwards) from neutral, the high velocity, low pressure air
passing over the lower (upper) gap decreases the air pressure under
(above) the balance panel and pulls it down (up). The force on the
balance panel is proportional to airspeed and control surface
deflection and assists the pilot in moving the controls
accordingly.
Aerodynamic Balance Panel Figure 34
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1.2 HIGH SPEED FLIGHT
Advancement in modern aircraft and engine design has produced
very large airliners capable of cruising at 87% of the speed of
sound. Typically at an altitude of 11,000 metres (approximately
36,000feet), this will amount to an airspeed of about 575 miles per
hour.
Earlier in the course the effects of subsonic air were
considered. As airspeed increases, the aerodynamic effects of
airflow passing over an aircraft, go through a series of changes,
which will now be considered.
1.2.1 SPEED OF SOUND
One of the most important measurements in high speed
aerodynamics is based on the speed of sound and so called mach
number.
Mach number is named after the Austrian physicist Ernst Mach
(1838-1916) and is the ratio of true airspeed of an aircraft to the
local speed of sound at that altitude. (This will be covered in
more detail later).
Sound waves, like those produced by a stationary object
vibrating at certain frequencies, will cause a continuous series of
pulses or pressure waves, to radiate outwards equally in all
directions from the point of origin and travel in exactly the same
manner as the ripples on a pond.
Pressure Waves Stationary Object Figure 35
The actual speed at which the waves radiate, depends on the type
and density of the material in which they are travelling. Air and
Water are both fluids but water is more dense than air, so sound
waves will travel faster (about 4 times) in water than in air.
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Additionally, in any one of the fluids, speed will vary with a
change in temperature. As temperature increases, the speed of sound
will increase and vice-versa, so that in Air on a standard day at
sea level (15oC approx), the waves will travel at 761mph (661.7
knots), whereas at 11,000 metres altitude, the speed will fall to
661mph, since the temperature has dropped to -56oC at this
altitude.
Note: At altitudes above 11,000 metres and up to about 27,000
metres, the temperature and hence the speed of sound, will remain
constant.
1.2.2 SUBSONIC FLIGHT
The propagation of the pressure waves from a stationary object
has been discussed above.
When an aircraft begins to move through the air at subsonic
speeds, (a speed less than pressure wave propagation speed) the
waves still travel forward and it is as if a message is sent ahead
of the aircraft to warn of its approach.
On receipt of this message, the air streams begin to divide to
make way for the aircraft but there is very little, if any change
in the density of the air as it flows over the aircraft. This
warning message can be detected perhaps 100metres in front of the
aircraft.
Consequently, anyone standing ahead of the aircraft, would hear
it coming and be able to detect the change in the nature of the
pressure waves as the aircraft passed by. It would be similar to
the change in the pitch of the siren of a passing emergency road
vehicle.
This is often referred to as Doppler shift or Doppler
effect.
Pressure waves Subsonic Flight Figure 36
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1.2.3 TRANSONIC FLIGHT
At subsonic speeds, the study of aerodynamics is simplified by
the fact that air passing over a wing experiences only very small
changes in pressure and density. The airflow is termed
incompressible as, when it passes through a venturi, the pressure
changes without the density changing
At higher speeds, the change in air pressure and density becomes
significant and is called the compressibility effect. When air
enters a venturi at supersonic speeds, the airflow slows down and
must compress in order to pass through its throat. Once a fluid
compresses, its pressure and density will both increase.
Subsonic Airflow Figure 37
Supersonic Airflow Figure 38
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The transonic flight range encompasses sound wave velocity and
consequently is the most difficult realm of flight since some of
the air flowing over the aircraft, particularly the wings, is
subsonic and some is supersonic. As the aircraft approaches the
speed of sound, the pressure waves ahead of it will be travelling
at the same speed as the aircraft and are therefore relatively
stationary. They accumulate to form a continuous pressure wave and
consequently will result in the removal of any advance warning of
the approach of the aircraft.
Transonic Flight Pressure Waves
Figure 39
At these speeds other pressure waves, or shock waves form
wherever the airflow reaches the speed of sound. These waves will
upset the aerodynamic balance of the wing and this phenomenon will
be covered later in the notes.
