JAR 66 CATEGORY B1
uk
engineering
MODULE 11.01
Theory of Flight
Contents
1 MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURES AND
SYSTEMS) .......................................................................................... 1-2
1.1
1.2
2
AEROPLANE AERODYNAMICS AND FLIGHT CONTROLS ..................... 1-2
1.1.1
Fixed Aerofoils .............................................................. 1-2
1.1.2
Moveable Control Surfaces ........................................... 1-7
1.1.3
High Lift Devices ........................................................... 1-14
1.1.4
Drag Inducing Devices .................................................. 1-15
1.1.5
Airflow Control Devices – Wing Fences......................... 1-18
1.1.6
Boundary Layer Control ................................................ 1-19
1.1.7
Trim Tabs ...................................................................... 1-22
1.1.8
Mass Balance ............................................................... 1-25
HIGH SPEED FLIGHT ..................................................................... 1-2
1.2.1
Speed of Sound ............................................................ 1-2
1.2.2
Subsonic Flight ............................................................. 1-3
1.2.3
Transonic Flight ............................................................ 1-4
1.2.4
Supersonic Flight .......................................................... 1-6
1.2.5
Aerodynamic Heating .................................................... 1-13
1.2.6
Area Rule ...................................................................... 1-14
1.2.7
Factors Affecting Airflow in Engine Intakes of High Speed Aircraft
1-15
1.2.8
Effects of Sweepback on Critical Mach Number ............ 1-17
SAFETY PRECAUTIONS – AIRCRAFT & WORKSHOP ............. 2-3
2.1
ACCIDENTS .................................................................................. 2-3
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MODULE 11.01
Theory of Flight
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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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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T-Tail Arrangement
The T-Tail arrangement places the complete stabiliser/tailplane and elevator
assembly at the top of the vertical stabiliser. The use of this system not only
makes the fin and rudder more effective by the so-called ‘end-plate effect’ but
also ensures pitch control is not affected by wing turbulence, (except during an
unwanted deep-stall condition).
However, this configuration has the disadvantage that the whole empennage
structure will be heavier than normal, due to the strengthening required to combat
the greater bending loads produced by this layout.
On some aircraft, the complete tailplane can be moved through several degrees
angle of attack to provide a trimming facility as an alternative to trim tabs.(later).
T–Tail Arrangement
Figure 3
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Theory of Flight
Unconventional Aircraft Design – Flying Wing
Minus the fuselage, which on a conventional aircraft generates minimum lift, but
produces a large amount of drag and without the vertical and horizontal
stabilising surfaces which also generate drag, the flying wing appears to be the
most perfect low-drag design.
However, without conventional wings, fuselage or empennage, the stability and
controllability of this design has until recently, been extremely poor.
Only with the advent of high computing power generating artificial stability and
actively operating unorthodox flying controls, has the design become a practically
feasibility.
On the example shown below, about 100 computers are used to control and
monitor every aspect of flight.
B2 Flying Wing
Figure 4
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Vertical Stabiliser
The vertical stabiliser for an aircraft is the aerofoil 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 5
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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 6
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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 sidestick.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 7
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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.
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 8
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 downgoing 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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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.
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 allflying 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 9
Stabilator Controls
Figure 10
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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 11
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 – Beach Starcraft
Figure 12
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).
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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.
Rudder controls
Figure 13
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 14
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 15
Additionally, on a few aircraft, ailerons designed to 'droop' when the trailing edge
flaps are lowered to certain positions, act as additional plain flaps. Roll control is
retained, but extra lift (and drag) is generated during landing.
These surfaces are usually referred to as Flaperons or sometimes called droop
ailerons.
Droop Aileron
Figure 16
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.
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 17
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.
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 18
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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 19
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 20
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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 21
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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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 22
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 23
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 24
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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.
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 25
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 26
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).
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 27
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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 28
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 29
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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 nonexistent 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 30
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 31
Integral Mass Weights
Figure 32
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Control Surface Bias
True control surface bias is achieved in manually operated controls by the use of
fixed or adjustable trim tabs, as previously discussed. In power operated controls
the input signal to the hydraulic servo valve is adjusted to offset the primary
control surface. (This will be covered later).
However 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.
Aerodynamic Balance – Horn Balance
In 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 33
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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 34
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 35
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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 36
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 37
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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 38
Supersonic Airflow
Figure 39
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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 40
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 41
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
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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 nonsupersonic 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 42
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 43
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.
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.
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 44
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 45
A Normal Shock Wave Begins to Form
Figure 46
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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 47
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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 48
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 49
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As can be seen in figure 50, 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 2º Angle of Attack
Figure 50
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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 fact, if
it were not for the special ceramic tile ‘heat-sink’ insulation on the structure of the
Space Shuttle, a similar fate might occur to it on re-entry.
In the commercial world, Concorde is probably the only airliner where
aerodynamic heating presents a significant problem. When the aircraft is 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 uses conventional aluminium alloys in its construction. If future aircraft
are 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 51
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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 aircraft’s 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 52
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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 is 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 which change the
intake duct shape from divergent to convergent cross-section as the aircraft
passes through Mach 1. See figure 53. 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 54
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Intake Moveable doors
Figure 53
‘Bullet Fairing’ Intake
Figure 54
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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
which is responsible for the pressure distribution and so for causing the shock
wave to develop. The other velocity component which 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 55
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SAFETY PRECAUTIONS – AIRCRAFT & WORKSHOP
2.1 ACCIDENTS
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