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1.2.4 SUPERSONIC FLIGHT
Once the aircraft is supersonic, all parts of it are considered
to be above the speed of sound and therefore travelling faster than
the rate of propagation of the pressure waves. An infinite number
of pressure waves are produced and form a cone, the inclination of
which will change as the aircraft speed changes.
Mach Cone Figure 40
1.2.4.1 Mach Number
As previously mentioned, Mach number is the ratio of the true
airspeed of the aircraft and the local speed of sound at that
altitude. An aircraft travelling at exactly the speed of sound is
said to be travelling at Mach 1.
It follows therefore that an aircraft travelling at twice the
speed of sound would be travelling at Mach 2 and at half the speed
of sound, Mach 0.5, etc,.
The following definitions regarding airflow and mach number
apply:
Subsonic Flow Mach Numbers below Mach 0.75
Transonic Flow Mach Numbers between Mach 0.75 and Mach 1.2
Supersonic Flow Mach Numbers between Mach 1.2 and 5.0
Hypersonic Flow Mach Numbers above Mach 5.0
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1.2.4.2 Critical Mach Number
At any constant aircraft forward speed, the speed of the airflow
will vary over the curves and cambers on the different areas of the
airframe. The behaviour of the airflow over the wing will be
particularly significant, since this is the major lift provider for
the aircraft.
As air flows over the camber on the upper surface of the wing,
its speed will increase as it flows rearwards from the leading
edge, reaching a maximum at the thickest part of the wing chord.
This means that although the aircraft itself may be travelling at
an airspeed well below Mach 1, the airflow over the thickest part
of the wing chord, may have already reached Mach 1
As will be discussed later, many unwanted effects occur when the
wing approaches and reaches Mach 1. Therefore, the designers may
either incorporate features that will lessen the unwanted effects,
or limit the aircraft to a predetermined maximum airspeed, that
will ensure the wing speed remains below Mach 1 and thus avoids the
unwanted effects altogether.
For each aircraft type therefore, a unique maximum aircraft
forward speed will be calculated, corresponding to a wing speed of
Mach 1. This aircraft speed (always be less than Mach 1) is called
the Critical Mach Number or M.crit and non-supersonic aircraft
flying in the transonic flight range, will normally be limited to a
maximum speed set below the Critical Mach number.
Critical Mach Number Figure 41
A thick wing will cause the airflow to speed up over the camber
and reach Mach 1 more quickly than a thin wing of similar chord
length. Consequently, the Critical Mach number for the thinner wing
will be a higher value than the thicker wing.
This in turn will mean that the aircraft with a thin wing, will
be able to fly faster in the transonic flight range than the one
with the thicker wing, before the unwanted effects caused by the
wing reaching Mach 1 ensue.
Conversely, less lift will be produced by a thin wing, than a
thick wing of similar chord length, but this can be overcome by the
so called Supercritical wing chord.
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In this design, the total amount of lift lost by the shallower
camber of the thin wing is restored by making the chord longer.
This is perfect for transonic cruise conditions, but at low
airspeeds, lift on a clean wing will be insufficient and so
extensive use of high lift devices (slots, slats and flaps) is
necessary
Supercritical Wing
Figure 42
1.2.4.3 Adverse Transonic Effects
Even though the onset of compressibility is gradual, it begins
to have a significant effect as the Critical Mach number is
approached. Unwanted adverse effects including, buffeting, shock
waves, increase in drag, decrease in lift and movement of the
centre of pressure occur.
If uncontrolled, these effects could result in the aircraft
becoming difficult to fly and to behave in a similar manner to a
low speed high incidence stall, even though the aircraft is at high
speed and low angle of incidence.
1.2.4.4 Compressibility Buffet
Previously discussed has been the build up of the pressure wave
in front of the aircraft as it approaches Mach 1, including the
fact that other parts of the airframe, in particular the wing, are
likely to reach Mach 1 well before the complete aircraft does.
When this occurs the smoothness of the airflow over the wing is
severely affected. This region, as well as those on the flying
control aerofoils, experience violent vibration and so-called
compressibility buffeting of the airframe. If allowed to continue,
control loss or possible structural damage can occur.
1.2.4.5 Shock Wave
Previously in the notes, the build up of pressure waves and the
change from incompressible to compressible flow as the aircraft or
an aerofoil surface approaches the speed of sound, has been
discussed. Transonic flight presents major design problems for the
aerofoil in particular, because only a portion of the airflow
passing over the wing becomes supersonic.
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When an aerofoil moves through the air at a speed below its
critical Mach number, all of the airflow is subsonic and the
pressure distribution is predictable.The first indication of a
change in the nature of the flow will be a breakaway of the airflow
from the aerofoil surface as described previously in boundary layer
control. Any turbulence resulting from the separation will cause an
increase in drag and a corresponding reduction in the amount of
lift. As speed begins to increase, the point of separation moves
forward, extending the turbulent wake.
Subsonic Flow Over all the Surface Figure 43
However, as flight speed reaches and exceeds the critical Mach
number, the airflow over the top of the wing speeds up to
supersonic velocity and a shock wave starts to form.
The First Sonic Flow is encountered
Figure 44
A Normal Shock Wave Begins to Form Figure 45
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Note: If the aerofoil is symmetrical and set at zero degrees
angle of attack, the incipient shock wave as it is called, would
form equally on the upper and lower surfaces. However, because the
wing is usually set to an angle of incidence of about 3 degrees,
even a symmetrical aerofoil section would produce the incipient
wave on the top surface first.
The wave extends outwards more or less at right angles to the
aerofoil surface and is referred to as a normal (perpendicular)
shock wave This normal shock wave forms a boundary between
supersonic and subsonic airflow.
As we have seen the high velocity airflow over the top of a wing
creates an area of low pressure. The shock wave causes it to
decelerate to subsonic speed, resulting in a rapid rise in
pressure. The separation point and turbulent wake will now start
from this point, resulting in a sudden and considerable increase in
drag (about 10 times) and therefore a large loss of lift. Severe
buffeting is likely, which could even lead to a shock stall and the
centre of pressure will be altered, affecting the pitching
moment.
This extra drag, so called Shock Drag, will be made up of two
components, namely Wave Drag, resistance caused by the wave itself
and Boundary Layer Drag, due to the increased turbulent region over
the surface of the wing. Furthermore, this shock-induced separation
is likely to reduce flying control effectiveness
The velocity of the air leaving the shock wave remains
supersonic, so both the static pressure and the density of the air
increase adding to the high drag/ low lift condition. Additionally,
some of the energy in the airstream will be dissipated in the form
of heat.
As the aircraft speed continues to increase, the wave will
extend outwards and begin to move aft towards the trailing edge of
the wing. A second wave begins to form on the lower surface, as the
airflow here also speeds up to supersonic velocity
Shock Induced Separation Occurs
Figure 46
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As the airspeed reaches the upper end of the transonic range,
both shock waves move aft, become stronger and will eventually
attach to the wing's trailing edge.
Almost all Flow is Supersonic, Some Shock Induced Separation
Figure 47
Further increases in forward speed will now result in the
characteristic normal shock wave forming ahead of the aerofoil.
This continuous wave, known as a Bow wave, will move towards and
subsequently attach itself, to the leading edge of the wing. Once
attached, all airflow over the wing will be supersonic and many of
the unwanted transonic effects are eliminated.
The Bow Wave is starting to Form
Figure 48
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As can be seen in figure 49, the transonic region has a great
affect on the lift and drag. Both values rise until Mach 0.81, when
shock induced separation drastically reduces the coefficient of
lift. As speed approaches Mach 0.99, a bow wave is forming and
airflow over the wing is slowed to subsonic speeds, resulting in an
increase in lift coefficient and a reduction of drag.
Lift / Drag Comparison at 2 Angle of Attack
Figure 49
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1.2.5 AERODYNAMIC HEATING
One of the biggest problems of sustained supersonic flight is
aerodynamic heating of the aircraft structure. An extreme example
of aerodynamic heating might be a shooting star, when its material
overheats to the point of destruction, from the heat generated by
friction heating with the earth's atmosphere.
In the commercial world, Concorde was probably the only airliner
where aerodynamic heating presents a significant problem. When the
aircraft was flown at Mach 2, the friction of the air passing
around the aircraft heats the skin considerably even at altitudes
in excess of 17,000 metres. The point of maximum heating is on the
nose where the rise in temperature could reach 1750C.
As a precaution, a probe on the nose of the aircraft monitors
the temperature during flight. When a reading of 1270C is reached,
the flight deck is directed to reduce the speed to about Mach 1.8,
to bring the temperature back within limits.
Concorde used conventional aluminium alloys in its construction.
If future aircraft were required to travel within the atmosphere at
even higher Mach numbers, other materials such as titanium alloy or
stainless steel would need to be considered.
Concord Skin Temperature
Figure 50
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1.2.6 AREA RULE
Area rule is an aerodynamic technique used in the design of
high-speed aircraft.
If drag is to be kept to a minimum at transonic speeds, aircraft
must be slim, smooth and streamlined. In general terms it means
that the wings, fuselage, empennage and other appendages have to be
considered together when working out the total streamlining. This
is necessary so that the cross-sectional area of successive slices
of the aircraft from nose to tail, conform to those of a simple
body of streamline shape.
Area rule is defined as: For the minimum drag at the
connections, (wing/fuselage), the variation of the aircrafts total
cross-sectional area along its length, should approximate that of
an ideal shape having minimum wave drag.
Without area rule, the greatest frontal cross-sectional area of
the fuselage would occur where the wings are attached to the
fuselage. Therefore, one method of achieving area rule in this
situation is to reduce the cross-sectional area of the fuselage,
thereby cancelling out the increase caused by the wings.
Alternatively, the fuselage cross-section could be increased
with the use of enlarged sections behind and in front of the wings
to eliminate sudden changes in the cross-sectional area and achieve
the same result.
Area Rule Figure 51
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1.2.7 FACTORS AFFECTING AIRFLOW IN ENGINE INTAKES OF HIGH SPEED
AIRCRAFT
Engine intakes on aircraft that operate in the subsonic flight
range only can be of almost any form.
The main criteria are that the airflow reaching the compressor
stage of the engine during cruise ideally does not exceed Mach 0.5.
This is normally achieved by the careful design of the intake
ducts.
Obviously, if the aircraft never exceeds Mach 0.5, a parallel
intake duct could be employed, but if the aircraft is to cruise at
airspeeds in excess of this, yet below Mach 1, a divergent duct
must be utilised to slow the airflow at the compressor down to Mach
0.5.
If the aircraft is designed to cruise above Mach 1, the air
entering the intakes will be supersonic and will behave in
accordance with the rules of supersonic flow. In this case a
convergent duct would be necessary to slow down the airflow to the
compressor.
However the aircraft must fly through the transonic range in
order to reach supersonic speed so both types of duct will be
necessary.
One way to overcome the problem is to have moveable doors that
change the intake duct shape from divergent to convergent
cross-section as the aircraft passes through Mach 1. See figure 52.
This technique can be found on the intakes of Concorde.
Other methods to control airflow reaching the compressor is to
make use of the fact that air passing through a shock wave slows
down to a lower speed. This type of intake design is usually
characterised by the bullet fairing, which on some aircraft can
translate in and out of the intake to reposition the shock wave
during low or high supersonic flight speeds. See Figure 53
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Intake Moveable doors Figure 52
Bullet Fairing Intake Figure 53
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1.2.8 EFFECTS OF SWEEPBACK ON CRITICAL MACH NUMBER
In order to fly at high speed in the transonic range without
encountering the problems caused by the production of shock waves,
the Critical Mach number needs to be as high as possible. As has
already been shown, one way is to have as thin a wing as possible.
This of course is an acceptable solution in theory, but in practice
there will be structural integrity problems, such as wing loading,
strength and flexibility.
Another way of raising the Critical Mach number without the
structural limitations is by the use of swept wings. Sweepback not
only delays the production of the shock wave, but reduces the
severity of the shock stall should it occur. The theory behind this
is that it is only the component of velocity over the wing chord
that is responsible for the pressure distribution and so for
causing the shock wave to develop. The other velocity component
that travels spanwise causes only frictional drag and has no effect
on shock wave production.
This theory is borne out by the fact that when it does appear,
the shock wave lies parallel to the span of the wing. Therefore
only that part of the velocity perpendicular to the shock wave,
i.e. across the chord, is reduced by the shock wave to subsonic
speeds.
The greater the sweepback, the smaller will be the component of
velocity affected, resulting in a higher Critical Mach number and a
reduction in drag at all transonic speeds. Additionally sweepback
results in a thinner mean aerodynamic chord, which raises the
Critical Mach number even more.
Effects of Sweepback Figure 54
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CONTENTS
2 AIRFRAME STRUCTURES GENERAL CONCEPTS ................ 2-1
2.1 AIRWORTHINESS REQUIREMENTS FOR STRUCTURAL STRENGTH .....
2-1
2.1.1 STRUCTURAL CLASSIFICATION
...................................................... 2-1 2.1.2
Primary structure
........................................................... 2-2
2.1.3 Secondary Structure
..................................................... 2-4 2.1.4
Tertiary Structure
.......................................................... 2-4
2.2 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS
............ 2-4 2.2.1 Fail Safe
........................................................................
2-4 2.2.2 Safe Life
........................................................................
2-4 2.2.3 Damage Tolerance
........................................................ 2-5
2.3 ZONAL AND STATION IDENTIFICATION SYSTEM
................................ 2-7 2.3.1 Zonal System
................................................................
2-7 2.3.2 Station Identification System
......................................... 2-8
2.4 LOADS FOUND WITHIN THE STRUCTURE STRESS AND STRAIN ......
2-9 2.4.1 Compression
.................................................................
2-10 2.4.2 Tension
.........................................................................
2-10 2.4.3 Bending
.........................................................................
2-11 2.4.4 Torsion
..........................................................................
2-12 2.4.5 Shear
............................................................................
2-12 2.4.6 Hoop Stress
..................................................................
2-13 2.4.7 Metal Fatigue
................................................................
2-13
2.5 DRAINAGE AND VENTILATION PROVISIONS
..................................... 2-16 2.5.1 External Drains
............................................................. 2-16
2.5.2 Internal Drains
...............................................................
2-18 2.5.3 Ventilation
.....................................................................
2-18
2.6 LIGHTNING STRIKE PROVISION
...................................................... 2-19
2.7 CONSTRUCTION METHODS
............................................................ 2-20
2.7.1 Stressed Skin Fuselage
................................................ 2-20 2.6.1 Frames
and Formers .....................................................
2-21 2.6.2 Bulkheads
.....................................................................
2-21 2.6.3 Longerons and Stringers
............................................... 2-22 2.6.4 Doublers
and Reinforcement ......................................... 2-23
2.6.5 Struts and Ties
.............................................................. 2-23
2.6.6 Beams and Floor Structures
.......................................... 2-24 2.6.7 Methods of
Skinning ......................................................
2-24 2.6.8 Anti-Corrosive Protection
.............................................. 2-26 2.6.9
Construction Methods Wing .......................................
2-27 2.6.10 Construction Methods Empennage
............................ 2-28 2.6.11 Construction Methods
Engine Attachments ................ 2-29 2.6.12 Structural Assembly
Techniques ................................... 2-31 2.6.13 Solid
Shank Rivets
........................................................ 2-31
2.6.14 Special and Blind Fasteners.
......................................... 2-33 2.6.15 Bolts and
Nuts
...............................................................
2-38 2.6.16 Adhesive Bonded Structures
......................................... 2-43 2.6.17 Methods of
Surface Protection ...................................... 2-45
2.6.18 Exterior Finish Maintenance
.......................................... 2-47
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2 AIRFRAME STRUCTURES GENERAL CONCEPTS
2.1 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
(EASA-25) and for aircraft below 5700kg MTWA, in EASA-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.
2.1.1 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.
2.1.2 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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2.1.3 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
2.1.4 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.
2.2 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS
2.2.1 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.
2.2.2 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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2.2.3 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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2.3 ZONAL AND STATION IDENTIFICATION SYSTEM
2.3.1 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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2.3.2 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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2.4 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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2.4.1 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
2.4.2 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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2.4.3 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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2.4.4 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
2.4.5 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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2.4.6 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
2.4.7 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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2.5 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
2.5.1 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 aerofoil